Flow Visualization Data Research Articles

Finned, multibody aerodynamic interference at transonic Mach numbers

A wind tunnel experiment involving single, double, and triple combinations of mutually interfering, generic, finned aircraft stores has been conducted. Each combination of stores was tested at Mach numbers from 0.60 to 1.20, at angles of attack from 0 to 20 deg for the single store, and from 0 to 10 deg for the double- and triplestore configurations. Extensive body and fin pressure and flow visualization data were obtained at each store location. The data indicate that finned, multibody interference (measured on the body) in the transonic region seems to be greatest at Mach numbers near 1, and subsides as the freestream Mach number increases or decreases from sonic conditions. Although somewhat dependent on fin orientation, the inboard fins seem to experience greater multibody interference at M^ = 0.60 than at Mach numbers closer to 1. An Euler solution for the doublestore configuration (x fin orientation) at A/^ = 0.95, 0 deg angle of attack was generated and compared with these data. The generally good correlation indicates that an Euler flow solver can yield accurate predictions of the location and magnitude of multibody interference on both the body and the fins, provided the viscous effects associated with these configurations remain small.

Experimental study of effects of forebody geometry on high angle-of-attack stability

A series of low-speed wind-tunnel tests on a generic airplane model with a cylindrical fuselage are made to investigate the effects of forebody shape and fineness ratio, and fuselage/wing proximity on static and dynamic lateral/directional stability. During the stability investigation ten forebodies were tested including three different cross-sectional shapes with fineness ratios of 2,3, and 4. In addition, the wing was tested at two longitudinal positions to provide a substantial variation in forebody/wing proximity. Conventional force tests were conducted to determine static stability characteristics, and single-degree-of-freedom free-to-roll tests were conducted to study the wing rock characteristics of the model with the various forebodies. Flow visualization data were obtained to aid in the analysis of the complex flow phenomena involved. The results show that the forebody cross-sectional shape and fineness ratio and forebody/wing proximity can strongly affect both static and dynamic (roll) stability at high angles of attack. These characteristics result from the impact of these factors on forebody vortex development, the behavior of the vortices in sideslip, and their interaction with the wing flowfield.

Study of multibody aerodynamic interference at transonic Mach numbers

A wind tunnel experiment involving single, double, and triple combinations of mutually interfering generic, unfinned aircraft stores has been conducted. Each combination of stores was tested at Mach numbers from 0.60 to 1.20 and at angles of attack from 0 to 25 deg for the single store and from 0 to 6 deg for the double and triple store configurations. Extensive axial and circumferential pressure and flow visualization data at each store location were obtained. Euler solutions for each configuration at 0 deg incidence have been generated and compared with experimental data. This comparison indicates an Euler flow solver can yield accurate predictions of the location and magnitude of multibody interference provided an appropriate grid is used and the viscous effects associated with these configurations remain small. The data indicate multibody interference in the transonic region increases as the freestream Mach number approaches 1 from either direction, and subsides as the Mach number moves away from sonic conditions. This interference is characterized by a large, localized reduction in pressure on the inboard surfaces of the bodies which results in forces that draw the configuration closer together.

A parametric analysis of vortex patterns visualized over airfoils in accelerating flow

Complex vertical patterns form over the surface of an airfoil in an accelerating flow started from rest. A parametric analysis of the development of the vortex patterns was conducted, using flow visualization data. We focused on the dependence of the vortex development on angle of attack and on Reynolds number. This analysis helps understanding of complex unsteady flows and it may serve as a reference for numerical models and for computer simulation

Comparison of experimental and computational techniques for plane mixing layers

Some studies of forced and unforced plane turbulent mixing layers have been conducted using two experimental and two computational approaches. The present paper contains an overall comparison and discussion of the measured and computed results. The experimental results include flow-visualization data using the smoke-laser technique and mean flow and turbulence measurements obtained with hot X wires and a two-component laser-Doppler velocimeter (LDV). The mean flow and turbulence results indicate that the two experimental techniques agree reasonably well for this shear flow. Two-dimensional computations of the measured mixing layers have also been conducted in a coordinated effort; one method uses the inviscid discrete vortex technique for a spatially developing layer and the other is based on an approximation to the Reynolds-averaged Navier–Stokes equations. The vortex method was found to give excellent results for the forced mixing layer while the Reynolds-averaged computations, with a modified turbulence model, were particularly successful at capturing the near-field viscous behavior that included the splitter plate wake.

Experimental flowfield visualization of a high alpha wing at Mach 1.62

Experimental oil-flow and tuft patterns and vapor-screen flow-visualization data were obtained on a cambered wing model at Mach = 1.62 for an angle-of attack range of 0-14 deg. These data were used as flow diagnostic tools along with surface-pressure and force data and full-potential theory calculations. A large separation bubble was found on the lower wing surface at low angle of attack. The high-angle-of-attack flowfield was characterized by a large attached-flow leading-edge expansion followed by a crossflow shock. At alpha = 14 deg the crossflow shock apparently induced discrete regions of streamwise separated flow, which were clearly indicated in the vapor-screen and oil-flow photographs.

Predictions of Endwall Losses and Secondary Flows in Axial Flow Turbine Cascades

This paper describes the development of a semi-empirical model for estimating end-wall losses. The model has been developed from improved understanding of complex endwall secondary flows, acquired through review of flow visualization and pressure loss data for axial flow turbomachine cascades. The flow visualization data together with detailed measurements of viscous flow development through cascades have permitted more realistic interpretation of the classical secondary flow theories for axial turbomachine cascades. The re-interpreted secondary flow theories together with integral boundary layer concepts are used to formulate a calculation procedure for predicting losses due to the endwall secondary flows. The proposed model is evaluated against data from published literature and improved agreement between the data and predictions is demonstrated.

Comparison of CW Doppler ultrasound spectra with the spectra derived from a flow visualization model

The methods and results of a study to determine the accuracy of continuous wave (CW) Doppler spectral recordings by comparison to the spectra derived from the flow profiles photographed simultaneously in a pulsatile flow visualization model are reported in this paper. A pulsatile pump produced a flow velocity waveform, similar to that seen in the human femoral artery, in a quartz glass tube. The velocity profiles, which were made visible by using a photochromic dye/laser technique, were photographed, and at the same time the instantaneous Doppler spectra were recorded. A comparison of the Doppler data and the photographed profiles gave the following results. The Doppler spectrograms and those reconstructed from the flow visualization data were quite similar. Excellent agreement was observed between the instantaneous maximum and mean Doppler waveforms. Individual spectra showed some differences and these differences were quantified by the novel application of certain statistical shape descriptor coefficients that are based on the estimation of the higher order moments of the spectra. The Doppler spectra are generally more skewed towards higher frequencies, narrower, and more peaked than the flow visualization spectra. Analysis of the assumptions and various sources of error lead to the conclusion that the differences were probably caused by ultrasound beam nonuniformity and the effects of refraction, causing a reduction of the beam field response at the tube edges. It is concluded that provided certain precautions are taken in the measurement technique, the CW Doppler ultrasound spectra fairly accurately represent the true velocity profile.

Assessment of preliminary prediction techniques for wing leading-edge vortex flows at supersonic speeds

The applicability of existing theoretical models for predicting wing leading edge separated flow characteristics for high-lift supersonic flight was examined with regard to previous experimental force, pressure and flow visualization data. Correlations of data on uncambered delta wings revealed that the upper surface normal force and minimum pressure coefficients decreased nonlinearly with increasing angles of attack. The attainable vacuum pressure decreased with increasing Mach number, while the lower surface normal force increased nonlinearly with increasing speed. The LISTAR and VORCAM linear theory aerodynamic codes generated predictions for comparison with the data. LISTAR displayed better agreement with measured vortex strength and position and lifting characteristics than did VORCAM. The Euler code SWINT was ill-suited to calculating wing performance in separated flows at high lift and low supersonic speeds.

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

The role of flow visualization in the study of afterbody and base flows

The role of the flow visualization in providing the necessary insight for the development of theoretical models of complex afterbody and base flows is described. Methods of calculating the turbulent base pressure for axisymmetric configurations are discussed. Emphasis is placed on the treatment of supersonic flows for cylindrical, boattail, and flare bodies as well as sharp and blunt cones and base flow nozzles. The guidance provided by flow visualization in the development of theoretical models for the transonic case is also discussed. The current interests in applying finite difference techniques to these-problems indicate that the extensive use of flow visualization data will continue.

The transfer of heat by natural convection between bodies and their enclosures

Natural convection heat transfer between concentrically located isothermal spherical, cylindrical and cubical inner bodies and their isothermal cubical enclosure was experimentally investigated. Comparisons were made with the existing data for these same inner bodies and a spherical enclosure. In addition, temperature distributions and flow visualization data were obtained for most of those geometries. Comparisons made between the spherical and cubical enclosure, for the same inner body types, showed that the cubical enclosure resulted in a larger Nusselt number for a given Rayleigh number and inner body size. Heat transfer data taken with a non-isothennal inner body showed that the isothermality conditions on the inner body could be relaxed. All of the heat transfer data was correlated with an average deviation of less than 14%. Overall, the effect of the enclosure shape was small so long as the appropriate length scale is employed.

Leeside flows over delta wings at supersonic speeds

An experimental investigation of the lee-side flow on sharp leading-edge delta wings at supersonic speeds has been conducted. Pressure data were obtained at Mach numbers from 1.5 to 2.8, and three types of flow-visualization data (oil-flow, tuft, and vapor-screen) were obtained at Mach numbers from 1.7 to 2.8 for wing leading-edge sweep angles from 52.5 deg to 75 deg. From the flow-visualization data, the lee-side flows were classified into seven distinct types and a chart was developed that defines the flow mechanism as a function of the conditions normal to the wing leading edge, specifically, angle of attack and Mach number. Pressure data obtained experimentally and by a semiempirical prediction method were employed to investigate the effects of angle of attack, leading-edge sweep, and Mach number on vortex strength and vortex position. In general, the predicted and measured values of vortex-induced normal force and vortex position obtained from experimental data have the same trends with angle of attack, Mach number, and leading-edge sweep; however, the vortex-induced normal force is underpredicted by 15 to 30 percent, and the vortex spanwise location is overpredicted by approximately 15 percent.

The Boundary Layer on an Axisymmetric Body With and Without Spin

The results from an experimental study of the boundary layer surrounding an axisymmetric model consisting of a 3-caliber secant ogive nose, a 2-caliber cylindrical midsection, and a 1-caliber 7-deg conical boattail are presented. Low-speed flow visualization data at zero angle of attack are shown for Reynolds numbers between 315,000 and 1,030,000 (based upon total body length), and ratios of peripheral to freestream velocity between 0 and 1.67. The smoke-flow photographs display the various stages of transition of the boundary layer. In addition, pressure distribution data are included for the nonspinning case. The pressure data are used in conjunction with the photographic data to present a physical picture of the boundary-layer development on the model.

Flow in a Rectangular Diffuser With Local Flow Detachment in the Corner Region

An experimental study was conducted in order to examine the nature of flow within a high-inlet-aspect-ratio rectangular diffuser with locally detached flow in the corner region, but attached flow elsewhere in the diffuser. The results include flow visualization data corresponding to tuft and oil flow patterns observed on the diffuser walls and flow angle data taken in the immediate vicinity of the corner. Axial mean velocity distributions and local wall shear stress profiles are also presented. Analysis of the results provides new insight into the structure of separated, three-dimensional corner flows, wherein combined attached and detached flow conditions exist simultaneously at each streamwise location.

Correlation of the detachment of two-dimensional turbulent boundary layers

An improved single-variable correlation is developed for incipient and full detachment, and for reattachment ,of turbulent boundary layers on two-dimensional surfaces. The correlation is shown to agree with 1) the wallwake law of Coles; 2) the criterion for detachment of Sandborn and Kline; 3) a large amount of old flowvisualization data; and 4) new measurements of time-averaged fraction of forward flow. Relationships among the correlation, the underlying physics, and the implications for computations are discussed. The results clarify long-standing confusion between incipient and full detachment of turbulent boundary layers with regard to both correlation and computation.

Experimental Investigation of a Wing with Controlled Midspan Flow Separation

A wing with a bounded region of separated flow at high angles of attack was tested in the McDonnell Aircraft Company low-speed wind tunnel. The model configuration was chosen to provide a test case for the develop- ment of an analytical method for the calculation of the flow about stalled wings. Extensive data were obtained on wing surface pressure, the boundaries of the separated flow bubble above and aft of the wing, and flow visualization. Flowfield velocity measurements were integrated to determine the displacement thicknesses of the viscous flow, which were subsequently used to determine an equivalent inviscid shape. The equivalency was verified by comparing the calculated potential flow pressures of this shape with the measured pressures. AVORABLE aerodynamic characteristics in the stalled flight regime are becoming increasingly important to modern fighter aircraft as flight envelopes are expanded to exploit the high-angle-of-attack combat environment. This requires a capability to design departure resistance into fighter aircraft in the early stages of development. Recent evidence for unswept wings suggests that controlled midspan separation, in concert with attached inboard and outboard flow, may significantly improve departure characteristics by softening the abrupt stall associated with other separation patterns.1-2 It is not obvious at this time that this type of controlled stall progression is the best design approach for increased departure resistance for fighter aircraft. However, it is desirable to determine analytically the stalling charac- teristics of fighter wings in order to supplement develop- mental wind-tunnel testing. To do this, it is necessary to predict the conditions under which flow separation will occur, regardless of the type of flow separation. In addition, it is necessary to predict the loads on the wings and control sur- faces when there are regions of separated flow present. Analytical techniques are being developed at McDonnell Aircraft Company to model regions of flow separation and to predict the aerodynamic forces under these conditions.3 To support this effort, it was necessary to provide experimental data for the description of the flow within the separation bubble and additional wing flowfield data. On the basis of a preliminary analysis, the required data were determined to be: 1) extensive surface pressure measurements to provide estimates of spanwise pressure gradients, 2) measurements of the boundaries of the separated flow bubble above and aft of the stalled wing, 3) an estimate of the shed vorticity, and 4) flow visualization data to identify regions of flow into and out of the bubble. A review of the pertinent experimental data indicated that no data were available that satisfied all of these requirements. Therefore, an experimental program was initiated, under MCAIR Independent Research and Development funding, to supply the necessary measurements.4 This paper discusses the results of this ex- perimental effort.

Remote Acoustic Detection of Zooplankton Response to Field Processes, Oceanographic Instrumentation, and Predators

A narrow-beam multifrequency acoustic backscattering system operating at discrete frequencies in the 10- to 600-kHz band has been used to detect predator–prey interaction, the passive response of Zooplankton to fluid processes such as internal waves, lee waves, hydraulic jumps, and the response of organisms to the presence of an oceanographic instrument in the water column. The acoustic flow visualization data illustrate the extent of the vertical and horizontal displacements and mixing of organisms by fluid processes. The data suggest that fluid processes may strongly influence predator–prey distribution in the continental shelf and sound regimes and that the interpretation of biological field data should include the influence of small-scale fluid motion on the vertical and horizontal distribution of organisms and nutrients. The predator–prey interaction data indicate that small organisms will move 7–20 m in the vertical to avoid larger organisms. Organisms appear to be able to detect the presence of oceanographic instruments to ranges of [Formula: see text]. At this time it is not known whether the organism detection range is related to the optical properties of the water column or other physical parameters.Key words: acoustic, Zooplankton, fluid processes, predator–prey interaction

A discrete vortex method for the non-steady separated flow over an airfoil

A discrete vortex method was used to analyse the separated non-steady flow about a cambered airfoil. The foil flow modelling is based on the thin lifting-surface approach, where the chordwise location of the separation point is assumed to be known from experiments or flow-visualization data. Calculated results provided good agreement when compared with the post-stall aerodynamic data of two airfoils. Those airfoil sections differed in the extent of travel of the separation point with increasing angle of attack. Furthermore, the periodic wake shedding was analysed and its time-dependent influence on the airfoil was investigated.

Shock Structure in Transonic Compressor Rotors

The inadequacy of axisymmetric theory in modeling the structure of the shock system in a transonic compressor rotor is shown with a simple model. A more accurate model for the flow near the sonic radius is proposed, which is supported by some flow visualization data and three-dimensional computations.