Showing papers on "Freestream published in 2010"

Simulation of Turbulent Mixing Behind a Strut Injector in Supersonic Flow

Abstract: The flowfield downstream of a strut-based injection system in a supersonic combustion ramjet is investigated using large-eddy simulation with a new localized dynamic subgrid closure for compressible turbulent mixing. Recirculations are formed at the base of the strut in the nonreacting flow and trap some of the injected fluid. The high levels of turbulence along the underexpanded hydrogen jets and in the shear layer lead to a high level of mixing of fuel and freestream fluids. Furthermore, the shear layer unsteadiness permits efficient large-scale mixing of freestream and injected fluids. In the reacting flowfield, the flame anchoring mechanism is, however, found to depend more on a recirculation region located downstream of the injectors than on their sides. A region of reverse flow is formed that traps hot products and radicals. Intermittent convection of hot fluid toward the injector occurs and preheats the reactants.

Instability and transition on the HIFiRE-5 in a Mach-6 Quiet Tunnel

Abstract: : A 2:1 aspect-ratio elliptic cone with a blunt nosetip was tested in the Boeing/ AFOSR Mach-6 Quiet Tunnel to investigate the effects of freestream noise level, surface roughness, angle of attack, and freestream Reynolds number on laminar-to turbulent boundary layer transition on the windward surface The cone had a minor axis half-angle of 7-deg and a nose radius of 095 mm Temperature-sensitive paint enabled a global measurement of the temperature distribution and detection of the transition front Transition apparently arising from two mechanisms was observed: transition along the centerline suspected to arise from the amplification of second mode waves in the inflected boundary layer, and transition roughly halfway between the centerline and leading edges probably due to the breakdown of cross flow vortices Reducing noise level from conventional (root-mean-square pressure 3% of the mean) to quiet (root-mean-square pressure less than 01% of the mean) substantially delayed transition due to both mechanisms Increasing the angle of attack from 0-deg to 4-deg delayed the cross flow transition mode on the windward side Transition moved forward as free stream unit Reynolds number increased from 26-deg 106 /m to 119-deg 106 /m PCB fast-response pressure transducers installed along the model centerline detected apparent instabilities at frequencies from 50 to 150 kHz prior to transition under noisy flow; quiet-flow results are less clear Mass flux profiles along the centerline were measured with a calibrated hot wire

Effect of Density Ratio on Flat Plate Film Cooling With Shaped Holes Using PSP

Abstract: Detailed film cooling effectiveness distributions are obtained on a flat plate using the pressure sensitive paint (PSP) technique. The applicability of the PSP technique is expanded to include a coolant-to-mainstream density ratio of 1.4. The effect of density ratio on the film cooling effectiveness is coupled with varying blowing ratio (M = 0.25–2.0), freestream turbulence intensity (Tu = 1%–12.5%), and film hole geometry. The effectiveness distributions are obtained on three separate flat plates containing either simple angle, cylindrical holes, simple angle, fanshaped holes (α = 10°), or simple angle, laidback, fanshaped holes (α = 10°, γ = 10°). In all three cases, the film cooling holes are angled at θ = 35° from the mainstream flow. Using the PSP technique, the combined effects of blowing ratio, turbulence intensity, and density ratio are captured for each film cooling geometry. The detailed film cooling effectiveness distributions, for cylindrical holes, clearly show the effectiveness at the lowest blowing ratio is enhanced at the lower density ratio (DR = 1). However, as the blowing ratio increases, a transition occurs, leading to increased effectiveness with the elevated density ratio (DR = 1.4). In addition, the PSP technique captures an upstream shift of the coolant jet reattachment point as the density ratio increases or the turbulence intensity increases (at moderate blowing ratios for cylindrical holes). With the decreased momentum of the shaped film cooling holes, the greatest film cooling effectiveness is obtained at the lower density ratio (DR = 1.0) over the entire range of blowing ratios considered. In all cases, as the freestream turbulence intensity increases, the film effectiveness decreases; this effect is reduced as the blowing ratio increases for all three film hole configurations.Copyright © 2010 by ASME

Navier-Stokes Simulations of Hypersonic Flows with Coupled Graphite Ablation

Abstract: equilibriumablationwithsurfacemassandenergybalancesfullycoupledwiththenumericalsolverandcanaccount for both surface oxidation and sublimation. The surface temperature is obtained from the steady-state ablation approximation. This numerical procedure can predict aerothermal heating, chemical species concentrations, and carbon material ablation rate over the heat-shield surface of reentry vehicles. Two-dimensional axisymmetric simulations have been performed to numerically reproduce the ablation of a graphite sphere cone that has been tested in the Interaction Heating Facility at the NASA Ames Research Center. The freestream conditions of the selected test case are typical for Earth reentry from a planetary mission. The predicted ablation rate and surface temperatureassumingfrozenchemistryinthe flowshowagoodagreementwiththeavailableexperimentaldata.The agreementisfurtherimprovedfreezingthenitrogenrecombinationreactionatthesurfacetobemoreconsistentwith experimental observation, which has shown nitrogen atom recombination not to occur at the graphite surface.

Dielectric barrier discharge control of a turbulent boundary layer in a supersonic flow

Abstract: We demonstrate effective manipulation of a turbulent boundary layer at Mach 4.7 conditions using a surface dielectric barrier discharge (DBD) actuator. The freestream conditions of low static pressure (1 kPa) and temperature (60 K) are conducive to the visualization of flow features using Rayleigh scattering from condensed CO2 particles. The boundary layer thinning is observed when spanwise momentum is induced by the low power (6.8 W), low frequency (28 kHz) single actuator pair oriented parallel to the freestream flow.

Influence of Coolant Density on Turbine Blade Film-Cooling Using Pressure Sensitive Paint Technique

Abstract: Adiabatic film-cooling effectiveness is examined on a high pressure turbine blade by varying three critical engine parameters, viz., coolant blowing ratio, coolant-to-mainstream density ratio and freestream turbulence intensity. Three average coolant blowing ratios (BR = 1.2, 1.7, and 2.2 on the pressure side and BR = 1.1, 1.4, and 1.8 on the suction side), three average coolant density ratios (DR = 1.0, 1.5, and 2.5), and two average freestream turbulence intensities (Tu = 4.2% and 10.5%) are considered. Conduction-free Pressure Sensitive Paint (PSP) technique is adopted to measure film-cooling effectiveness. Three foreign gases— N2 for low density, CO2 for medium density, and a mixture of SF6 and Argon for high density are selected to study the effect of coolant density. The test blade features 2 rows of cylindrical film-cooling holes on the suction side (45° compound), 4 rows on the pressure side (45° compound) and 3 around the leading edge (30° radial). The inlet and the exit Mach numbers are 0.24 and 0.44, respectively. Reynolds number of the mainstream flow is 7.5E105 based on the exit velocity and blade chord length. Results suggest that the PSP is a powerful technique capable of producing clear and detailed film effectiveness contours with diverse foreign gases. Large improvement on the pressure side and moderate improvement on the suction side effectiveness is witnessed when blowing ratio is raised from 1.2 to 1.7 and 1.1 to 1.4, respectively. No major improvement is seen thereafter with the downstream half of the suction side showing drop in effectiveness. The effect of increasing coolant density is to increase effectiveness everywhere on the pressure surface and suction surface except for the small region on the suction side, xss /Cx <0.2. Higher freestream turbulence causes effectiveness to drop everywhere except in the region downstream of the suction side where significant improvement in effectiveness is seen.Copyright © 2010 by ASME

Effect of Freestream Turbulence Intensity on Film Cooling Jet Structure and Surface Effectiveness Using PIV and PSP

Abstract: An experimental investigation of film cooling jet structure using two-dimensional, particle image velocimetry (PIV) has been completed for cylindrical, simple angle (θ = 35°) film cooling holes. The PIV measurements are coupled with detailed film cooling effectiveness distributions on the flat plate obtained using a steady state, pressure sensitive paint (PSP) technique. Both the flow and surface measurements were performed in a low speed wind tunnel where the freestream turbulence intensity was varied from 1.2% to 12.5%. With this traditional film cooling configuration, the blowing ratio was varied from 0.5–1.5 to compare the jet structure of relatively low and high momentum cooling flows. Velocity maps of the coolant flow (in the streamwise direction) are obtained on three planes spanning a single hole: centerline, 0.25D, and 0.5D (outer edge of the film cooling hole). From the seeded jets, time averaged, mean velocity distributions of the film cooling jets are obtained near the cooled surface. In addition, turbulent fluctuations are obtained for each flow condition. Combining the detailed flow field measurements obtained using PIV (both instantaneous and time averaged) with detailed film cooling effectiveness distributions on the surface (PSP), provides a more complete view of the coolant jet–mainstream flow interaction. Near the edge of the film cooling holes, the turbulent mixing increases, and as a result the film cooling effectiveness decreases. Furthermore, the PIV measurements show the increased mixing of the coolant jet with the mainstream at the elevated freestream turbulence level resulting in a reduction of the jet to effectively protect the film cooled surface.Copyright © 2010 by ASME

A coaxial thermocouple for shock tunnel applications

Abstract: A chromel-constantan coaxial surface junction thermocouple has been designed, fabricated, calibrated, and tested to measure the temperature-time history on the surface of a body in a hypersonic freestream of Mach 8 in a shock tunnel. The coaxial thermocouple with a diameter of 3.25 mm was flush mounted in the surface of a hemisphere of 25 mm diameter. The hypersonic freestream was of a very low temperature and density, and had a flow time of about a millisecond. Preliminary test results indicate that the thermocouple is quite sensitive to low temperature-rarefied freestreams, and also has a response time of a few microseconds (≈5 μs) to meet the requirements of short duration transient measurements. The sensor developed is accurate, robust, reproducible, and is highly inexpensive.

Experimental Study of Near Wake of Micro Vortex Generators in Supersonic Flow

Abstract: Detailed schlieren and laser lightsheet visualizations of the near wake of micro vortex generator (MVG) revealed large structures that were different from those of the undisturbed turbulent boundary layer. These structures were attributed to the rapid breakdown of the primary trailing vortex pair. The breakdown was thought to arise from a cylindrical Kelvin–Helmholtz-like instability surface. The structures appear to be hairpin or ring-like in nature that showed eruptions into the freestream flow, entraining it.

High-Speed Boundary-Layer Transition Induced by an Isolated Roughness Element

Abstract: Progress on an experimental effort to quantify the instability mechanisms associated with roughness-induced transition in a high-speed boundary layer is reported in this paper. To simulate the low-disturbance environment encountered during high-altitude flight, the experimental study was performed in the NASA-Langley Mach 3.5 Supersonic Low-Disturbance Tunnel. A flat plate trip sizing study was performed first to identify the roughness height required to force transition. That study, which included transition onset measurements under both quiet and noisy freestream conditions, confirmed the sensitivity of roughness-induced transition to freestream disturbance levels. Surveys of the laminar boundary layer on a 7deg half-angle sharp-tipped cone were performed via hot-wire anemometry and pitot-pressure measurements. The measured mean mass-flux and Mach-number profiles agreed very well with computed mean-flow profiles. Finally, surveys of the boundary layer developing downstream of an isolated roughness element on the cone were performed. The measurements revealed an instability in the far wake of the roughness element that grows exponentially and has peak frequencies in the 150 to 250 kHz range.

Aero-Optical Measurements in a Heated, Subsonic, Turbulent Boundary Layer

Abstract: A scaling relationship for OPD rms that accounts for the fluctuating total temperature profile within a turbulent boundary is derived from the modified Crocco relation. Experimental data from heated, compressible boundary layers at six subsonic Mach numbers in two wind tunnel facilities are shown to be consistent with the theory. The results show that a temperature mismatch between the freestream and underlying wall has a significant impact on the overall optical aberration.

Description and Flow Assessment of the Delft Hypersonic Ludwieg Tube

Abstract: An assessment of the hypersonic Ludwieg tube of Delft University of Technology is given. The facility is described theoretically, and an experimental evaluation is performed to infer the facility performance. Experiments are performed using conventional techniques such as static and total head pressuremeasurements andFay–Riddell heat flux evaluations by means of infrared thermography. Furthermore, particle image velocimetry is used to deduce nozzle boundary-layer parameters aswell as the freestreamvelocityfield and the static and total temperatures for the Mach 7 nozzle. For the Mach 9 nozzle, stagnation heat flux measurements are performed to obtain the total temperature of the flow. The freestream values were determined experimentally in two different ways and the results showed good agreement. The application of particle image velocimetry allows the freestream flowfield to be directly obtained and gives a directmeasure for the flowfield uniformity (0.2%) and repeatability (0.4%). The static and total temperatures calculated from the particle image velocimetry results showed that there is a large mismatch between the theoretical total temperature and measured total temperature, which is attributed to heat losses present in the throat and nozzle.

Large-eddy simulations of a turbulent Coanda jet on a circulation control airfoil

Abstract: Large-eddy simulations are performed of a turbulent Coanda jet separating from a rounded trailing edge of a simplified circulation control airfoil model. The freestream Reynolds number based on the airfoil chord is 0.49×106, the jet Reynolds number based on the jet slot height is 4470, and the ratio of the peak jet velocity to the freestream velocity is 3.96. Three different grid resolutions are used to show that their effect is very small on the mean surface pressure distribution, which agrees very well with experiments, as well as on the mean velocity profiles over the Coanda surface. It is observed that the Coanda jet becomes fully turbulent just downstream of the jet exit, accompanied by asymmetric alternating vortex shedding behind a thin (but blunt) jet blade splitting the jet and the external flow. A number of “backward-tilted” hairpin vortices (i.e., the head of each hairpin being located upstream of the legs) are observed around the outer edge of the jet over the Coanda surface. These hairpins cr...

Compressibility effects on the vortical flow over a 65° sweep delta wing

Abstract: The aim of this study is to investigate numerically the compressibility effects on the vortical flow developing over the VFE-2 delta wing. This wing is equipped with a sharp leading-edge and a sweep angle equal to 65°. The angle of attack is set equal to 25.5°, and two different freestream Mach numbers are considered: M∞=0.4 and M∞=0.8. The simulations are based on a turbulent modeling coupling Delayed Detached Eddy Simulation and Zonal Detached Eddy Simulation approaches, allowing a faster decay of the eddy-viscosity in Large Eddy Simulation regions. Such a method allows a good agreement with available experimental data. These computations highlight a modification of the flow behavior with an increase in the freestream Mach number. Indeed, the leading-edge vortex moves toward the wing and its core is dilated. Moreover, it is demonstrated in this study that the leading-edge vortex interacts with shock wave at midchord. This interaction induces a decay of the Rossby number in the leading-edge region and th...

Numerical modeling of detonation combustion of hydrogen-air mixtures in a convergent-divergent nozzle

Abstract: The flow of igniting hydrogen-air mixtures entering an axisymmetric convergent-divergent nozzle at a supersonic velocity is considered. A possibility of stabilizing detonation combustion is numerically investigated at different freestream Mach numbers with account for nonuniform distribution of hydrogen concentration at the nozzle entry. The investigation is performed on the basis of the two-dimensional gasdynamic Euler equations for a multicomponent reacting gas. A detailed model of chemical reactions is used. The calculated thrust is compared with the drag of a conical housing containing the supersonic nozzle considered.

Hypersonic interference heating in the vicinity of surface protuberances

Abstract: The understanding of the behaviour of the flow around surface protuberances in hypersonic vehicles is developed and an engineering approach to predict the location and magnitude of the highest heat transfer rates in their vicinity is presented To this end, an experimental investigation was performed in a hypersonic facility at freestream Mach numbers of 82 and 123 and Reynolds numbers ranging from Re ∞/m = 335 × 106 to Re ∞/m = 935 × 106 The effects of protuberance geometry, boundary layer state, freestream Reynolds number and freestream Mach numbers were assessed based on thin-film heat transfer measurements Further understanding of the flowfield was obtained through oil-dot visualizations and high-speed schlieren videos The local interference interaction was shown to be strongly 3-D and to be dominated by the incipient separation angle induced by the protuberance In interactions in which the incoming boundary layer remains unseparated upstream of the protuberance, the highest heating occurs adjacent to the device In interactions in which the incoming boundary layer is fully separated ahead of the protuberance, the highest heating generally occurs on the surface just upstream of it except for low-deflection protuberances under low Reynolds freestream flow conditions in which case the heat flux to the side is greater

Experimental and Numerical Studies of Diamond-Shaped Injector in a Supersonic Flow

Abstract: IMPROVING fuel–air mixing and flame holding are current research areas for combined-cycle engines for hypersonic propulsion [1–6]. The scramjet mode is particularly difficult because fuel and air residence times within the engine are of the order of milliseconds; during this time, the fuel must penetrate into and mix with the freestream air and then substantially react to completion. Ideally, penetration and then mixing of the fuel should be done with minimal pressure loss, and thus there is an incentive, especially for small-scale engines, to accomplish fuel injection through so-called nonintrusive injection ports, the most basic of which is the circular, flush-wall, normal injector. Furthermore, while penetration and mixing with the crossflow are typically of interest, other considerations may be important too, such as entrainment into a flameholding device. In the present study, the focus is on flush-wall injection through a diamond-shaped orifice [7,8]. Recently, Srinivasan and Bowersox [9,10] investigated with computational fluid dynamics (CFD) the possibility of tailoring the flow structure to enhance mixing and produce a stable vortex for gas-dynamically induced flame holding. The goal was to control the shape of the interaction barrel shock to produce a flow structure that resembled a blunt bodywith a truncated transverse plane at the trailing edge. A diamond-shaped port was found to produce the desired flow structure. Boundary-layer and injector fluid would then be entrained into this recirculation zone, termed the lateral counter-rotating vortex pair (LCVP), because the structure contains a vortex pair spanning the width of the barrel shock. To determine the robustness of the LCVP, parametric studies were performed for freestream Mach numbers from 2 to 5; it was found that by controlling the injector pressure, the LCVP could be created. The residence time within the LCVP was estimated to be an order of magnitude longer than the flow time through the solution domain, indicating that LCVP may be sufficient for flame holding under some circumstances. The objective for the present studywas to visualize the structure of the LCVP and assess reactivity within it. As with the previous study [11], the tools employed include planar laser-induced fluorescence (PLIF) of the hydroxyl (OH) and nitric oxide (NO) molecules and complementary 3-D nonreacting CFD simulations.

Experimental Study of Graphite Ablation in Nitrogen Flow, Part II: Further Numerical Analysis

Abstract: DOI: 10.2514/1.43264 The ablative behavior of graphite heated in an inductively coupled plasma wind tunnel is analyzed using an integrated computational method to examine the probability value of nitridation reaction occurring at graphite surface. In this method, the plasma torch freestream condition at the entrance of the test chamber is evaluated by calculating the flows in the plasma torch. The thermal response of the graphite test piece is calculated by loosely coupling the thermochemical nonequilibrium fluid dynamics code and two-dimensional heat conduction equation solver using the freestream condition so valuated. Using a computational tool, surface contours of the graphite test piece at several time steps are calculated by varying the probability value of nitridation reaction and are compared with those measured in the heating test. The effect of impurities remaining in the test chamber is also taken into account.Asaresultofthestudy,itissuggestedthataprobabilityvalueof0.003fornitridation reactionisreasonable to explain the amount of mass loss of the graphite test piece in the heating test.

Dielectric- Barrier Discharge Plasmas for Flow Control at Higher Mach Numbers

Abstract: The main aim of this study is to investigate the control authority of Dielectric-Barrier Discharge (DBD) plasma actuators at higher Mach numbers and higher Reynolds numbers. Different strategies are pursued to influence the flow, all attempting to create coherent structures to transport momentum from the freestream into the near-wall region of the boundary-layer. To achieve this various actuator geometries and configurations are used and tested on a NACA0012 airfoil. To obtain a better insight into the effects limiting the actuator performance under high-speed flow conditions, separate investigations on a flat plate are presented.

A rotating elliptic airfoil in fluid at rest and in a parallel freestream

Abstract: This paper reports results of DPIV measurements on a two-dimensional elliptic airfoil rotating about its own axis of symmetry in a fluid at rest and in a parallel freestream. In the former case, we examined three rotating speeds (Re c,Ω = 400, 1,000 and 2,000), and in the later case, four rotating speeds (Ro c,Ω = 2.4, 1.2, 0.6 and 0.4), together with two freestream velocities (Re c,u = 200 and 1,000) and two starting configurations of the airfoil (i.e., chord parallel to (α 0 = 0°) or normal (α 0 = 90°) to the freestream). Results show that a rotating airfoil in a stationary fluid produces two distinct types of vortex structures depending on the Reynolds number. The first type occurs at the lowest Reynolds number (Re c,Ω = 400), where vortices shed from the two edges or tips of the airfoil dissipated quickly, resulting in the airfoil rotating in a layer of diffused vorticity. The second type occurs at higher Reynolds numbers (i.e., Re c,Ω = 1,000 and 2,000), where the corresponding vortices rotated together with the airfoil. Due to the vortex suction effect, the torque characteristics are likely to be heavily damped for the first type because of the rapidly subsiding vortex shedding, and more oscillatory for the second type due to persistent presence of tip vortices. In a parallel freestream, increasing the tip-speed ratio (V/U) of the airfoil (i.e., decreasing the Rossby number, Ro c,Ω) transformed the flow topology from periodic vortex shedding at Ro c,Ω = 2.4 to the generation of a “hovering vortex” at Ro c,Ω = 0.6 and 0.4. The presence of the hovering vortex, which has not been reported in literature before, is likely to enhance the lift characteristics of the airfoil. Freestream Reynolds number is found to have minimal effect on the vortex formation and shedding process, although it enhances shear layer instability and produces more small-scale flow structures that affect the dynamics of the hovering vortex. Likewise, initial starting configuration of the airfoil, while affecting the flow transient during the initial phase of rotation, has insignificant effect on the overall flow topology. Unfortunately, technical constraint of our apparatus prevented us from carrying out complimentary force measurements; nevertheless, the results presented herein, which are more extensive than those computed by Lugt and Ohring (1977), will provide useful benchmark data, from which more advanced numerical calculations can be carried out to ascertain the corresponding force characteristics, particularly for those conditions with the presence of hovering vortex.

Numerical investigation of boundary-layer transition initiated by a wave packet for a cone at Mach 6

Abstract: Direct Numerical Simulations are performed to investigate transition initiated by a wave packet in a sharp cone boundary layer at Mach 6. In order to understand the natural transition process in hypersonic cone boundary layers, the ∞ow was pulsed through a hole on the cone surface to generate a wave packet which consisted of a wide range of disturbance frequencies and wave numbers. The ∞ow parameters for the simulations are based on the experimental conditions of the Boeing/AFOSR Mach 6 quiet{∞ow Ludwieg Tube at Purdue University. 1 First, the linear development of the wave packet was studied by forcing the ∞ow with a low amplitude pulse (0:001% of the freestream velocity). The dominant waves within the resulting wave packet were identifled as the second mode two{dimensional disturbance waves. In addition, weaker flrst mode oblique waves were also observed on the lateral sides of the wave packet. In order to investigate the weakly nonlinear transition regime, medium amplitude pulse disturbances (0:5% of the freestream velocity) were introduced. The response of the ∞ow to the medium amplitude pulse disturbances indicated the presence of a fundamental resonance mechanism. Lower secondary peaks in the disturbance wave spectrum were identifled at approximately half the frequency of the high amplitude frequency band for azimuthal mode numbers kc§55, which would be an indication of a subharmonic resonance mechanism. Finally, in order to identify more clearly which of these mechanisms ultimately leads to turbulent breakdown, a simulation with a higher forcing amplitude (5% of the freestream velocity) was performed. The developing strongly nonlinear wave packet eventually leads to localized patches of turbulent ∞ow (turbulent spots). In these nascent turbulent spots various known properties of mature turbulent spots could be identifled.

Experimental Study of Test Medium Vitiation Effects on Dual-Mode Scramjet Mode Transition

Abstract: An experimental study was performed to characterize the effect of vitiates due to combustion air preheating on dual-mode scramjet combustion. Major vitiate species (H2O and CO2) were added to the freestream of an electrically heated, direct connect facility simulating Mach 5 flight enthalpy. With dry air, the combustor operated in the supersonic mode at fuel equivalence ratios below 0.22 and in the subsonic mode for equivalence ratios above 0.26. Hysteresis was observed in the transitional region between 0.22 and 0.26, as the mode of combustion was dependent on whether the fuel flow rate was increasing or decreasing. Adding increasing amounts of water vapor and carbon dioxide to the freestream decreased combustor pressures by 10% to 30% for the same fuel equivalence ratio. Vitiates also caused the transition between supersonic to subsonic combustion to occur at a higher fuel equivalence ratio than with clean air. These results indicate the importance of accounting for vitiates when extrapolating from ground testing to flight, particularly in the transition region between subsonic and supersonic combustion regimes.

Computational and experimental study of intake ground vortices

Abstract: The ground vortices generated by an intake under both headwind and crosswind configurations have been investigated using computational and experimental approaches. The flow field of a scale-model intake was experimentally studied using stereoscopic particle image velocimetry to measure the ground vortex in conjunction with induct total pressure measurements for the internal flow. The computational predictions were performed using an unsteady Reynolds averaged Navier-Stokes approach. The experimental results show that under crosswind conditions a single ground vortex forms which becomes stronger as the crossflow velocity is increased. Under headwind conditions the measured ground vortex strength initially increases with freestream velocity before it reaches a local maximum and then reduces thereafter. The computations also exhibit the same characteristics and show good agreement with the measurements for some configurations. Based on the predictions, the complex flow field topology is investigated and a detailed flow model of the vortex flow field under crosswind conditions is proposed.

Suppression of Fluctuating Surface Pressures in a Supersonic Cavity Flow

Abstract: An experimental investigation of an open cavity with a length to depth ratio of six was conducted with and without passive suppression using a cylinder placed near the leading edge of the cavity. Five rod diameters were immersed at various gap heights in the approaching supersonic boundary layer where the freestream Mach number was 1.4. The study included surface pressure measurements and flowfield visualization using both Schlieren imaging and particle image velocimetry. The results indicate a rod sized at least 40% of the boundary layer height and placed with it’s top near the top of the boundary layer exhibited the most effective suppression configuration. In controlled cases the shear layer is spread and slightly lofted above the trailing edge of the cavity when an appropriately sized rod was placed at the cavity leading edge. Evidence of the disruption of the feedback mechanism well known to cause resonant tones were highlighted with correlation and coherence plots and further substantiated with particle image velocimetry and Schlieren images. The rod was able to reduce overall pressure fluctuations measured on the aftwall by nearly 35% and the peak tone reductions by 65%.

Numerical Study of Laser-Spot Effects on Boundary-Layer Receptivity for Blunt Compression-Cones in Mach-6 Freestream

Abstract: This paper presents the direct numerical simulation (DNS) study of the boundary layer receptivity for blunt compression cones in Mach-6 flow with freestream laser-spot (hotspot) perturbation. The flow conditions are the same as the Boeing/AFOSR Mach-6 Quiet tunnel (BAM6QT) in Purdue University. Compression-cone geometry is expected to cause laminar/turbulence transition in shorter stream-wise distance than straight-wedged cone geometry due to adverse pressure gradient occurs along the body. Therefore, using compression cones is advantageous to study the transition mechanisms. The DNS will be carried out in two parts: simulation of the steady flow behind the bow-shock, and simulation of the unsteady flow behind the bow-shock. The aim of the DNS is to generate the results that are agreeable with Purdue’s laser induced hotspot experiment for compression cones.

Momentum and Thermal Boundary Layer Development on an Internally Cooled Turbine Vane

Abstract: Recent advances in computing power have made conjugate heat transfer simulations of turbine components increasingly popular; however, limited experimental data exists with which to evaluate these simulations. The primary parameter used to evaluate simulations is often the external surface temperature distribution, or overall effectiveness. In this paper, the overlying momentum and thermal boundary layers at various streamwise positions around a conducting, internally cooled simulated turbine vane were measured under low (Tu = 0.5%) and high (Tu = 20%) freestream turbulence conditions. Furthermore, experimental results were compared to computational predictions. In regions were a favorable pressure gradient existed, the thermal boundary layer was found to be significantly thicker than the accompanying momentum boundary layer. Elevated freestream turbulence had the effect of thickening the thermal boundary layer much more effectively than the momentum boundary layer over the entire vane. This data is valuable in understanding the conjugate heat transfer effects on the vane as well as serving as a tool for computational code evaluation.© 2010 ASME

Supersonic Flow Over Straight Shallow Cavities Including Leading and Trailing Edge Modifications

Abstract: An experimental investigation was conducted to examine the effects of adding serrated edges to shallow straight cavities at a freestream Mach number of 2. Measurements consisted of mean and time resolved schlieren visualisation and unsteady pressure spectra. The length to depth ratio of the cavity was 8. The tests conducted at different Reynolds numbers with the straight cavity showed that increasing the Reynolds number increases the root mean square pressure values inside the cavity. Addition of serrations to the cavity leading edge or the trailing edge did not have any significant effect on the separating shear layer nor in stabilising the oscillations of the shear layer. There was also no noticeable effect on the overall sound pressure levels (OASPL) inside the cavity.

Local non-similarity solutions for a forced-free boundary layer flow with viscous dissipation

Abstract: The boundary layer flow over a horizontal plate with power law variations in the freestream velocity and wall temperature of the form Ue ∼ xn and Tw − T∞ ∼ xm and with viscous dissipation, is studied. The boundary layer equations are transformed to a dimensionless system of equations using a non–similarity variable ξ(x) and a pseudo-similarity variable η (x, y). The effects of the various parameters of the flow on velocity and temperature distribution in the boundary layer, on the local skin friction and local heat transfer coefficients and on the non–similar terms, are investigated.