Showing papers on "Freestream published in 1976"

High frequency thin-airfoil theory for subsonic flow

Abstract: A large parameter solution procedure of Schwartzschild and Landahl is adapted to the task of calculating closed-form approximate solutions for the pressure and lift of a flat-plate, infinite-span airfoil. Two general cases are treated: 1) the two-dimensional subsonic flow problem, in which the large parameter is the upwash frequency, and 2) the three-dimensional incompressible flow problem, in which the large parameter is the spanwise wavenumber of the upwash. For the first case, the four problems of a gust drifting with the freestream, a gust moving at other than the freestream velocity, a plunging motion, and a linear upwash are treated. For the second case, the two problems of a gust drifting with the freestream and a generalized gust moving at other than the freestream velocity are considered. Comparison of the solutions with available numerical results generally shows good agreement when the appropriate parameter is large. The solutions for the gust convecting with the freestream for both the two-dimensional compressible and the three-dimensional incompressible cases were derived previously by Adamczyk using the Wiener-Hopf technique.

Supersonic Shock Wave Turbulent Boundary-Layer Interactions

Abstract: Results are presented of an experimental investigation of shock wave turbulent boundary-layer interactions in supersonic flow. The experiments were conducted at a freestream Mach number of 2.96 and over a boundarylayer Reynolds number range of 10 5 to 10 6. Surface static pressure measurements, oil flow photographs, and interferograms were obtained to define the length of separation and the incipient separation angles for 1) twodimensional compression corner and 2) planar shock wave interactions with a turbulent boundary layer. The tests were conducted in a high unit Reynolds number freestream on a long flat plate with a turbulent boundarylayer thickness in the interaction region of from 0.12 to 0.18 in. Direct comparisons were made between the compression corner and incident shock wave interactions to determine the effects of configuration on turbulent boundary-layer separation. For both configurations the length of the separated region was found to decrease and the incipient separation angle to increase with increasing Reynolds number. For constant Reynolds number, the overall pressure rise for incipient separation was approximately the same for the compression corner interaction and the incident shock wave interaction. Turbulent boundary-layer separation was found to be of the "free interaction" type whereby the separation angle and pressure distribution through separation were independent of Reynolds number, overall pressure rise, and configuration. fo Nomenclature skin friction coefficient at beginning of interaction freestream Mach number pressurepressure freestream pressure Reynolds number based on freestream condition and boundary-layer thickness at beginning of interaction . span °f shock generator stagnation temperature wall temperature axial distance from flat plate leading edge axial location of center of interaction axial distance from flat plate/ramp hinge line to shock generator leading edge axial location of separation point axial offset distance vertical distance between flat plate and shock generator leading edge effective incipient separation corner angle compression ramp angle shock generator angle boundary-layer thickness at beginning of interaction

Diffusion of a tangential drag-reducing polymer injection on a flat plate at high reynolds numbers

Abstract: The diffusion of a thin tangential jet of an aqueous solution of drag-reducing polymer injected into the water- turbulent boundary layer of a flat plate at a freestream Reynolds number, 3.6 x 10 to the seventh power, and the accompanying drag reduction are investigated for a variety of initial concentrations and ratios of injection to freestream velocities. The concentration distribution along the wall is found to be mainly represented by two regions. In the first region the wall concentration is practically constant and equal to the injected one; in the second region the concentration varies approximately as the inverse of the distance from the injection slit. The length of the first region is significantly increased by the polymer solution injection as compared with the pure solvent injection. The drag-reduction effect associated with the polymer injection depends on the trailing-edge concentration achieved as a result of the diffusion process.

Boundary-layer profile measurements in a combustion driven MHD generator

Abstract: The velocity, temperature, and electron number density profiles were measured in the electrode wall boundary layer of a combustion driven MHD generator. Both subsonic and supersonic conditions were run. The experimental results are compared with predictions of a two-dimensional turbulent boundary-layer computation. For the subsonic condition, high levels of freestream turbulence were measured, about 10-12 percent. The measured velocity profile was fatter than that predicted, although the temperature and electron number density profiles were in agreement. This difference is tentatively ascribed to the high freestream turbulence levels. There was no measurable MHD effect for the subsonic case. For the supersonic condition, the measured velocity, temperature, and electron number density profiles fell under the predicted profiles. The discrepancy may be due to three-dimensional recirculation effects. There was a small amount of MHD interaction, the degree of which was in agreement with predictions. Electron number density nonequilibrium was not identified, but the degree of nonequilibrium predicted was small. Under the appropriate supersonic conditions, primarily at freestream temperatures below 2400K, ionization nonequilibrium is predicted to occur.

Experimental Investigation of the Boundary Layer on a Rotating Cylinder

Abstract: This paper documents the experimental analyses done in determining the stability, transition, and growth of boundary layers on a spinning cylinder at angle of attack. It has been shown that spin alters the boundary-layer growth as well as skewing and moving forward the transition line. These effects can have a significant influence on the thickness distribution of the boundary layer. F r R = - Re = u U(y) x,y, z a. rotating cylinder is shown schematically in Fig. 1. The model was installed in a 100 cm x 150 cm subsonic tunnel. A hot-wire probe is remotely positioned in the boundary layer by a traversing mechanism with three degrees of freedom: jc, the longitudinal distance along the cylinder; y, the radial distance from the cylinder; and z, the distance along the circumference of the cylinder. The traversing mechanism can also be ad- justed so that the probe may be positioned at any angle about the axis of the cylinder. Tests were conducted at a freestream velocity of 34 fps. As shown in Fig. 1, there are several unique features in- corporated in the design of the model. First, an annular suc- tion slot is used at the nose of the model to remove all of the boundary layer which develops along the nose. In this way, the effect of the particular design of the nose upon the flow on the cylindrical test section is eliminated. The second feature concerns the injection of a velocity per- turbation into the boundary layer through the use of a tuned acoustic driver which sets up standing waves within the in- terior of the cylinder. These waves interact with the flow through the annular suction slot to produce the desired signal. Hot-wire measurements show that this signal is confined almost entirely to the boundary layer. The growth and decay of this signal is tracked along the body and the coordinate of the maximum and minimum points of the perturbation serve to determine the location of the neutral stability curve. A third aspect of the model design is the use of an annular disk suction slot located to the rear of the rotating test section. By adjusting the suction flow rate through this slot, the longitudinal pressure gradient along the cylinder is removed without resorting to cumbersome tunnel wall adjustments. Since reported measurements of the neutral stability curve for a flat plate show considerable scatter, a two-channel hot- wire measuring system was used to improve the consistency of the results. A fixed hot-wire is positioned in the boundary layer near the middle of the cylinder and is used to monitor the signal. Any change in the effective Reynolds number or in the injected signal strength immediately shows up as a change

Experimental Poststall Rotary Aerodynamic Coefficients for Airplane-like Configurations

Abstract: Military and civilian airplane losses caused by out-of-contro l spin motions are significant. Knowledge of rotary coefficients is necessary to understand the cause of spin entry and to devise proper recovery techniques. An exploratory wind-tunnel investigation has been conducted at Ames Research Center on simple airplane-like configurations on a rotary sting apparatus at rotation rates up to 10 rps. Rotary coefficients have been measured at unit Reynolds numbers from 2xl06nT1 to 24.6xl06nT1 and at angles of attack from 45° to 90°. Results show that the aerodynamic characteristics at steady spin rates are highly dependent on both spin rate and Reynolds number. Nomenclature A = body reference area, ird2/4 b = wing span, 0.457 m (1.5 ft) CD = body drag force per unit length/qd c'N = body normal force per unit length/qd CR = moment about spin axis/qSb, positive clockwise (viewed from rear) c'y = body side force per unit length/qd Cy = body side force/qA d = diameter of center body, 0.0762 m (0.25 ft) £ = length of nose section M = freestream Mach number Nj = tangent ogive nose, see Fig. 4 q = freestream dynamic pressure Rd = Reynolds number based on d S -wing reference area, 0.0491 m2 (0.5234 ft2) TI = aft body, circular cylinder tail section, see Fig. 4 T2 = T] plus tail surfaces, see Fig. 4 U = freestream velocity W = wing, see Fig. 4 xn = distances, see Fig. 9 a.f = local angle of attack, see Fig. 9 Of = angle between the freestream velocity vector and the body x axis \l/ = angle of roll about the body x axis co = angular velocity of the rotary sting Q = reduced spin rate, ub/2U

Numerical Solutions for Transonic Flows Past Wing-Body Combinations

Abstract: A numerical method has been developed for calculating the inviscid flow past lifting wing-body combinations by solving a form of the transonic small-perturbation equation for the velocity potential. An outline is given of the formulation of the problem and the procedure for numerical solution. Numerical results are compared with data from wind-tunnel tests on a wing-body configuration, and the variation of body interference with angle of incidence and freestream Mach number is illustrated.

Subsonic loads on wings having sharp leading edges and tips

Abstract: A vortex-lattice method for predicting the aerodynamics of wings having separation at the sharp edges in incompressible flows is extended to compressible subsonic flows using a modified Prandtl-Glauert transformation. Numerical results showing the effect of freestream Mach number on the aerodynamic coefficients are compared with available experimental data for several planforms. It is shown that the proposed method is suitable for predicting the aerodynamic loads on low-aspect wings at moderate angles of attack for high subsonic freestream Mach number. The method is limited to angles of attack up to 12 deg for high subsonic freestream Mach number and to angles of attack up to 20 deg for Mach number not exceeding 0.5.

Spatially decaying array of vortices

Abstract: An analysis of two‐dimensional vorticity disturbances introduced at a grid and propagating downstream reveals that viscosity causes the x wavenumber to diminish and the phase speed to increase from their inviscid values, while also causing the amplitude to decay exponentially downstream. This decaying disturbance field introduces an adverse mean pressure gradient of strength O (1/RΛ), where RΛ is the Reynolds number based on freestream velocity and vortex diameter. Viscous effects introduce a discrepancy between the two‐point and two‐time velocity correlations when Taylor’s hypothesis r=U∞t is used to relate the space and time separations. This discrepancy arises because the vortices propagate at a phase speed c greater than the freestream. As RΛ→∞, c→U∞. If Taylor’s hypothesis is modified to r=ct, then the two‐point and two‐time correlations agree.

Re-examination of Compliant Wall Experiments in Air with Water Substrates

Abstract: A possible alternative explanation is proposed for compliant wall drag reductions measured in previous investigations. Standing waves were observed to form on the surfaces of compliant wall models in air with water substrates as the freestream velocity was increased from 15 to 30 m/s. These waves resembled sine waves with half of the wave protruding over the upstream portion of the model and the other half being recessed over the downstream end of the model. These data coupled with results of recent drag reduction experiments suggest that standing waves could have caused a shift in the model center of gravity creating a bending moment that was interpreted as a reduction in the skin friction drag.

Modified thin shock layer method for supersonic flow past an oscillating cone

Abstract: The problem of supersonic flow past a circular cone oscillating about its vertex is considered. The amplitude and the frequency parameter of the oscillation are assumed to be small, and a perturbation solution in the amplitude and frequency is sought. Furthermore, thin shock layer expansion is used to derive the flowfield solution in the form of a series. The first three terms in the series are obtained, showing that the series solution tends to converge when the shock layer is very thin and otherwise it tends to diverge. The technique of parameter straining then is applied which greatly improves the accuracy and extends the range of applicability of the thin shock layer solution. In particular, simple explicit formulas for the stability derivatives are valid for moderate as well as high freestream Mach numbers and for thick as well as slender cones. Variations of the stability derivatives with the freestream Mach number, specific heat ratio, and the cone semiangle are investigated and comparisons with existing theories are included. The relation of limiting gasdynamic theory with unsteady Newtonian flow theory also is discussed. k £ MOO m n p r t t utv,w

Similarity solution of the boundary-layer equations for laser heated flows

Abstract: The laminar boundary-layer equations with local similarity approximation are solved in order to estimate the heat transfer rate to the wall in a laser-heated rocket thruster using pure hydrogen, where the average temperature of the hot plasma core is about 14,000 K and the core Reynolds number based on nozzle throat diameter is about 2000. Under these conditions, the density-viscosity product at the wall can be 10 times the freestream value, and the hydrogen is completely dissociated. Hence the boundary layer equations with variable transport properties are solved by a quasi-linearization technique, and the equilibrium properties of hydrogen are used in the calculations. Velocity profiles are obtained, and the wall shear and wall stagnation enthalpy gradient are plotted against the pressure gradient parameter.

Analytical Model of Axisymmetric Af terbody Flow Separation

Abstract: A practical engineering approach to the prediction of pressure distribution and boundary layer separation point location on afterbodies in subsonic flow is developed. Experimental data are reviewed and the inadequacy of currently available separation prediction methods is demonstrated. A control volume technique is developed as an alternative and is shown to have merit. The separation bubble is then modeled and another control volume technique is developed to predict its outer boundary. Finally, these two developments are combined with a conventional inviscid flowfield calculation and a boundary-layer analysis to produce an iterative procedure to predict pressure distribution and separation point location on an afterbody given only body shape and freestream flow condition.

Rough wall turbulent heat transfer with variable velocity, wall temperature, and blowing

Abstract: An integral prediction method is presented which accurately describes Stanton number behavior for a fully rough turbulent boundary layer flowing over a uniformly rough surface. The kernel function which represents the response of such a system to an unheated starting length is shown to also describe the response to variations in freestream velocity, surface temperature and blowing. Predictions are compared with experimental data for cases of variable wall temperature, favorable pressure gradients and variable blowing. Agreement is excellent in all cases. (auth)

Engineering Flowfields and Heating Rates for Highly Swept Wing Leading Edges

Abstract: An engineering flow model is developed (and verified experimentally) which describes the flowfield that arises when a supersonic stream encounters a wedge/cylinder configuration whose angles are such that the flow includes only weak shock waves. A numerical code using the perfect gas relations is used to describe the flow in the plane of symmetry inboard of the shock interaction region. Theoretical surface-pressure and heat-transfer distributions are computed for freestream velocities ranging from 1167 to 7610 m/sec. Nondimensionalization of the heat-transfer rates in terms of local flow parameters produced a correlation of Stanton number as a function of the local Reynolds number, which is independent of the freestream flow conditions and of the surface temperature.

Correlation of freestream pressure disturbances in supersonic wind tunnels

Abstract: An attempt is made to correlate data for freestream disturbance levels in terms of parameters that can be easily calculated for most wind tunnels. A correlation expression is derived assuming that the fluctuating pressure at the wall of the supersonic wind tunnel are related to fluctuating pressures in the freestream. The acoustic origins, flow properties, and geometric factors are taken into account. The suggested correlations of wind tunnel disturbances provide good predictions of rms pressure disturbances in the freestream of wind tunnels or on the surface of cones at Mach numbers greater than 2.5.

Steady flow separation along a straight streamwise corner

Abstract: This note considers the steady incompressible flow of a laminar or turbulent boundary layer along a straight streamwise corner, formed by two flat plates intersecting each other at an arbitrary included angle. It is assumed that the undisturbed freestream is parallel or roughly parallel to the cornerline and that the intersecting corner-walls are impermeable. Attention is focussed on the necessary and sufficient conditions for steady flow separation from the sharp cornerline.

The effect of varying freestream velocity on dynamic stall characteristics

Abstract: A low speed wind tunnel equipped with an axial gust generator to simulate the aerodynamic environment of a helicopter rotor was used to study the dynamic stall of a pitching blade. The objective of this investigation was to find out to what extent harmonic velocity perturbations in the freestream affect dynamic stall. The study involved making measurements of the aerodynamic moment on a two-dimensional, pitching blade model in both constant and pulsating airstreams. Using an operational analog computer to perform on-line data reduction, plots of moment versus angle of attack and work done by the moment were obtained. The data taken in the varying freestream were then compared to constant freestream data, and to the results of two analytical methods. These comparisons showed that the velocity perturbations had a significant effect on the pitching moment which could not be consistently predicted by the analytical methods, but had no drastic effect on the blade stability.

Stagnation Region Gas Film Cooling Spanwise Angled Coolant Injection.

Abstract: : This experimental investigation involved the study of gas film cooling from a single row of spanwise angled holes using the stagnation region of a cylinder in cross flow to model the leading edge of a turbine vane. The objective was to obtain data for the local convective heat transfer rates to a highly cooled, curved surface exposed to a turbulent hot mainstream flow and a secondary, film coolant flow. Since the leading edge of the first stage, inlet turbine vane experiences some of the most severe thermal loads found in the turbine engine, effective film cooling is most important in this area. Film cooling of the leading edge area was modeled by making heat transfer measurements on the front stagnation region of a cylinder in cross flow. Experiments were conducted in a rectangular duct using a film cooled cylindrical test surface normal to a two-dimensional freestream flow. A gas turbine combustor provided heated air flow to simulate a Reynolds number typical of a high pressure, high temperature turbine vane. Internal convection cooling of the cylinder allowed a gas-to-wall temperature ratio of 2.1 to be achieved while using a moderate freestream gas temperature (1000R; 555K. The film coolant was chilled to obtain a coolant-to-freestream density ratio of 2.2, representative of the gas turbine environment. The cylindrical test surface was instrumented with miniature heat flux gages, and wall thermocouples to determine the influence of the film coolant blowing ratio and the injection hole geometry on the film cooling performance.

Investigation of Slot Cooling at High Subsonic Speeds

Abstract: The paper discusses the results of an experimental and numerical investigation of tangential slot injection film cooling with zero pressure gradients in subsonic boundary layers at freestream Mach numbers of 0.4, 0.6, and 0.8. The results are compared with the predictions obtained from a finite-difference boundary-layer program developed by NASA for slot injection into turbulent supersonic boundary layers. The two sets of results are found to compare favorably. The numerical results point to the existence of a unique relation between isothermal effectiveness and adiabatic effectiveness, thereby confirming the existence of similarity conditions for temperature and velocity profiles.

Computational Aspects of Nonsimilar Solutions of the Rarefied Leading Edge on a Cone Wesley

Abstract: viscous layer is modeled after that of Yasuhara with modified external pressure field and boundary conditions. The original freestream boundary conditions of Yasuhara and the classical Rankine-Hugoniot conditions downstream of the conical shock are replaced by matching the mass flux, the tangential shear, velocity components, and state variables at the shock wave-viscous layer interface. Attention is given to the details of the initial data profiles. We also note the finite-difference form of the model equations which are required for the generation of a solution which is physically acceptable and mathematically selfconsistent. Heat transfer rate and viscous shear at the surface of the cone are computed. First-order corrections for shock curvature, velocity slip, and temperature jump at the cone surface are not considered in this paper. The proposed method does provide a straightforwar d means for including these effects.1 The following relations can be obtained by integrating the zeroth order conservation equations for the shock structure and applying the freestream boundary conditions: p°v°=-smds

Transition Simulation for Incompressible Boundary Layers and Heated Hydrodynamic Bodies.

Abstract: : Incompressible boundary-layer transition has been analyzed using a second-order closure turbulence model. With no transition-specific modifications, the turbulence model predicts salient features of incompressible, zero-pressure-gradient boundary-layer transition including sensitivity to freestream turbulence and surface roughness, transition width, and transitional velocity profiles. With transition modifications based on linear stability theory, the model accurately predicts transition sensitivity to surface heat transfer, pressure gradient, and suction. With no further modifications, transition predictions have been made for several hydrodynamic bodies, including effects of surface heating. (Author)

A theoretical investigation of over-wing-blowing aerodynamics

Abstract: A theoretical method is established for determining the aerodynamic characteristics of over-wing-blowing configurations. The method accounts for both jet entrainment and jet interaction effects because of the differences in freestream and jet dynamic pressures and Mach numbers. The predicted lift increments agree well with available data. It is shown that the lift is underpredicted with entrainment effect alone when the jet is close to the wing surface.

An investigation of a jet in a crosswind

Abstract: Wind tunnel experiments were conducted to determine the complete longitudinal interference characteristics of a turbulent jet exhausting from a flat plate into a turbulent subsonic freestream. The apparatus was designed so that the trends from systematic variations in one parameter, while the others remain fixed, could be established. The variable parameters were jet inclination, the plate Incidence and the ratio of the jet exit velocity to the freestream velocity (the velocity ratio.) The angles of jet inclination, measured from the normal to the plate surface, varied from 0 to 60 degrees downstream in increments of 15 degrees. The angle of incidence of the plate to the freestream direction varied from 0 to 8 degrees in increments of 2 degrees. The values of the velocity ratio ranged from 4 to 12, values that are pertinent to the range of interest for V/STOL aircraft in transitional flight, The surface pressure distribution about the jet and the jet trajectory, defined as the locus of the maximum total pressure, were measured for each configuration. In addition, the surface pressure distribution was integrated numerically to provide a surface force distribution about the jet, a suction force coefficient, a pitching moment coefficient and the centre of pressure. The results are summarised by presenting the variation of the suction force coefficient, centre of pressure, pitching moment coefficient and jet trajectory with the velocity ratio for a given jet inclination and plate incidence. These curves can be crossplotted to provide the variation of these quantities with the jet inclination or the plate incidence as the independent variable In addition, selected isobar plots are presented. The extent of the low pressure field in the lateral and forward regions was reduced as the jet inclination increased. The contribution from these regions to the lift loss and the magnitude of the lift loss decreased. The centre of pressure moved downstream accordingly. The jet penetrated the freestream less and was deflected less as the jet inclination increased. These observations were attributed to a change in the entrainment rate of the jet. The jet entrainment rate decreased as the jet inclination increased. The changes in the surface pressure distribution resulting from a change in incidence of the plate were detailed rather then gross. The variation of the lift loss with incidence exhibited a maximum between 40 and 60 incidence. The change in jet penetration and. deflection was small. The centre of pressure appeared to be independent of incidence. A change in incidence appeared to cause an effective change in the inclination of the jet. The entrainment rate of the jet was only moderately affected by a change in incidence. The low pressures spread to the lateral and forward region as the velocity ratio-increased. The contribution of these regions to the lift loss increased while that from, the wake region decreased. The magnitude of the lift loss increased as the velocity ratio increased. The centre of pressure moved upstream accordingly. Both the lift loss and the centre of pressure showed a weak dependence on the velocity ratio for large values of the velocity ratio. The jet penetrated the freestream more and suffered a less severe initial deflection as the velocity ratio was increase. These observations were attributed to an increase in the entrainment rate of the jet as the velocity ratio increased, A two-dimensional potential flow model was proposed to predict the surface pressure distribution, the surface force distribution and the suction force coefficient on the flat plate for the normal jet at aero incidence. The model successfully predicted the surface pressures and surface forces close to the jet in the lateral and forward regions, for velocity ratios less than 10, The agreement between the predicted and experimental value of the suction force coefficient was particularly good for velocity ratios less than 10, The model was unable to allow for the increasing three-dimensional effects at high ratios. The trends in the variation of the model parameters with the velocity ratio agreed well with the experimentally observed trends of the physical characteristics which they represented.

Strong interaction effects on upstream wall pressure in a separated flow with a forward-facing wall jet

Abstract: Results are presented for an experimental study of the effect of separation on the upstream wall pressure at large Reynolds number in the case of separation of an incompressible turbulent boundary layer on a flat surface. Interpretation of observed upstream wall pressure is facilitated by introducing the concept of reference and interaction flows. The interaction flow is found to be in the forward direction upstream of the separation point, and its magnitude ranges from 10 to 30% of the freestream velocity. A correlation which could be interpreted as a coordinate expansion for large distances is proposed. A discussion is presented of two interpretations of the observed wall pressure distribution for the separation of the boundary layer under favorable pressure gradient and for the separation of the wall jet under adverse pressure gradient.