Showing papers on "Freestream published in 1994"

Suppression of flow-induced pressure oscillations in cavities

Abstract: Experimental methods for suppressing flow-induced pressure oscillations in a shallow cavity resulting from tangential flows over the cavity are described. The effects of manipulating the shear layer over the cavity leading edge are examined. Static and oscillating fences and steady and pulsating flow injection at the leading edge are studied for their effect on cavity sound pressure levels. Both subsonic and supersonic flow conditions are considered. Of the methods tested, static fences at the leading edge were found to provide the most suppression. Suppression was dependent on the frequency mode and the flow Mach number. Nomenclature a = speed of sound D = cavity depth / = frequency k = vortex convection velocity to freestream velocity ratio L = cavity length M = Mach number m = frequency mode number P = pressure P0 = stagnation pressure Re = Reynolds number, Uxlv S = Strouhal number S * = modified Strouhal number U = freestream velocity x = distance from nozzle exit Z = cavity span, lateral dimension a = phase delay parameter y = ratio of specific heats 8 = boundary-layer thickness v = kinematic viscosity

The Effect of High Freestream Turbulence on Film Cooling Effectiveness

Abstract: This study investigated the adiabatic wall cooling effectiveness of a single row of film cooling holes injecting into a turbulent flat plate boundary layer below a turbulent, zero pressure gradient free stream. Levels of free-stream turbulence (Tu) up to 17.4 percent were generated using a method that simulates conditions at a gas turbine combustor exit. Film cooling was injected from a single row of five 35 deg slant-hole injectors (length/diameter = 3.5, pitch/diameter = 3.0) at blowing ratios from 0.55 to 1.85 and at a nearly constant density ratio (coolant density/free-stream density) of 0.95. Film cooling effectiveness data are presented for Tu levels ranging from 0.9 to 17 percent at a constant free-stream Reynolds number based on injection hole diameter of 19,000. Results show that elevated levels of free-stream turbulence reduce film cooling effectiveness by up to 70 percent in the region directly downstream of the injection hole due to enhanced mixing. At the same time, high free-stream turbulence also produces a 50--100 percent increase in film cooling effectiveness in the region between injection holes. This is due to accelerated spanwise diffusion of the cooling fluid, which also produces an earlier merger of the coolant jets from adjacent holes.

Influence of injectant Mach number and temperature on supersonic film cooling

Abstract: The current work is an experimental investigation of the dependence of film cooling effectiveness on the injection Mach number, velocity, and mass flux. The freestream Mach number is 2.4, and the injection Mach numbers range from 1.2 to 2.2 for both air and helium injection. The adiabatic wall temperature is measured directly. The injection velocity and mass flux are varied by changing the total temperature and Mach number while maintaining matched pressure conditions between the injected flow and that of the freestream. The total temperature of the freestream is 295 K, and for the injection it ranges from 215-390 K. The results indicate an increase in film cooling effectiveness as the injection rate is increased. With the exception of heated helium runs, larger injection Mach numbers slightly increase the effective cooling length per mass injection rate. The results for helium injection indicate an increase in effectiveness as compared to that for air injection. Heated injection, with the injectant to freestream velocity ratios greater than 1, exhibit a rise in wall temperature downstream of the slot resulting in effectiveness values greater than 1. The experimental results are also compared with earlier studies in the literature.

Flow alteration and drag reduction by riblets in a turbulent boundary layer

Abstract: Detailed measurements with a miniature single-sensor hot-wire probe of the streamwise velocity field within a riblet groove of dimensionless size h+ ~ 14 and s+ ~ 28 (based on the riblet surface average friction velocity) have been made. The wall shear stress, when integrated over the riblet surface, yields about 4% drag reduction com- pared with a smooth surface with the same projected area. This is largely due to the greatly diminished wall shear stress near the bottom of the riblet valley. Four-sensor hot-wire probe measurements reveal that riblets sig- nificantly reduce the vertical flux of streamwise momentum within the riblet valley. HAT, under certain conditions, configuring the boundary be- neath a turbulent flow into small grooves aligned with the mean flow direction, called riblets, will reduce viscous drag has been well documented for over a decade.1'2 However, very few de- tailed measurements of the flowfield within and just above these riblet grooves have been made for the obvious reason that they are very difficult. In most laboratory air boundary layers the physical dimensions of such riblets are considerably less than 1 mm, and the velocities to be measured are extremely small. Vukoslavcevic et al.3 have, however, measured properties of the streamwise velocity field along the vertical planes of symmetry above the valley and above the peak of a triangular riblet surface in an air boundary layer at RQ ~ 1.0 X 10 3. These measurements showed that the mean ve- locity gradient at the valley surface is much lower than at a compar- ison smooth flat plate surface for flows with the same freestream velocity. The mean velocity gradient over the peak was somewhat greater than that over the smooth surface. Vukoslavcevic et al. 3 also determined that the turbulence intensity of the streamwise ve- locity fluctuations was greatly reduced deep in the riblet valley. The skewness and flatness factors indicated that the turbulence only rarely penetrates deep into the valley from the layers above. These measurements, however, were not made normal to or at a sufficient number of stations along the riblet wall surface to esti- mate whether and, if so, how drag reduction was achieved. Further- more, no measurements of the vertical velocity fluctuations v were made, and so information about the vertical flux of streamwise mo- mentum could only be inferred indirectly. Chu et al. 4 and Choi et al. 5 have carried out direct numerical simulations of flow over riblet walls in channel flows. Most of the conclusions made from the laboratory experiment reported herein are qualitatively consistent with their simulation results, although there are some quantitative differences in detail, as will be dis- cussed later when comparisons are made.

Flowfield measurements in supersonic film cooling including the effect of shock-wave interaction

Abstract: A study has been made to investigate the flowfield of supersonic slot injection and its interaction with a two- dimensional shock wave. Air and helium were injected at Mach numbers of approximately 1.3 and 2.2 into an airstream of Mach 2.4. Measurements of the total pressure profiles perpendicular to the wall were made at sev- eral axial locations, the farthest being at 90 slot heights. The profiles provided details of the structure of the flow for the different injection conditions. With heated gas injection, experiments were conducted to determine the adiabatic wall temperatures and the wall static pressures. These measurements were then repeated with the impingement of two-dimensional shock waves at 60 slot heights downstream of the slot. The shock strengths were chosen to illustrate the differences between separated and attached flows. The shock strength that produced incipient separation was found to be smaller when helium was injected than when no film coolant was present. Conversely, the shock strength that produced incipient separation with air injection was slightly larger than that obtained without film cooling. develop relations for effectiveness as a function of downstream position divided by slot height x/s and the ratio of mass flux for the injected flow to that in the freestream, A, = (pw)//( pu)^. The phys- ical basis for these relations has also been motivated from simple integral analysis of the flowfield.5 The variation in the effectiveness with downstream position is accompanied by changes in flowfield structure. Some subsonic film-cooling experiments6 and other subsonic experiments involv- ing wall jets with moving freestream7 have shown that the flow- field can be divided into three regions: a potential core region, a wall-jet region, and a boundary-layer region. The potential core region, like a freejet, contains a viscous layer that emanates from the lip and ends when it meets the slot-flow boundary layer. In this region the wall temperature remains at a constant value equal to that of the injected fluid (in the case of subsonic injection) or equal to the recovery value (in the case of supersonic injection). Thus, the effectiveness in the potential core region is unity. The wall-jet region starts when the viscous layer emanating from the lip merges with the injectant boundary layer. In this region intense mixing takes place, and the wall temperature increases toward the freestream value. In the boundary-layer region, the flow then relaxes to that of a boundary layer. Consequently, the effectiveness decreases from unity near the injector and approaches zero far downstream. Thus, film-cooling flows combine different types of familiar flows: a freejet flow, a wake, a shear layer, and a bound- ary layer. The different hydrodynamic features of the flow in each region suggest using different scaling laws to predict effectiveness. This approach was attempted with some success in low-speed flow,8 where empirical data from jet flows and boundary-layer flows were applied for near and far regions, respectively. These approxi- mate flowfields were used in the energy equation to solve for the distribution in wall temperature and the film-cooling effectiveness. In these incompressible analyses, the thermodynamic properties were considered invariant within the flowfield.5 The film-cooling effectiveness in high-speed flow is often defined as T| = T -T (2)

Quantitative investigation of compressible mixing - Staged transverse injection into Mach 2 flow

Abstract: Planar measurements of the injectant mole fraction distribution and the velocity field within a supersonic mixing flowfield have been made using laser-induced iodine fluorescence. The flowfield investigated in this work is staged transverse injection of air into a Mach 2 freestream. A complete three-dimensional survey of the injectant mole fraction distribution has been generated, and a single planar velocity measurement has been completed. The measurements reveal the dramatic effect of streamwise vortices on the mixing in the near field of the injectors, as well as the rapid mixing generated by staging two fuel injectors. Analysis of the downstream decay of the maximum injectant mole fraction in this and other supersonic mixing flowfields indicates that the relative rate of injectant mixing well downstream of the injectors is independent of injection geometry, freestream Mach number, and injectant molecular weight. Mixing within this region of the flowfield is dominated by small-scale turbulence within the injectant plume. The transition of the dominant mixing mechanism, from vortex-driven mixing in the near field to small-scale turbulent mixing in the far field, was found to occur in the region about 10 diameters downstream of the injectors.

Physical description of boundary-layer transition: Experimental evidence

Abstract: The problems of understanding the origins of turbulent flow and transition to turbulent flow are the most important unsolved problems of fluid mechanics and aerodynamics. It is well known that the stability, transition, and turbulent characteristics of bounded shear layers are fundamentally different from those of free shear layers. Likewise, the stability, transition, and turbulent characteristics of open systems are fundamentally different from those of closed systems. Because of the influence of indigenous disturbances, surface geometry and roughness, sound, heat transfer, and ablation, it is not possible to develop general prediction schemes for transition location and the nature of turbulent structures in boundary-layer flows. At the present time no mathematical model exists that can predict the transition Reynolds number on a flat plate. The recent progress in this area is encouraging, in that a number of distinct transition mechanisms have been found experimentally. The theoretical work finds them to be amplitude and Reynolds-number dependent. The theory remains rather incomplete with regard to predicting transition. Amplitude and spectral characteristics of the disturbances inside the laminar viscous layer strongly influence which type of transition occurs. The major need in this area is to understand how freestream disturbances are entrained into the boundary layer, i.e., to answer the question of receptivity. We refer receptivity to the mechanism(s) that cause freestream disturbances to enter the boundary layer and create the initial amplitudes for unstable waves.

Generating high freestream turbulence levels

Abstract: This paper presents a method of generating a highly turbulent freestream flow, up to levels of 20% with a relatively uniform mean velocity field. This method was developed as a result of a combined water channel and wind tunnel study. The method for generating these high turbulence levels includes using high-velocity jets issuing into a mainstream cross-flow. A range of turbulence levels can be generated, using this same flow geometry, by adjusting the jet-to-mainstream velocity ratio or the Reynolds number of the flow.

Analysis of thermal ignition in the supersonic mixing layer

Abstract: Ignition in a laminar supersonic mixing layer between two parallel streams of initially separated reactants is studied both numerically and through the use of large activation energy asymptotics. The asymptotic analysis provides a description of ignition characteristics over the entire range of system parameters. In particular, it is demonstrated that, for small values of viscous heating, the ignition distance scales approximately linearly with the freestream Mach number, whereas for large viscous heating it decreases rapidly due to the temperature-sensitive nature of the reaction rate. This indicates the potential of using local flow retardation to enhance ignition rather than relying solely on external heating. The asymptotic analysis further identifies several distinct ignition situations, yielding results that compare well with those obtained from the full numerical calculation. The effects of flow nonsimilarity are also assessed and are found to be more prominent for the mixing layer flow in comparison to the flat-plate configuration studied previously.

Experimental investigation of the flowfield about an unswept afterbody

Abstract: In general, the flowfield about the aft portion of an aircraft fuselage employing an upswept afterbody is complex and can have a detailed vortex structure. Directional pressure probe measurements show that the afterbody wake evolution is weakly dependent on the Reynolds number. Pressure taps are used to investigate the effect of base slant on the base pressure distribution. The base pressure distribution is found to increase along the upswept ramp in the freestream direction. Using oil flow visualizations general conclusions are drawn about the surface flow topology. Two distinct regions of flow separation are identified. An analytical model of the crossflow wake structure is developed from the experimental data. This model allows for conclusions to be drawn about the crossflow vortex structure and provides simple analytical expressions for the vorticity and swirl velocity distributions in the crossflow plane.

Computational/experimental investigation of staged injection into a Mach 2 flow

Abstract: The stage normal injection of two jets of air into a Mach 2 freestream behind a rearward-facing step has been investigated using laser-induced iodine fluorescence and laser doppler anemometry techniques. A detailed data set has been compiled that includes profiles of pressure, temperature, two components of velocity, and mole fraction of the injectant at 44 locations, plus planar surveys of pressure, temperature, velocity, and mole fraction. A companion numerical study was performed using the SPARK Navier-Stokes code. Good overall agreement was observed for this complex, highly three-dimensional flowfield. Discrepancies were observed in the computed and measured total temperatures, turbulent mixing, and shock strengths. The effect of grid resolution was investigated by calculating solutions on grids of 60,000 points, 200,000 points, and 450,000 points. Differences in the solutions on the two finer grids were small. The effect of turbulence modeling was investigated by calculating solutions with three different algebraic models for the jet turbulence. Overall, the turbulence models were found to have the greatest effect on the numerical solutions, followed by the grid resolution and the injectant Mach number.

Drag reduction of airplane fuselages through shaping by the inverse method

Abstract: The axial singularity solution for the axisymmetric inverse problem has been extended to utilize doublet elements with linear intensity distribution. The solution converges faster than the source-based method, and is therefore quite promising. A procedure based on this solution has been used to design low-drag laminar fuselage shapes for small aircraft applications with a volumetric Reynolds number range of 10-30 million. A profile with a fineness ratio of 6, transition at 40% of body length, and volumetric drag coefficient of 0.012 at a nominal /?v of 15 million, has been developed. The present inverse procedure was shown to be a powerful alternative to optimization methods. Several transition criteria were investigated in the course of the study. The Crabtree criterion appears to be the most consistent. Experimental transition data for axisymmetric bodies at high (flight) Reynolds numbers are urgently needed. Nomenclature CDV = volumetric drag coefficient based on V 2/3 as characteristic area /,. = fineness ratio, body length/maximum diameter H = boundary-layer shape factor, 8*/9 L = body length / = length of an axial-singularity element n = number of doublet elements q = local velocity on body surface (at the edge of the boundary layer) RL = length Reynolds number based on UM and L Rs = Reynolds number based on local velocity and s Rx = Reynolds number based on local velocity and jc Re = Reynolds number based on local velocity and 9 7?v = volumetric Reynolds number based on Ux and V1/3 r. — local body radius and radial coordinate 5 = length along body surface starting at nose Ux = freestream velocity u = axial component of q v = radial component of q X = nondimensional axial coordinate, x/L

End-Stage Boundary Layer Transition Models for Engineering Calculations

Abstract: The paper presents some recent new data on the combined effects of pressure gradient and freestream turbulence level on the onset and length of the latter stages of the boundary layer transition process. Generalized correlations for the transition length Reynolds number are developed from considerations of the non-dimensional turbulent spot formation rate. The optimized correlation is built into a popular linear combination integral computer code to predict the growth of the transitional boundary layer in a number of practical engineering flows.

Mean and turbulent velocity measurements of supersonic mixing layers

Abstract: The behavior of supersonic mixing layers under three conditions has been examined by schlieren photography and laser Doppler velocimetry. In the schlieren photographs, some largescale, repetitive patterns were observed within the mixing layer; however, these structures do not appear to dominate the mixing layer character under the present flow conditions. It was found that higher levels of secondary freestream turbulence did not increase the peak turbulence intensity observed within the mixing layer, but slightly increased the growth rate. Higher levels of freestream turbulence also reduced the axial distance required for development of the mean velocity. At higher convective Mach numbers, the mixing layer growth rate was found to be smaller than that of an incompressible mixing layer at the same velocity and freestream density ratio. The increase in convective Mach number also caused a decrease in the turbulence intensity (σ u /ΔU).

Role of Splitter Plates in Modifying Cylinder Wake Flows

Abstract: The effects of a splitter plate as well as the gap g between the splitter plate and the trailing edge of a cylinder on the frequency characteristics of the cylinder were investigated over a Reynolds number ReD range of 1.2 x 104 to 2.4 x 10 4. A splitter plate attached at the trailing edge of the cylinder decreased the frequency of the peak observed in the velocity spectrum compared with the global shedding frequency of the wake for the cylinder alone. For g/D 3, the frequency increased to the global shedding frequency. It has been shown that the detailed pressure distribution around the cylinder can be used to predict the increase or decrease of the frequency. For the first time, a coupling between the local shear layer instability and the wake instability on the frequency characteristics of such a flow has been suggested. Nomenclature Cp = pressure coefficient, 2(P - />«,) / (p U2J D = cylinder diameter or side of the square cylinder Dp = maximum width of the cylinder normal to flow / = frequency, Hz g = gap between the cylinder and the splitter plate L = length of the splitter plate / = length of the cylinders P = pressure at any point P^ = freestream pressure ReD = Reynolds number, U^D/v StD = Strouhal numbers,/D/LL LL = freestream longitudinal mean velocity x = downstream distance from the trailing edge of the cylinder y = transverse distance from the x axis a = angle of incidence v = kinematic viscosity of air p = density of air

Prediction of supersonic inlet unstart caused by freestream disturbances

Abstract: The objective of this article is to report progress toward the development of an Euler analysis procedure for predicting the unstart tolerance of supersonic inlets. As an aid to understanding boundary condition issues, a one-dimensional, linear-analysis procedure was developed and used to analyze inlet unstart behavior. Using these results as a guide, an Euler analysis procedure was extended through the addition of a new bleed boundary condition, a new compressor face boundary condition, and an engine demand model for the simulation of unsteady inlet flows caused by freestream flow disturbances. Five unstart conditions were identified with the Euler analysis of the axisymmetric inlet for both 20- and 90-deg throat bleed configurations. Results show that both increases and decreases in temperature or velocity will unstart the inlet, whereas only pressure decreases will unstart the inlet. It was also found that 90-deg throat bleed improves the unstart tolerance relative to 20-deg throat bleed for freestream pressure decreases, temperature increases, and changes in velocity.

Thrust loss due to supersonic mixing

Abstract: The maximum thrust developed by a device in which two streams mix in a parallel configuration at supersonic velocities is estimated. Total pressure profiles in a two-dimensional, compressible shear layer are calculated by assuming turbulent Prandtl and Lewis numbers of unity. As the convective Mach number Mc rises, the total pressure acquires a defect that becomes large for Mc > 1. For shear layers with equal freestream total pressures, an analytical relation for the defect vs Mc is found. The extent and magnitude of the defect agrees well with experimental data. The loss in total pressure is connected to the loss in thrust of a simplified model of a scramjet. The thrust loss is about 30% for Mc = 2 and 50% for Mc 3. The trends are insensitive to details of the shear-layer velocity profile and to the ratios of freestream quantities. The role of turbulent-energy dissipation in the reduction of total pressure is discussed.

Minimum-Drag Axisymmetric Bodies in the Supersonic/Hypersonic Flow Regimes

Abstract: A study of minimum-drag body shapes was conducted over a Mach range from 3 to 12. Numerical results show that power-law bodies result in low-drag shapes, where the power n = 0.69 (l/d = 3) or n = 0.70 (l/d = 5) shapes have lower drag than theoretical minimum results (« = 0.75 or 0.66, depending on the particular form of the theory). To evaluate the results, a numerical analysis was made, including viscous effects and the effect of a gas model. None of these considerations altered the conclusions. The Hayes minimum-drag body was analyzed and had a higher drag than the optimum power-law body. d i n r T rref x, y, 0 Nomenclature = drag coefficient based on the maximum cross-sectional area = skin-friction coefficient = pressure coefficient = body diameter at the base = marching plane index = unified supersonic-hypersonic similarity parameter, = body length = freestream Mach number = power-law exponent = body radius = temperature at the body surface = freestream temperature z - physical coordinates = circumferential angle

Near-wake structure for a generic configuration of aeroassisted space transfer vehicles

Abstract: Results of a numerical study are presented for hypersonic low-density nitrogen gas flow about a 70-deg blunt cone using the direct simulation Monte Carlo method.The flow conditions simulated are attainable in existing lowdensity hypersonic wind tunnels; encompassing freestream Knudsen numbers of 0.03 to 0.001. Particular emphasis is given to the near-wake flow and its sensitivity to rarefaction and other parametric variations. A stable vortex forms in the near wake at and below a freestream Knudsen number of 0.01, and the size of the vortex increases with decreasing freestream Knudsen number. The base region of the flow remains in thermal nonequilibrium for all cases. There is no formation of a lip separation shock or a distinct wake shock at these rarefied conditions.

Numerical study of the effects of wall catalysis on shock wave/boundary-layer interaction

Abstract: This article presents a numerical study to investigate the effects of nonequilibrium chemistry, and in particular, wall catalysis on the separated flow region generated by an oblique shock wave impinging upon a flat plate boundary layer for a highly dissociated air flowfield. The results focus on the effects of the nonequilibrium chemistry upon the surface heat transfer and the separation zone size. Comparative results are given for chemically reacting (both noncatalytic and fully catalytic walls) and nonreacting flow cases. Furthermore, this comparison is extended over a wide range of freestream pressures (143-123,500 Pa) with a constant Reynolds number, Re = 1793. A direct comparison of all three cases, at low pressures, reveals a minimal change in the peak heat transfer for the noncatalytic wall case as compared to the calorically perfect gas case. In contrast, the fully catalytic wall exerted a tremendous increase in the surface heat transfer. However, as the freestream pressure is increased, significant recombination occurs, so the increase in the peak heat transfer for the noncatalytic wall is more pronounced. Whereas for the fully catalytic wall, at higher pressures, the increase in peak heat transfer is somewhat diminished due to the chemical recombination upstream of the reattachment point.

Parallel fuel injection from the base of an extended strut into supersonic flow

Abstract: Planar Rayleigh/Mie scattering and acetone PLIF flow visualizations along with CFD results are presented for helium injected at sonic velocity into a nominal Mach 2 freestream air flow. The helium is injected parallel to the freestream from an extended strut with three different nozzle-to-freestream air static pressure ratios. Jet spread is insignificant for all three pressure cases. However, large scale, spatially periodic and organized structures are observed primarily in the under-expanded cases. The jet interaction is markedly three dimensional as exhibited by the irregular helium jet contour and the appearance of a conical shock in the highly under-expanded case. The Mach disk, jet spread, barrel shock and recirculation zone shown in the CFD results compare reasonably well to the planar Rayleigh/Mie scattering and the acetone PLIF images. --

Hypersonic nonequilibrium Navier-Stokes solutions over an ablating graphite nosetip

Abstract: The general boundary conditions, including mass and energy balances, of chemically equilibrated or nonequilibrated gas adjacent to ablating surfaces have been derived. A computer procedure based on these conditions was developed and interfaced with the Navier-Stokes solver GASP (General Aerodynamics Simulation Program). A test case with a proposed hypersonic test-vehicle configuration and associated freestream conditions was developed. The solutions of the GASP code with various surface boundary conditions were obtained and compared with those of the ASCC (ABRES Shape Change) code, and the effect of nonequilibrium gas as well as surface chemistry on surface heating and ablation rate were examined. Nomenclature A = area, m2 B' = quantity defined in Eq. (7) CH = #-type heat-transfer coefficient, kg/m2 • s CM = mass-transfer coefficient

Aerodynamic design of Pegasus - Concept to flight with computational fluid dynamics

Abstract: Pegasus™, a three-stage, air-launched, winged space booster, was developed to provide fast and efficient commercial launch services for small satellites The aerodynamic design and analysis of the vehicle were conducted without wind-tunnel and subscale model testing, using only computational aerodynamic and fluid-dynamic methods All levels of codes, ranging in complexity from empirical database methods to three-dimensional NavierStokes codes, were used in the design This article describes the design and analysis requirements, the unique and conservative design philosophy, and the analysis methods considered for the various technical areas of interest and concern Nomenclature AR = aspect ratio CD = drag coefficient CL = lift coefficient Cm = pitching-moment coefficient CAT = normal-force coefficient c = mean aerodynamic chord g = gravitational acceleration h = altitude / = reference length MOO = Mach number p - local pressure Poo = freestream pressure q^ - dynamic pressure S = reference area t = time x = axial coordinate measured from the nose y = lateral coordinate z = vertical coordinate a = angle of attack, deg ft = angle of sideslip, deg 8 = horizontal tail deflection angle, deg 0 = angle of roll, deg i = taper ratio

Numerical investigation of dynamic stall of an oscillating aerofoil

Abstract: The flow structure around an NACA 0012 aerofoil oscillating in pitch around the quarter-chord is numerically investigated by solving the two-dimensional compressible N-S equations using a special matrix-splitting scheme. This scheme is of second-order accuracy in time and space and is computationally more efficient than the conventional flux-splitting scheme. A 'rigid' C-grid with 149 x 51 points is used for the computation of unsteady flow. The freestream Mach number varies from 0.2 to 0.6 and the Reynolds number from 5000 to 20,000. The reduced frequency equals 0.25-0.5. The basic flow structure of dynamic stall is described and the Reynolds number effect on dynamic stall is briefly discussed. The influence of the compressibility on dynamic stall is analysed in detail. Numerical results show that there is a significant influence of the compressibility on the formation and convection of the dynamic stall vortex. There is a certain influence of the Reynolds number on the flow structure. The average convection velocity of the dynamic stall vortex is approximately 0.348 times the freestream velocity.

Laminar and turbulent flow computations of Type 4 shock-shock interference aerothermal loads using unstructured grids

Abstract: This report presents computations for the Type 4 shock-shock interference flow under laminar and turbulent conditions using unstructured grids Mesh adaptation was accomplished by remeshing, refinement, and mesh movement Two two-equation turbulence models were used to analyze turbulent flows The mean flow governing equations and the turbulence governing equations are solved in a coupled manner The solution algorithm and the details pertaining to its implementation on unstructured grids are described Computations were performed at two different freestream Reynolds numbers at a freestream Mach number of 11 Effects of the variation in the impinging shock location are studied The comparison of the results in terms of wall heat flux and wall pressure distributions is presented

Behavior of separation bubble and reattached boundary layer around a circular leading edge

Abstract: The flow around a circular leading edge airfoil is investigated in an incompressible, low turbulence freestream. Hot-wire measurements are performed through the separation bubble, the reattachment and the recovery region till development of the fully turbulent boundary layer. The results of the experiments in the range of Reynolds numbers 1.7×103 to 11.8×103 are analysed and presented in this paper. A separation bubble is present near the leading edge at all Reynolds numbers. At the lowest Reynolds number investigated, the transition is preceded by strong low frequency oscillations. The correlation given by Mayle for prediction of transition of short separation bubbles is successful at the lower Reynolds number cases. The length of the separation bubble reduces considerably with increasing Reynolds number in the range investigated. The turbulence in the reattached flow persists even when the Reynolds number based on momentum thickness of the reattached boundary layer is small. The recovery length of the reattached layer is relatively short and the mean velocity profile follows logarithmic law within a short distance downstream of the reattachment point and the friction coefficient conforms to Prandtl-Schlichting skin-friction formula for a smooth flat plate at zero incidence.Copyright © 1994 by ASME