Showing papers on "Freestream published in 1982"

Aerodynamic Loads and Performance of the Darrieus Rotor

Abstract: The double-multiple-streamtube model was used for determining the aerodynamic blade loads and roto performance on the Darrieus vertical-axis wind turbine with curved blades. This analytical model i capable of predicting the difference in the induced velocities at the upstream and downstream passes. The upwind and downwind interference factors are calculated by a double iteration, one for each half of the rotor and vertical variations in the freestream velocity are accounted for. Under local aerodynamic conditions a closed-form analytical solution is obtained for two functions, one for the upwind and one for the downwind hal of the turbine. Thus the local aerodynamic loads on the blade and the rotor performance are calculated more accurately than by other streamtube methods. Comparison of the analytical results obtained with the doublemultiple-streamtube model and the available field test data for the Sandia 17-m machine shows good agreement. This method can therefore be used for generating a suitable aerodynamic-load model for structural-design analysis of the Darrieus rotor.

An experimental investigation of the unsteady behavior of blunt fin-induced shock wave turbulent boundary layer interactions

Abstract: : An experimental study has been made of blunt fin-induced shock wave turbulent boundary layer interactions. This type of interaction is known to be highly unsteady. The objective of this experiment was to determine the characteristics of the fluctuating surface pressure distribution and the parameters controlling it. Tests have been made using fins of different diameter, D, with incoming turbulent boundary layers varying in thickness, Delta, in the ratio of about 5:1. Measurements have been made on the fin centerline and up to four diameters outboard of it. All tests were made at a Mach number of 2.95 and a unit Reynolds number of 6.3 billion/m, and under approximately adiabatic wall conditions. The measurements show that very high intensity r.m.s. pressure levels occur--up to almost two orders of magnitude above that of the incoming boundary layer. The highest intensities occur on centerline ahead of the fin. Here, the r.m.s. pressure distribution is characterized by three distinct peaks which decrease at different rates with distance outboard. Even four diameters off centerline, the maximum r.m.s. value in the distribution is still an order of magnitude larger than that of the incoming boundary layer. Outboard of the centerline, the r.m.s. pressure level downstream of the freestream shock wave steadily decreases. Within a distance of six to eight diameters it is close to the undisturbed value. With different diameter fins and different boundary layers, the qualitative characteristics are the same. The quantitative results depend on the ratio D/Delta. (Author)

Comparison of sharp and blunt fin-induced shock wave/turbulent boundary-layer interaction

Abstract: In this paper, results are presented from an experimental study of fin-induced shock wave/turbulent boundary-layer interaction. Semi-infinite fin models, with sharp and hemicylindrically blunted leading edges, were tested at Mach 3 in two high Reynolds number, adiabatic wall, turbulent boundary layers. Detailed streamwise surface pressure distributions were measured at several spanwise stations for angles of attack from 0 to 12 deg. The objective was to examine the spanwise development of the interaction region and determine its dependence on leading edge geometry and the incoming-flow parameters. The results show that for blunt fins there is a region of the interaction in the vicinity of the centerline where the scale and characteristics of the disturbed flowfield are controlled primarily by the leading edge diameter. Outboard of this inner, or "leading-edge dominated," region the interaction properties and spanwise development are essentially the same as if the leading edge were sharp. Thus there is an outer region of the blunt fin-induced flowfield in which the properties are effectively "independent of leading edge blunting." Nomenclature D = blunt fin leading-edge diameter h - fin height Lu = upstream influence measured relative to the freestream shock wave M = Mach number P = static pressure P2 = wall pressure downstream of interaction Re = Reynolds number X = coordinate in the plane of the test surface and aligned with the tunnel axis, with X= 0 at the fin leading edge Xs = coordinate in the X direction measured relative to the freestream shock wave Y = coordinate normal to the X axis in the plane of the test surface, with Y= 0 at the fin leading edge

Equilibrium Parameters for Two-Dimensional Turbulent Wakes

Abstract: The parameters characterizing a plane turbulent wake in its equilibrium state of development are determined on the basis of experiment and analysis. Measurements are made in an open-circuit suction-type wind tunnel with a contraction ratio of about 10 and a test-section of about 30 cm sq and 4.27 m long. Less than 1.5% variation in wind speed along the test section is attained by applying suitable divergence for the boundary layer growth. A two-dimensional wake is created behind a twin plate configuration at a freestream velocity of 21.3 m/s. The freestream turbulence level at this speed is about 0.15%.

Shuttle Boundary-Layer Transition Due to Distributed Roughness and Surface Cooling

Abstract: The transition criteria for the windward surface of the Shuttle Orbiter, which is composed of a large number of thermal protection tiles, must include the effect of distributed roughness arising from joints and possible tile misalignments. Theoretical flowfield parameters and heat-transfer distributions were used to analyze boundarylayer transition data. Data were obtained for Mach numbers from 8 to 12 over a range of Reynolds number based on model length from 1.8 to 17.6 x 106 with surface temperatures from 0.114 to 0.4357,. The transition correlations were approximately the same both for the misaligned-tile models and for the grit-roughened model. The incipient value of Rektr was 30, the critical value was 110, and the effective value was 180. Nomenclature h = local heat-transfer coefficient ^/,ref = heat -transfer coefficient for the stagnation point of the reference sphere k = height of the misaligned tiles L = axial model length, 0.5734m (1.881 ft) M = Mach number Rek = Reynolds number based on conditions at the top of the misaligned tile [Eq. (1)] Rey = Reynolds number based on the flow conditions at a distance y from the wall Ree = Reynolds number based on local flow properties and the momentum thickness ReWfL = Reynolds number based on freestream flow properties and the model length T = temperature U = streamwise component of the velocity x = axial coordinate a = angle of attack <5* = displacement thickness 6 = momentum thickness jji = viscosity £ = relative transition location defined in Eq. (2) p —density Subscripts

Measurement and Analysis of Total-Pressure Unsteadiness Data From an Axial-Flow Compressor Stage

Abstract: A fast-response, total-pressure probe was used with a periodically sampling and averaging data acquisition system to study the unsteady total-pressure field in an axial-flow turbomachine. Periodically unsteady total-pressure data were used to demonstrate some of the ways in which turbomachine blade wake transport and interaction influences the energy transfer involved. Observed trends of periodic variations in local total-pressure values could be explained in terms of the details of energy transfer associated with the different kinds of fluid particles (freestream, wake segment, blade surface boundary layer, mixed) moving through a blade row.

Computation of the steady viscous flow over a tri-element 'augmentor wing' airfoil

Abstract: The augmentor wing consists of a main airfoil with a slotted trailing edge for blowing, and two smaller aft airfoils which shroud the jet. This configuration has been modeled for numerical simulation by a novel discretization procedure which generates four separate grids: three surface-oriented airfoil grids and one outer free-stream grid. Grid lines and slopes are continuous across boundaries, so grid overlap at common boundaries provides boundary information without interpolation. A two-dimensional unsteady thin-layer Navier-Stokes code is used to calculate the flow for the no-blowing case at freestream Mach number = 0.7, Re = 12,600.000, and angles-of-incidence = 1.05 deg. Qualitative agreement with experimental data indicates the utility of this procedure in the analysis of multi-element configurations.

Turbulent boundary layer heat transfer experiments: Convex curvature effects including introduction and recovery

Abstract: Measurements were made of the heat transfer rate through turbulent and transitional boundary layers on an isothermal, convexly curved wall and downstream flat plate. The effect of convex curvature on the fully turbulent boundary layer was a reduction of the local Stanton numbers 20% to 50% below those predicted for a flat wall under the same circumstances. The recovery of the heat transfer rates on the downstream flat wall was extremely slow. After 60 cm of recovery length, the Stanton number was still typically 15% to 20% below the flat wall predicted value. Various effects important in the modeling of curved flows were studied separately. These are: the effect of initial boundary layer thickness, the effect of freestream velocity, the effect of freestream acceleration, the effect of unheated starting length, and the effect of the maturity of the boundary layer. An existing curvature prediction model was tested against this broad heat transfer data base to determine where it could appropriately be used for heat transfer predictions.

Flight, Wind Tunnel, and Numerical Experiments with a Slender Cone at Incidence

Abstract: The three-dimensional leeward separation about a 5-deg semi-angle cone of 11 deg angle of incidence was Investigated in night, in the wind tunnel, and by numerical computations. The test conditions were Mach numbers of 0.6, 1.5, and 1.8 at Reynolds numbers between 7 and 10 million based on freestream conditions and a 76.2-cm (30-in.) length of surface. The surface pressure conditions measured included those of fluctuating and mean static, as well as recovery pressures generated by obstacle blocks to provide skin friction and separation-line locations. The mean static pressures from flight and wind tunnel were in reasonably good agreement. The computed results gave the same distributions, but were slightly more positive in magnitude. The experimentally measured primary and secondary separation line locations compared closely with computed results. There were substantial differences In level between the surface root-mean-square pressure fluctuations obtained in night and in the wind tunnel, due, It Is thought, to a relatively high acoustic disturbance level in the tunnel compared with the quiescent atmospheric conditions in night.

Turbulent boundary layer over solid and porous surfaces with small roughness

Abstract: The wind tunnel models and instrumentation used as well as data reduction and error analysis techniques employed are described for an experimental study conducted to measure directly skin friction and obtain profiles of mean velocity, axial and normal turbulence intensity, and Reynolds stress in the untripped boundary on a large diameter axisymmetric body. Results are given for such a body with a (1) smooth, solid surface; (2) a sandpaper roughened, solid surface; (3) a sintered metal, porous surface; (4) a ""smooth'' performated titanium surface; (5) a rough, solid surface made of fine diffusion bonded screening; and (6) a rough, porous surface made of the same screening. The roughness values were in low range (k+ 5 to 7) just above what is normally considered ""hydraulically smooth''. Measurements were taken at several axial locations and tow or normal stream freestream velocities, 45.1 m/sec and 53.5 m/sec.

Analysis of Combustion in Recirculating Flow for Rocket Exhausts in Supersonic Streams

Abstract: A new computer program has been written to calculate the properties of the recirculating base flow region of a rocket exhaust plume. The nozzle stream and the freestream, either of which may be supersonic, mix turbulently to form an axisymmetric compressible free boundary layer. Solutions of the elliptic differential equations, which incorporate a two-equation model of turbulence closure, are obtained using an iterative finite-difference technique coupled to a point-by-point solution of the chemical kinetics. Predictions are presented for the velocity, pressure, temperature, and species concentration distributions in the base recirculation region of a rocket exhaust plume under supersonic flight conditions.

Measurement of small normal velocity components in subsonic flows byuse of a static pipe

Abstract: A concept for measuring small normal velocity components in subsonic flows by use of a static pipe has been developed and tested. A theory is presented which allows one to relate measurements of the difference of static pressures across a static pipe in a slightly disturbed flow to the axial derivative of the velocity component normal to the pipe. This derivative may be integrated, after fixing the constant of integration, to give the normal velocity component distribution along the pipe. The theory is restricted to small normal velocity components relative to the main freestream flow which is aligned with the pipe axis. The theoretical development presented here is primarily concerned with two-dimensional disturbance fields, but results are given for three-dimensional disturbance fields also. An experimental application of the concept to measure the disturbance fields arising in a two-dimensional wind tunnel is described. These experiments show how the present static-pipe concept can be used in a practical experimental situation.

Aerodynamic behavior of a slender slot in a wind tunnel wall

Abstract: A theoretical model for the flow through a single slot of finite length in a wall separating a uniform freestream and a quiescent fluid at different static pressures is constructed. This problem is relevant to understanding the aerodynamic behavior of slots which are used in the test sections of some ventilated wall transonic wind tunnels. The theoretical relationship which is obtained between the pressure differential across the slot and the flow through the slot shows both the linear and quadratic regimes observed in experiments. The linear behavior arises from the acceleration of the cross flow into the slot downstream of the leading edge and from the interaction of streamwise stations along the slot, as well as from the effect of slot taper. Analytical solutions are obtained for two slot planform shapes, and some other cases are solved numerically. The quantitative agreement with experimental data is very encouraging.

Self-excited wave oscillations in a water table

Abstract: An experimental investigation has been performed on self-excited wave oscillations on cavity, spike-tipped, and inlet models in a water table. Buzzing was generated by positioning the models at a small angle of attack with respect to the freestream flow. The hydraulic analogy was used to compare the results obtained in water to results obtained in a gas. High-speed and real-time photography were used in the experiment. The frequencies of oscillations in water and air were consistent with the hydraulic analogy. Numerical solutions of the phenomenon were also obtained. and from considerable savings in energy arid funds. It is in fact far more inexpensive and energy saving to operate a water table than a supersonic wind tunnel. In addition, some phenomena in water can be observed directly without sophisticated equipment, while in a wind tunnel it is necessary to mafe use of schlieren high-speed photography, holography, interferometry, etc. An experimental in- vestigation was therefore performed on self-excited wave oscillations occurring on two-dimension al bodies in water. The objective was to study the "buzzing" phenomenon in water and to compare the results, whenever possible, with previous experimental results obtained in air. Necessary conditions for buzzing to occur are the presence of a shear layer with an inflection point in the velocity profile (separation), a reflecting surface (cavity wall, body shoulder, inlet cowl, etc.), and the appropriate body length to permit "in-phase" reflection of pressure waves. If these conditions are met, resonance occurs in the frequency range of the oscillations and amplification of the disturbance takes place until a "limit cycle" is reached. Photographic data were obtained, i.e., high-speed and real-time movies, to visually illustrate the phenomenon. Furthermore, a stability analysis was developed and numerical solutions were obtained to determine the conditions for instabilities and disturbances like buzzing to occur in water. Cavity, spike-tipped, and inlet models were used in the water table.

A Comparison of the Second-Order Triple-Deck Theory with Interacting Boundary Layers

Abstract: We consider viscous flows over a flat plate with a small hump situated downstream of the leading edge. The characteristic Reynolds number Re based on the freestream conditions (ρ*∞, U*∞, μ*) and the distance L* from the leading edge of the plate to the position of the hump is assumed to be large. The response of the boundary-layer flow to such a protuberance has been reviewed by Sedney [2.118]. Understanding this response is important in studying the effect of a roughness element on the transition of laminar flows to turbulent flows. Furthermore, in high-speed applications, the presence of such protrusions may contribute to the total drag of the vehicle and it may cause local high heating rates. The study is also relevant to atmospheric boundary-layer flows over hills and mountains [2.65, 2.119]; however, the two-dimensional and laminar-flow assumptions are oversimplifications for these applications.

Direct Simulation Monte Carlo Technique For Modeling Of The Environment In The Vicinity Of The Space Shuttle Orbiter

Abstract: A program for analyzing the flowfield parameters in the neighborhood of the Space Shuttle Orbiter has been developed. The program uses the direct simulation Monte Carlo method, which is a completely probabilistic Monte Carlo technique capable of analyzing 3-dimensional steady or unsteady flow with prescribed internal and external boundary conditions. The freestream flux densities incident on the external flowfield boundaries are calculated from the drifting Maxwellian gas properties of the freestream. The flux entering the flowfield from the internal boundary is calculated from the outgassing flux density distribution over the Shuttle external surface and the prescribed discrete source fluxes. This technique produces a numerical flowfield solution which is the probabilistic equivalent of a complete solution of the time-dependent, 3-dimensional Boltzmann equation. Flowfield results are presented for the following configurations: (1) Shuttle angles of attack of 0 and 90 deg, (2) freestream density 10 to the 9th - 10 to the 11th per cu cm, (3) Shuttle outgassing flux density, and (4) operation of the aft downfiring vernier Reaction Control System (RCS) engine. Results are presented for column density distribution of outgassed and engine species and for the flux density of outgassed and engine species incident on the Shuttle bay and the windshield.

Hot-wire investigation of an unseparated shock-wave/turbulent boundary layer interaction

Abstract: Detailed hot-wire measurements of the longitudinal component of the mass-flow fluctuations have been made for an adiabatic fully attached shock-wave/turbulent boundary-layer interaction. The shock wave is induced by an 8-deg compression corner at a Mach number of 2.85 and a unit Reynolds number of 6.3 X10 m ~ l . The data indicate that the absolute mass-flow turbulence intensity increases abruptly through the shock wave, and continues to increase with further distance downstream. When nondimensionalized by the local freestream massflow rate the maximum turbulence intensity actually exceeds the upstream equilibrium value at the last measurement station. Measurements of the probability density of the fluctuations reveal only minor difference between the boundary-layer behavior upstream and downstream of the shock. A significant dip in the kurtosis was observed near the point of maximum intensity, which also coincided with the onset of intermittency.

Near-wall similarity in three-dimensional turbulent boundary layers. I - Model review. II - Pressure-driven flow results

Abstract: Eleven proposed models for near-wall similarity for three-dimensional turbulent boundary layer flows are reviewed. Six of these models are comparatively simple scalar models and five are more complex and/or two-component vector models. Ten of the models can be tested as to their validity or predictive capability with the aid of measured mean velocity field, wall pressure field, and direct wall shear stress field (magnitude and direction) data. One of the models cannot be tested owing to its dependence on two parameters that are at present extremely difficult (if not impossible) to measure. Ten three-dimensional near-wall similarity models are then evaluated with direct wall shear, velocity field, and pressure gradient data from a three-dimensional pressure-driven boundary layer flow. In a primary focus of the interval where y+ is between 50 and 300, graphical results suggest that six simpler models and the freestream component of one complex model are adequate for profiles with monotone increasing skew up to about 15 deg.

Study of boundary-layer transition using transonic cone Preston tube data

Abstract: Laminar layer Preston tube data on a sharp nose, ten degree cone obtained in the Ames 11 ft TWT and in flight tests are analyzed. During analyses of the laminar-boundary layer data, errors were discovered in both the wind tunnel and the flight data. A correction procedure for errors in the flight data is recommended which forces the flight data to exhibit some of the orderly characteristics of the wind tunnel data. From corrected wind tunnel data, a correlation is developed between Preston tube pressures and the corresponding values of theoretical laminar skin friction. Because of the uncertainty in correcting the flight data, a correlation for the unmodified data is developed, and, in addition, three other correlations are developed based on different correction procedures. Each of these correlations are used in conjunction with the wind tunnel correlation to define effective freestream unit Reynolds numbers for the 11 ft TWT over a Mach number range of 0.30 to 0.95. The maximum effective Reynolds numbers are approximately 6.5% higher than the normal values. These maximum values occur between freestream Mach numbers of 0.60 and 0.80. Smaller values are found outside this Mach number range. These results indicate wind tunnel noise affects the average laminar skin friction much less than it affects boundary layer transition. Data on the onset, extent, and end of boundary layer transition are summarized. Application of a procedure for studying the relative effects of varying nose radius on a ten degree cone at supercritical speeds indicates that increasing nose radius promotes boundary layer transition and separation of laminar boundary layers.

Pointwise and scanning laser anemometer measurements in steady and unsteady separated turbulent boundary layers

Abstract: The physical features of steady and unsteady freestream separating turbulent boundary layers that have been determined by pointwise laser anemometer measurements are outlined. It is seen that the large-scale structures control the outer region's backflow behavior. Near the wall, the mean backflow velocity profile for both the steady and unsteady cases is found to scale on the maximum negative mean velocity and its distance from the wall. A description is given of a scanning laser anemometer that produces nearly instantaneous velocity profiles for examing the temporal features of these large-scale structures. Also described is a 'zero-wake' seeder that supplies particles to the outer shear layer and freestream flow with a minimal disturbance.

Turbulent Boundary Layers Developing over Compliant Surfaces.

Abstract: : The flow examined is the two-dimensional turbulent boundary layer over sinusoidal wavy surfaces The surfaces executed prescribed motion, that of a progressive water-wave The main conclusions are: The pressure dominates the small skin friction reduction that occurs At wavespeeds about 7/10 times the freestream speed and higher, the pressure becomes thrust producing for the case of two-dimensional waves When the waves are swept, the pressure becomes thrust producing as wavespeeds approach the component of the freestream in the direction normal to the wavefront Therefore the larger the sweep, the smaller the wavespeeds at which the pressure produces thrust Because of lack of flexible wall experiments, with well defined motion of the sinusoidal wall and high wavespeeds, comparisons were made with water-wave experiments Reasonable agreement was obtained for measured quantities inside the boundary layer It was estimated that drag reduction, for the cases considered, is small Limited comparison with available experiments indicates that the computed trends in the physical quantities are correct Computations using other approaches and pressure measurements on wavy walls with well defined motion are needed, in order to examine if the turbulence model used in this study is adequate for detailed quantitative predictions Based on the results of this study, a practical working system with a drag reducing surface with progressive waves does not seem feasible

On the Definition of Freestream Conditions in Wind-Tunnel Testing

Abstract: The concept of adaptive wall control in wind tunnels was advanced [4209, 4210] as a solution to the problem of boundary interference, especially in testing models in the transonic regime It had become apparent that wall effects could be significant, their character and magnitudes unknown, and reliable theory for the correction of measured data nonexistent Adaptive wall control is based on the observation that the existence of unconfined-flow conditions in a wind tunnel—and, in fact, a quantitative assessment of departure from such conditions—can be ascertained by calculation from data measured in the flow field Changes can then be made, systematically, in the tunnel boundary configuration, leading in an iterative way to unconfined flow, for any arbitrary model configuration

Investigation of the Three-Dimensional Transonic Flow around an Air Intake by a Finite-Volume Method for the Euler Equations

Abstract: The inviscid flow around a Pitot-type air intake has been calculated for a fixed freestream Mach number of 0.8 and three angles of attack (-6°, 0°, +6°). The ratio of static pressures at channel exit and infinity has been held constant. The analysis was carried out for an isolated air intake, i.e. the fuselage is omitted and the splitter plate is extended to an infinite flat plate. The Euler equations written in integral-conservation form are solved by a finite volume method applying a Runge-Kutta scheme with local time-stepping. The computational mesh is composed by a set of two-dimensional grids, constructed by a complex transformation combined with the solution of geometric partial differential equations. The main features of the predicted flow fields are described and compared with experimental results for a realistic inlet-fuselage combination.

A method of predicting unsteady turbulent flows and its application to diffusers with unsteady inlet conditions

Abstract: An integral method for computing the main characteristics of unsteady, incompressible, turbulent flows is developed. The method is applied to unsteady boundary layers with a prescribed freestream velocity and to diffusers with unsteady inlet conditions. For prescribed pressure gradient flows, the method gives as good results and is more than an order of magnitude faster than finite difference methods. For diff user flows, no other methods have been found in the literature.