Showing papers on "Freestream published in 2005"

Review of Unsteady Vortex Flows over Slender Delta Wings

Abstract: Nomenclature AR = aspect ratio; amplitude ratio B = probability density function of velocity c = root chord length f = frequency k = reduced frequency; axial wave number n =w ave number in angular direction P = probability p = pressure fluctuation Re =R eynolds number based on chord length S = spectral density s = local semispan T = period t = time U∞ = freestream velocity u = axial velocity v = swirl velocity x = streamwise distance xbd = breakdown location y = spanwise distance z =v ertical distance above wing surface α = angle of attack � = circulation δ = flap angle � = sweep angle ν = kinematic viscosity τ = time constant � =fi nangle ω =v orticity; radial frequency

A Critical Assessment of Reynolds Analogy for Turbine Flows

Abstract: The application of Reynolds analogy (2St/c f ≅1) for turbine flows is critically evaluated using experimental data collected in a low-speed wind tunnel. Independent measurements of St and c f over a wide variety of test conditions permit assessments of the variation of the Reynolds analogy factor (i.e., 2St/c f ) with Reynolds number, freestream pressure gradient, surface roughness, and freestream turbulence. While the factor is fairly independent of Reynolds number, it increases with positive (adverse) pressure gradient and decreases with negative (favorable) pressure gradient. This variation can be traced directly to the governing equations for momentum and energy which dictate a more direct influence of pressure gradient on wall shear than on energy (heat) transfer. Surface roughness introduces a large pressure drag component to the net skin friction measurement without a corresponding mechanism for a comparable increase in heat transfer. Accordingly, the Reynolds analogy factor decreases dramatically with surface roughness. Freestream turbulence has the opposite effect of increasing heat transfer more than skin friction, thus the Reynolds analogy factor increases with turbulence level. Physical mechanisms responsible for the observed variations are offered in each case. Finally, synergies resulting from the combinations of pressure gradient and freestream turbulence with surface roughness are evaluated.

An Experiment of Two Dimensional Roughness Effect on Hypersonic Boundary-Layer Transition

Abstract: tripping method of the hypersonic boundary layer is conducted at a freestream Mach number of 7.1 and a wide range of stagnation conditions with a 5 deg half angle sharp cone at Hypersonic Wind Tunnel #1, JAXA. Aerodynamic heating distributions and fluctuations of the surface pressure are measured both with and without roughness elements. A wavy-wall roughness with a wave length of 2 99%, intending to introduce ‘right’ frequency disturbances into the boundary layer, resulted in delaying the transition at nominal conditions instead of expediting it. Comparisons in power spectrum densities of the surface pressure fluctuations between test conditions as well as roughness types were made to discuss the roll of the wavy-wall roughness. Based on the accurate repeatability results obtained in the present tests, it is concluded that the wavy-wall roughness has a delaying eect on transition at a specific condition.

Numerical Study of Resonant Interactions and Flow Control in a Canonical Separated Flow

Abstract: A novel numerical configuration has been devised in order to investigate active control of separated airfoil flows in a comprehensive and systematic manner. The configuration consists of a flat plate at zero degrees angle-of-attack in a freestream on which a separation bubble of prescribed size is created at a prescribed location through blowing and suction on the top boundary of the computational domain. Numerical simulations of this configuration show that these canonical separated airfoil flows exhibit three distinct characteristic time scales corresponding to the shear layer, the separation zone and the wake vortex shedding. The vortex dynamics associated with these distinct phenomena are described. Preliminary simulations of this flow subjected to zero-net-mass-flux perturbation are also presented.

Separation-Bubble-Transition Measurements on a Low-Re Airfoil Using Particle Image Velocimetry

Abstract: This paper presents experimental results on separation-bubble transition at low Reynolds number and low freestream turbulence, measured on an airfoil using particle image velocimetry (PIV). The two-dimensional PIV measurements have been performed over the suction surface of a low-Reynolds-number airfoil in a water tow-tank facility. Reynolds numbers, based on airfoil chord length and towing speed, of 40,000 and 65,000 have been examined at various angles of incidence, providing a range of streamwise pressure distributions and transitional separation-bubble geometries. The types of bubbles observed range from a short and thick bubble with separation near the leading edge of the airfoil, to a long and thin bubble with separation far downstream of the suction peak. The PIV measurements facilitate visualization of the vortex dynamics associated with separation-bubble transition. The growth of instability waves within the separated shear layer and eventual breakdown into turbulence is documented through the instantaneous vector fields. For all cases examined, large-scale vortex shedding and multiple reverse-flow zones are observed in the reattachment region. A technique for estimating the location of transition onset based on statistical turbulence quantities is presented, and comparisons are made to existing transition models.Copyright © 2005 by ASME

Separated Shock-Boundary-Layer Interaction Control Using Streamwise Slots

Abstract: Experiments have been performed to assess the potential of discretely placed arrays of streamwise slots to control a separated normal shock wave‐turbulent boundary-layer interaction. The supersonic blowdown wind tunnel was operated at a Mach number of 1.5 and a freestream Reynolds number of 26 × × 10 6 m −1 .A tM = 1.5 slot control bifurcated the shock to give a λ shock structure that was significantly larger than that seen without control. The effect on the shock was fairly two-dimensional and persisted in the region between control devices, showing that three-dimensional control devices can have a global effect on the shock structure. In addition, slot control altered the nature of the separated boundary layer from a two-dimensional separation bubble to give highly three-dimensional regions of attached and separated flow. There is evidence that slot control also introduced streamwise vortices, which may help delay or prevent downstream separation.

The behaviour of circular synthetic jets in a laminar boundary layer

Abstract: Dye flow visualisation of circular synthetic jets was carried out in laminar boundary layers developing over a flat plate at a range of actuator operating conditions and freestream velocities of 0.05 and 0.1ms -1 . The purpose of this work was to study the interaction of synthetic jets with the boundary layer and the nature of vortical structures produced as a result of this interaction. The effects of Reynolds number (Re), velocity ratio (V R ) and Strouhal number (St) on the behaviour of synthetic jets were studied. At low Re and V R , the vortical structures produced by synthetic jets appear as highly stretched hairpin vortices attached to the wall. At intermediate Re and V R , these structures roll up into vortex rings which experience a considerable amount of tilting and stretching as they enter the boundary layer. These vortex rings will eventually propagate outside the boundary layer hence the influence of the synthetic jets on the near wall flow will be confined in the near field of the jet exit. At high Re and V R , the vortex rings appear to experience a certain amount of tilting but no obvious stretching. They penetrate the edge of the boundary layer quickly, producing very limited impact on the near wall flow. Hence it is believed that the hairpin vortices produced at low Re and V R are likely to be the desirable structures for effective flow separation control. In this paper, a vortex model was also described to explain the mechanism of vortex tilting.

The Laminar Boundary Layer over a Permeable Wall

Abstract: An analysis is given of the laminar boundary layer over a permeable/porous wall. The porous wall is passive in the sense that no suction or blowing velocity is imposed. To describe the flow inside and above the porous wall a continuum approach is employed based on the Volume-Averaging Method (S. Whitaker The method of volume averaging). With help of an order-of-magnitude analysis the boundary-layer equations are derived. The analysis is constrained by: (a) a low wall permeability; (b) a low Reynolds number for the flow inside the porous wall; (c) a sufficiently high Reynolds number for the freestream flow above the porous wall. Two boundary layers lying on top of each other can be distinguished: the Prandtl boundary layer above the porous wall, and the Brinkman boundary layer inside the porous wall. Based on the analytical solution for the Brinkman boundary layer in combination with the momentum transfer model of Ochoa-Tapia and Whitaker (Int. J. Heat Mass Transfer 38 (1995) 2635). for the interface region, a closed set of equations is derived for the Prandtl boundary layer. For the stream function a power series expansion in the perturbation parameter κ is adopted, where κ is proportional to ratio of the Brinkman to the Prandtl boundary-layer thickness. A generalization of the Falkner–Skan equation for boundary-layer flow past a wedge is derived, in which wall permeability is incorporated. Numerical solutions of the Falkner–Skan equation for various wedge angles are presented. Up to the first order in κ wall permeability causes a positive streamwise velocity at the interface and inside the porous wall, but a wall-normal interface velocity is a second-order effect. Furthermore, wall permeability causes a decrease in the wall shear stress when the freestream flow accelerates, but an increase in the wall shear stress when the freestream flow decelerates. From the latter it follows that separation, as indicated by zero wall shear stress, is delayed to a larger positive pressure gradient.

Skin Friction Correlation for Smooth and Rough Wall Turbulent Boundary Layers

Abstract: In this paper, we propose a novel skin friction correlation for a zero pressure gradient turbulent boundary layer over surfaces with different roughness characteristics. The experimental data sets were obtained on a hydraulically smooth and ten different rough surfaces created from sand paper, perforated sheet, and woven wire mesh. The physical size and geometry of the roughness elements and freestream velocity were chosen to encompass both transitionally rough and fully rough flow regimes. The flow Reynolds number based on momentum thickness ranged from 3730 to 13,550. We propose a correlation that relates the skin friction, Cf, to the ratio of the displacement and boundary layer thicknesses, δ*∕δ, which is valid for both smooth and rough wall flows. The results indicate that the ratio Cf1∕2∕(δ*∕δ) is approximately constant, irrespective of the Reynolds number and surface condition.

Laminar Boundary-Layer Instabilities on Hypersonic Cones: Computations for Benchmark Experiments

Abstract: : Although significant advances have been made in hypersonic boundary-layer transition prediction in the last several decades, most design work still relies on unreliable empirical correlations or wind-tunnel tests. Codes using the semi-empirical eN method will need to be well verified and validated before being used for expensive flight vehicles. The code package STABL and its PSE-Chem stability solver are used to compute first and second-mode instabilities for both sharp and blunt cones at wind-tunnel conditions using a Xavier-Stokes mean-flow solution. Computations are performed for Stetson's 3.81 mm nose-radius cone, a sharp cone at Mach 3.5, a large-bluntness cone at Mach 8, and sharp and blunt cones corresponding to the experiments of Rufer. Comparisons to previous computations by other researchers show differences on the order of 10% in local amplification rates and frequencies, but better agreement is obtained for the transition location. Many issues are examined for verification and validation, including the laminar transport properties, the freestream boundary conditions, and the effect of freestream thermal nonequilibrium. This work helps to verify and validate STABL, extend its applicability to low-temperature flows, and develop the methodologies for using STABL.

Rapid Engineering Approach to Modeling Hypersonic Laminar-To-Turbulent Transitional Flows

Abstract: Modeling the transitional behavior of hypersonic flows from an engineering point of view is addressed. More specifically, the accuracy of transition onset predictions by using a modified transition onset transport model is examined. Transitional blending is predicted by using the well-established Dhawan and Narasimha algebraic model. To employ the rapid engineering approach, the methodology is encompassed within the So, Sarkar, Gerodimos, and Zhang low-Reynolds-number k‐e turbulence model framework that has been extended to include compressibility effects. Model calibration and validation is achieved by using fundamental flat-plate data sets of Mee as well as recent data sets obtained by Holden for scramjet forebody geometries. It is shown that the transition onset transport equation predicts onset in accord with experimental values by using freestream velocity fluctuation levels consistent with those of the shock tunnel facilities. An additional simulation of the reentry F vehicle under low noise flight conditions was also performed. Under these conditions, the onset model predicts the location of transition very close to that measured. Overall, the framework developed can be used for very rapid turnaround engineering applications. The model can account for both freestream noise as well as amplification effects caused by shock/boundary interactions, specifically those dealing with compression corners.

Planar Velocity and Temperature Measurements in Rarefied Hypersonic Flow Using Iodine LIF

Abstract: A planar laser-induced fluorescence (PLIF) technique is discussed and applied to measurement of time-averaged values of velocity and temperature in an I2-seeded N2 hypersonic free jet facility. A low temperature, non-reacting, hypersonic flow over a simplified model of a reaction control system (RCS) was investigated using the PLIF technique. Data are presented of rarefied Mach 12 flow over a sharp leading edge flat plate at zero incidence, both with and without an interacting jet issuing from a nozzle built into the plate. The velocity profile in the boundary layer on the plate was resolved. The slip velocity along the plate, extrapolated from the velocity profile data, varied from nearly 100% down to 10% of the freestream value. These measurements are compared with results of a DSMC solution. The velocity variation along the centerline of a jet issuing from the plate was measured and found to match closely with the correlation of Ashkenas and Sherman. The velocity variation in the oblique shock terminating the jet was resolved sufficiently to measure the shock wave thickness. I. Iodine fluorescence measurement technique Numerous methods have been reported in the literature for measurement of thermodynamic properties in supersonic flows using laser-induced fluorescence. This paper presents results of a technique based on narrowband excitation of fluorescence from nuclear hyperfine components in the B↔X electronic transition of molecular iodine. This approach was developed to study low temperature flows of I2-seeded N2 gas. A tunable single frequency laser beam is expanded and collimated into a thin sheet and passed into a low density wind tunnel flow, causing I2 molecules to fluoresce within a selected planar section of the flow-field. A series of images are taken of the fluorescence with a charge-coupled device (CCD) array as the laser frequency is incrementally tuned over selected lines in the I2 absorption spectrum. The fluorescence signal versus frequency is thereby obtained for any point in the plane of the laser sheet, within the resolution of the digital image. These absorption spectra can provide information on the local thermodynamic conditions in the flow. This paper describes planar laser-induced fluorescence (PLIF) measurements of velocity and temperature made in a low density, low temperature, hypersonic flow over a simplified model of a reaction control system (RCS) jet, 1 intended to support assessment of a hybrid numerical solution method for such flows. 2, 3

Fluid Structure Interaction on a Generic Body-Flap Model in Hypersonic Flow

Abstract: A numerical investigation of a generic reentry vehicle flap configuration at hypersonic speed is presented, which includes considerations of the generic hinge line gap and the outer edges of the flap. Numerical solutions were obtained with a Navier–Stokes code and were coupled via a surface interpolation routine to a structural solver. Two- and three-dimensional solutions were compared for the generic flap model, consisting of a forebody and a body flap deflected 20 deg relative to the forebody with a gap between them. Presented are computational results for a freestream Mach number of 7.3 and an angle of attack of 15 deg with both open- and closed-gap configurations, which are compared to available experimental data. Numerical error, either from the code or the grid, is assessed with the impact of any error on the magnitude of the heat fluxes and the associated gradients. The three-dimensional solutions predict that a strong outflow from the center of the model to the outer edges takes place over the model surface, impacting the flow topology significantly. This was verified with comparison to the experimental results. Temperature peaks predicted by stand-alone computational fluid dynamics solutions were not observed in the experimental data. The investigations confirmed that this is due to strong coupling effects between fluid and structure. Heat conduction into and inside the structure leads to significantly reduced surface temperatures in critical regions.

Experiments With Three-Dimensional Passive Flow Control Devices on Low-Pressure Turbine Airfoils

Abstract: The effectiveness of three dimensional passive devices for flow control on low pressure turbine airfoils was investigated experimentally. A row of small cylinders was placed at the pressure minimum on the suction side of a typical airfoil. Cases with Reynolds numbers ranging from 25,000 to 300,000 (based on suction surface length and exit velocity) were considered under low freestream turbulence conditions. Streamwise pressure profiles and velocity profiles near the trailing edge were documented. Without flow control a separation bubble was present, and at the lower Reynolds numbers the bubble did not close. Cylinders with two different heights and a wide range of spanwise spacings were considered. Reattachment moved upstream as the cylinder height was increased or the spacing was decreased. If the spanwise spacing was sufficiently small, the flow at the trailing edge was essentially uniform across the span. The cylinder size and spacing could be optimized to minimize losses at a given Reynolds number, but cylinders optimized for low Reynolds number conditions caused increased losses at high Reynolds numbers. The effectiveness of two-dimensional bars had been studied previously under the same flow conditions. The cylinders were not as effective for maintaining low losses over a range of Reynolds numbers as the bars.Copyright © 2005 by ASME

The Structure of Shock Cell Noise from Supersonic Jets

Abstract: Shock cell noise is created by supersonic jets operating at off-design conditions due to the interaction between the evolving turbulence and the quasi-fixed shock cell structure. Transonic wind tunnel tests are conducted to identify the structure of the noise produced by this process in the presence of external flow. The data is compared to a simple kinematic model that incorporates the freestream velocity. It predicts the relationship between frequency and radiation angle and also predicts the screech frequency as the limiting case of upstream propagation. Results indicate that the freestream flow has a significant effect on the noise production. The noise level decreases and the radiation angle rotates toward the upstream axis as the external flow increases. But, when convective effects are considered the direction of noise is 90 o to 100 o for all conditions.

Large-structure topology in a three-dimensional supersonic base flow

Abstract: The near wake of a blunt-base cylinder at 10-deg angle of attack to a Mach 2.46 freestream flow is studied using Mie scattering flow visualization. Large-scale structures are clearly evident in side-view images of the windward and lateral parts of the shear layer but not in the leeward shear layer

The effect of wake parameters on the transitional boundary layer on turbine blade

Abstract: This paper presents an experimental analysis of the interaction between wakes and boundary layers on an aerodynamic blade profile. The main objective of this research was to investigate the simultaneous influence of the freestream conditions and wake parameters on the wake-induced transition process. The investigations were performed for two levels of inlet freestream turbulence Tu = 0.4 and 3 per cent. The wakes were generated by cylindrical bars mounted in the rotating wheel. It was shown that wake parameters and especially the turbulence in the wake and wake width have a great influence on the position and extent of the induced transition. The role of freestream turbulence on the wake structure passing over the blade surface was also emphasized.

Temporal Linear Stability Analysis of Three-Dimensional Compressible Binary Shear Layers

Abstract: Temporal linear stability analysis is performed on the three-dimensional compressible binary shear layer for density ratios ranging from 1 to 32 and convective Mach numbers up to 2. Comparison is made with the results of spatial theory. Some stability properties are shared by both approaches, for example, the independence of the direction of propagation of the most unstable modes from the freestream density ratio, but other features are different: a single secondary mode appears instead of two at high compressibility. An empirical model is proposed for the wave length and phase speed of the most amplified waves. Direct numerical simulations show good agreement with linear inviscid theory for initial Reynolds numbers greater than 400.

Mach number and freestream turbulence effects on turbine vane aerodynamic losses

Abstract: The aerodynamic performance of a smooth. cambered turbine vane (which replicates one used in an operating gas turbine engine) is investigated in this paper. Three different Mach-number distributions are employed, which result in one transonic flow and two subsonic flows. All of these distributions match flow conditions in different industrial applications. A fine mesh grid and cross bars are used to augment the magnitudes of longitudinal turbulence intensity at the inlet of the test section. Wake-profile data are presented for two different locations downstream of the vane trailing edge (one axial chord length and 0.25 axial chord length). The contributions of varying Mach number and varying freestream turbulence intensity to aerodynamic losses, normalized kinetic energy profiles, normalized Mach-number profiles, integrated aerodynamies losses, and area-averaged loss coefficients are quantified. Results show that wake profiles are more sensitive to turbulence intensity variations at lower subsonic flow conditions than when transonic flow is present. Wake profiles are also broadened either as the exit Mach number increases or as the freestream turbulence intensity level increases. Higher losses in the freestream flow are present as the inlet turbulence intensity level increases. Also described are effects of increased turbulent diffusion, streamwise development, and profile asymmetry. Corresponding integrated aerodynamic losses and area-averaged loss coefficient Y A magnitudes increase with increasing Mach number or with increasing turbulence intensity level. Results additionally show larger loss magnitudes with flow turning and cambered airfoils, relative to symmetric airfoils, when compared at the same exit Mach number.

Planar Laser‐Induced Iodine Fluorescence Measurements in Rarefied Hypersonic Flow

Abstract: A planar laser-induced fluorescence (PLIF) technique is discussed and applied to measurement of time-averaged values of velocity and temperature in an I(sub 2)-seeded N(sub 2) hypersonic free jet facility. Using this technique, a low temperature, non-reacting, hypersonic flow over a simplified model of a reaction control system (RCS) was investigated. Data are presented of rarefied Mach 12 flow over a sharp leading edge flat plate at zero incidence, both with and without an interacting jet issuing from a nozzle built into the plate. The velocity profile in the boundary layer on the plate was resolved. The slip velocity along the plate, extrapolated from the velocity profile data, varied from nearly 100% down to 10% of the freestream value. These measurements are compared with results of a DSMC solution. The velocity variation along the centerline of a jet issuing from the plate was measured and found to match closely with the correlation of Ashkenas and Sherman. The velocity variation in the oblique shock terminating the jet was resolved sufficiently to measure the shock wave thickness.

The Effects of Elevated Freestream Turbulence on Separation Control with Vortex-Generating Jets

Abstract: Vortex-generating jets (VGJs) have proven to be an effective form of active separation control on low-pressure turbine blades in low freestream turbulence flows. Twocomponent flow velocity measurements were taken in a low-speed linear cascade to assess the effectiveness of VGJs with highly turbulent freestream conditions. Blade losses associated with flow separation were evaluated using a boundary layer momentum flux loss parameter which compared favorably with prior experiments performed in a larger cascade facility using the traditional wake total pressure loss coefficient. Losses were calculated for freestream turbulence levels of 0.4%, 3%, 6%, and 10%, and steady jet blowing ratios (B = jet velocity/freestream velocity) of 0, 2, and 4 at a Reynolds number of 25,000. At the lowest freestream turbulence level of 0.4%, the separation losses decreased by 57% and 51% for jet blowing ratios of 2 and 4 respectively. Increasing the freestream turbulence level monotonically reduced the effectiveness of the VGJs. At the highest freestream turbulence level of 10%, the decreases in loss were 18% and 6% for blowing ratios of 2 and 4 respectively. This decrease in the effectiveness of VGJs at high turbulence levels can be attributed to a more rapid transition from a laminar to a turbulent boundary layer on the blade suction surface and a more rapid dissipation of the vortices induced by the VGJs. The lower jet blowing ratio (B=2) was up to three times more effective than the high blowing ratio (B=4) with elevated freestream turbulence, due to increased jet penetration at the higher blowing ratio. Nomenclature

Aerothermodynamics of the gas giants

Abstract: The aerothermodynamics of the gas giant planets was studied in a superorbitalnexpansion tube. A mixture of H2/Ne was used as a substitute to the Jovian atmosphere ofnH2/He, which allowed Jovian entry aerothermodynamics to take place at flight speeds ofn15 km/s, whilst still maintaining chemical similarity. Results for shock standoff distance andnelectron concentration along the stagnation streamline of a blunt body demonstrated thatnnonequilibrium effects occurring in the vehicle shock layer during Jovian entry can benexperimentally simulated. Two-wavelength and near-resonant holographic interferometrynwere used to obtain experimental results.n n A one-dimensional inviscid analytical model was developed to model shock layernproperty distribution along the stagnation line of a blunt body in nonequilibrium hypersonicnflow. Shock standoff was calculated from the average density along the post-shock stagnationnline and controlling non-dimensional flow parameters were identified. The analytical modelnwas validated for dissociating flow using published theoretical, experimental and CFD shocknstandoff results. For ionising flow, experimental results for shock standoff and electronnconcentration distribution were compared with predictions from the theoretical model. Goodnagreement was obtained for shock standoff results using near-resonant holographicninterferometry, however, for two-wavelength holographic interferometry, experimental resultsnwere consistently lower than predictions, primarily due to low resolution of density variationnin the shock layer. Shock standoff distances modified for the effect of low density resolutionnshowed significantly improved agreement with both analytical and CFD results. Thenexperimental electron concentration distribution was reasonably consistent with the analyticalnmodel. Radiative cooling of the shock layer, three-dimensional cylinder end effects andnuncertainty in reaction rate coefficients were identified as second-order effects that mayncontribute to differences between predictions made using the analytical model and CFD, andnexperimental results.n n New scaling parameters have emerged during analysis that can be used to simplifynthe experimental simulation of hypersonic nonequilibrium dissociating and ionising gasnmixtures. Flow properties along the shock layer stagnation line were found to dependnprimarily on the chemical scaling parameter for highly nonequilibrium flight, regardless ofnindividual values of freestream density, freestream velocity, freestream dissociation, bodynradius and degree of gas dilution.n

Shock Interaction Control for Scramjet Cowl Leading Edges

Abstract: An experimental study was conducted to qualitatively determine the effectiveness of stagnation-region gas injection in protecting a scramjet cowl leading edge from the intense heating produced by Type III and Type IV shock interactions. The model consisted of a two-dimensional leading edge, representative of that of a scramjet cowl. Tests were conducted at a nominal freestream Mach number of 6. Gaseous nitrogen was supersonically injected through the leading-edge nozzles at various mass flux ratios and with the model pitched at angles of 0deg and -20deg relative to the freestream flow. Qualitative data, in the form of focusing and conventional schlieren images, were obtained of the shock interaction patterns. Results indicate that large shock displacements can be achieved and both the Type III and IV interactions can be altered such that the interaction does not impinge on the leading edge surface.

Instability and Bifurcation of Transonic Flow Over Airfoils

Abstract: An investigation of inviscid transonic flow over a flattened airfoil whose curvature is small in the midch ord region is performed. Numerical simulation reveals a bifurcation of steady solutions to the Euler equations in a range of freestream Mach numbers �� and angles of attack � . The dependence of the lift coefficient CL on �� andis studied in the intervals �1.2 < �, deg < 1.2 , 0.79 < �� < 0.85. A theoretical interpretation of the bifurcation and instability phenomena is presented. Nomenclature CL = lift coefficient � = angle of attack �� = freestream Mach number �s = singular Mach number �min = left endpoint of the bifurcation interval at �=0 �max = right endpoint of the bifurcation interval at �=0 x,y = Cartesian coordinates I. Introduction ECENT studies of inviscid transonic flow over a flattened bump in a channel 1-3 have revealed an instability associated with the splitting/merging of local supersonic regions over the bump. The instability was found to trigger an abrupt change of the structure of steady flow under a small perturbation of boundary conditions. The concept of structural instability has shed some light on the nonuniqueness of inv iscid flow over airfoils demonstrated earlier by Jameson , 4 McGrattan , 5 and Hafez & Guo. 6,7 A detailed analysis of the link between the non -uniqueness and instability was performed by Kuz'min & Ivanova 8,9 and Semyonov 10 for the symmetric airfoil

Characterization of Flow Structures and Turbulence in Hypersonic Jet Interaction Flowfields.

Abstract: Numerical simulations of normal sonic gaseous injection through diamond shaped orifices into a Mach 5.0 freestream were performed using RANS and DES turbulence models. The GASP and Cobalt flow solvers were used in this work. A detailed study of the secondary turbulent flow structure was performed. Flow structures predicted by RANS solutions were compared with instantaneous and time-averaged DES solutions. Results indicate that RANS and DES predict similar flow structures. Turbulent kinetic energy results show higher predictions by DES in the downstream shear layer and lower prediction near the injector as compared to RANS model results. Also, clear connection to secondary flow structure was observed; for example, a TKE peak was observed downstream of the injector and was concluded to be due to the interaction of the wake vortices.

The Boundary Layer Over Turbine Blade Models With Realistic Rough Surfaces

Abstract: Results are presented of extensive boundary layer measurements taken over a flat, smooth plate model of the front one-third of a turbine blade and over the model with an embedded strip of realistic rough surface. The turbine blade model also included elevated freestream turbulence and an accelerating freestream in order to simulate conditions on the suction side of a high-pressure turbine blade. The realistic rough surface was developed by scaling actual turbine blade surface data provided by U.S. Air Force Research Laboratory. The rough patch can be considered to be an idealized area of distributed spalls with realistic surface roughness. The results indicate that bypass transition occurred very early in the flow over the model and that the boundary layer remained unstable (transitional) throughout the entire length of the test plate. Results from the rough patch study indicate the boundary layer thickness and momentum thickness Reynolds numbers increased over the rough patch and the shape factor increased over the rough patch but then decreased downstream of the patch. It was also found that flow downstream of the patch experienced a gradual retransition to laminar-like behavior but in less time and distance than in the smooth plate case. Additionally, the rough patch caused a significant increase in streamwise turbulence intensity and normal turbulence intensity over the rough patch and downstream of the patch. In addition, the skin friction coefficient over the rough patch increased by nearly 2.5 times the smooth plate value. Finally, the rough patch caused the Reynolds shear stresses to increase in the region close the plate surface.

Aerodynamics of Low-Blowing-Ratio Fuselage Injection into a Supersonic Freestream

Abstract: A parametric characterization of the surface flow structure and associated aerodynamic loading for injection into a Mach 3.0 freestream through jets mounted into the side of missile fuselage (sharp-coned cylinder) was experimentally achieved. The experimental methods included surface oil flow visualization, shadowgraph photography, and pressure-sensitive paint. The range of jet-to-freestream pressure ratios spanned 0.6-19.0, which corresponds to an effective backpressure ratio range of 0.19-6.15. Hence, overexpanded, perfectly expanded, and underexpanded jets were tested. The temperature ratios spanned 1.0-13.6, where helium was used to simulate the highest temperature. The nondimensional momentum parameter ratio (MPR) varied over a range of 0.005-0.09. It was observed that the injection temperature for a given MPR had a small effect on the boundary-layer separation distance. For the range of conditions tested, the interaction force increased linearly with increasing MPR. The amplification factors for the lowest MPR values were found to be higher than expected trends.