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Flow visualization and containment tests in directed-wall-jet vortex tube with radial outflow and moderate superimposed axial flows
An experimental investigation was conducted to assess the vortex flow-field interactions on an advanced, twin-jet fighter aircraft configuration at high angles of attack. Flow-field surveys were conducted on a small-scale model in the Northrop 0.41 - by 0.60-meter water tunnel and, where appropriate, the qualitative observations were correlated with low-speed wind tunnel data trends obtained on a large-scale model of the advanced fighter in the NASA Langley Research Center 30- by 60-foot (9.1- by 18.3-meter) facility. Emphasis was placed on understanding the interactions of the forebody and LEX-wing vortical flows, defining the effects on rolling moment variation with sideslip, and identifying modifications to control or regulate the vortex interactions at high angles of attack. The water tunnel flow visualization results and wind tunnel data trend analysis revealed the potential for strong interactions between the forebody and LEX vortices at high angles of attack. In particular, the forebody flow development near the nose could be controlled by means of carefully-positioned radome strakes. The resultant strake-induced flow-field changes were amplified downstream by the more powerful LEX vortical motions with subsequent large effects on wing flow separation characteristics.
Axisymmetric flow of rotating stream analyzed to determine conditions allowing isolated vortex breakdown to develop
We report numerical simulations of the nonlinear dynamics of Josephson vortices driven by strong dc currents in layered superconductors. Dynamic equations for interlayer phase differences in a stack of coupled superconducting layers were solved to calculate a drag coefficient η(J) of the vortex as a function of the perpendicular dc current density J. It is shown that Cherenkov radiation produced by a moving vortex causes significant radiation drag increasing η(v) at high vortex velocities v and striking instabilities of driven Josephson vortices moving faster than a terminal velocity v c . The steady-state flux flow breaks down at v > v c as the vortex starts producing a cascade of expanding vortex-antivortex pairs evolving into either planar macrovortex structures or branching flux patterns propagating both along and across the layers. This vortex-antivortex pair production triggered by a rapidly moving vortex is most pronounced in a stack of underdamped planar junctions where it can occur at J > J s well below the interlayer Josephson critical current density. Both v c and J s were calculated as functions of the quasiparticle damping parameter, and the dc magnetic field was applied parallel to the layers. The effects of vortex interaction on the Cherenkov instability of moving vortex chains and lattices in annular stacks of Josephson junctions were considered. It is shown that a vortex driven by a current density J > J s in a multilayer of finite length excites self-sustained large-amplitude standing waves of magnetic flux, resulting in temporal oscillations of the total magnetic moment. We evaluated a contribution of this effect to the power W radiated by the sample and showed that W increases strongly as the number of layers increases. Furthermore, these mechanisms can result in nonlinearity of the c-axis electromagnetic response and contribute to THz radiation from the layered cuprates at high dc current densities flowing perpendicular to the ab planes.
A discrete vortex method was used to analyze 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 analyzed and its time-dependent influence on the airfoil was investigated.
The free-stream interference caused by the flow through the slotted walls of the test sections of transonic wind tunnels has continuously a problem in transonic tunnel testing. The adaptive-wall transonic tunnel is designed to actively control the near-wall boundary conditions by sucking or blowing through the wall. In order to make the adaptive-wall concept work, parameters for computational boundary conditions must be known. These parameters must be measured with sufficient accuracy to allow numerical convergence of the flow field computations and must be measured in an inviscid region away from the model that is placed inside the wind tunnel. The near-wall flow field was mapped in detail using a five-port cone probe that was traversed in a plane transverse to the free-stream flow. The initial experiments were made using a single slot and recent measurements used multiple slots, all with the tunnel empty. The projection of the flow field velocity vectors on the transverse plane revealed the presence of a vortex-like flow with vorticity in the free stream. The current research involves the measurement of the flow field above a multislotted system with segmented plenums behind it, in which the flow is controlled through several plenums simultaneously. This system would be used to control a three-dimensional flow field.
Euler and Navier-Stokes solutions of the supersonic shear flow past a circular cylinder are obtained. These solutions are used to study the basic flow structure around the cylinder. Both the inviscid and viscous calculations show the formation of a large recirculating flow region around the front stagnation point. The calculations further show that the overall size of the recirculating region is approximately the same for the Euler and Navier-Stokes solutions but the inside structure is quite different. The inviscid flow shows only one vortex whereas the viscous flow shows two vortices inside the recirculating flow region. The inner vortex in the Navier-Stokes solution is formed primarily due to the viscous effects near the body surface and its size depends upon the Reynolds number. It is found that with increasing Reynolds number, the inner vortex diminishes in size and the Navier-Stokes solution asymptotically approaches the Euler solution. These results indicate that the Euler equations may correctly predict certain high Reynolds number separation phenomenon in flows with natural inviscid vorticity source.
The suction analogy concept of Polhamus for predicting vortex lift in conjunction with an appropriate potential-flow solution is called the present method. This method is applied to an aspect ratio 0.25 sharp-edge delta wing from a Mach number of 0.143 to 10.4 in free air and at 0.074 in ground effect, and also to an aspect ratio 0.35 triangular crosssectional body at a Mach number of 6.9. The models had subsonic leading edges at the test Mach numbers. Vortex-flow effects could be neither confirmed nor denied to exist at high speeds because of the lack of flow visualization above a Mach number of 0.143. The data, however, could be better predicted by including a vortex-flow effect, although not always to the extent predicted from the present method because of the presence of actual and hypothesized unmodeled flow situations. The method of Nenni and Tung (NASA CR-1860) tended to confirm the existence of vortex flow at hypersonic speeds. The hypersonic-tangent-cone method predicted best the delta-wing results over the test angle-of-attack range and hypersonic Mach number range and did equally as well as the present method for the triangular body.
The application of a new, general, potential flow computational technique to the solution of the subsonic, three-dimensional flow over wings with leading-edge vortex separation is presented. The present method is capable of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings of rather arbitrary planform. The wing geometry is arbitrary in the sense that leading and trailing edges may be curved or kinked and the wing may have arbitrary camber and twist. The method employs an inviscid flow model in which the wing, the rolled-up vortex sheets, and the wake are represented by piecewise continuous quadratic doublet sheet distributions. The Kutta condition is imposed along all wing edges. Strengths of the doublet distributions as well as shape and position of the free fortex sheet spirals are computed in iterative fashion starting with an assumed initial sheet geometry. The method is verified by numerous computed results.
Fast Fourier transform (FFT) techniques are applied to the problem of finding the flow due to source or vortex distributions in the field outside an airfoil or other two-dimensional body. Either the complex potential or the complex velocity may be obtained to a high order of accuracy, with computational effort similar to that required by second-order fast Poisson solvers. These techniques are applicable to general flow problems with compressibility and rotation. An example is given of their use for inviscid compressible flow.
This paper describes an analytical model of an incompressible, isothermal reacting jet in crossflow. The model represents the flow in the jet cross-section by a counter rotating vortex pair, a flow structure that has been observed to dominate the jet behavior. The reaction surface surrounding the fuel jet is represented as a composite of strained diffusion flames that are stretched and deformed by the vortex pair flow. The results shed new light on the interaction between the vortex pair circulation and flame structure evolution and their relation to the concept of entrainment.
Taylor vortices and turbulence in flow between eccentric rotating cylinders
The buffet response of the flexible twin-tail configuration of the F/A-18 and a generic F-111 aircraft are computationally simulated and experimentally validated. The problem is a multidisciplinary one which requires the sequential solution of three sets of equations on a multi-block grid structure. The first set is the unsteady, compressible, full Navier-Stokes equations. The second set is the aeroelastic equations for bending and torsional twin-tail responses. The third set is the grid-displacement equations which are used to update the grid coordinates due to the tail deflections. The computational models consist of a 76 deg. swept back, sharp edged delta wing of aspect ratio of one and a swept-back F/A-18 or F-117 twin-tail. The configuration is pitched at 30 deg. angle-of-attack. The problem is solved for the initial flow conditions with the twin tails kept rigid. Next, the aeroelastic equations of the tails are turned on along with the grid-displacement equations to solve for the bending and torsional tails responses due to the unsteady loads produced by the vortex breakdown flow of the leading-edge vortex cores of the delta wing. Several spanwise locations of the twin tails are investigated. The computational results are validated using several existing experimental data.
Analytical investigations of fluid dynamics problems of relevance to the gaseous core nuclear reactor program are presented. The vortex type flow which appears in the nuclear light bulb concept is analyzed along with the fluid flow in the fuel inlet region for the coaxial flow gaseous core nuclear reactor concept. The development of numerical methods for the solution of the Navier-Stokes equations for appropriate geometries is extended to the case of rotating flows and almost completes the gas core program requirements in this area. The investigations demonstrate that the conceptual design of the coaxial flow reactor needs further development.
Because the potential flow suction along the leading and side edges of a planform can be used to determine both leading- and side-edge vortex lift, the present investigation was undertaken to apply the vortex-lattice method to computing side-edge suction force for isolated or interacting planforms. Although there is a small effect of bound vortex sweep on the computation of the side-edge suction force, the results obtained for a number of different isolated planforms produced acceptable agreement with results obtained from a method employing continuous induced-velocity distributions. By using the method outlined, better agreement between theory and experiment was noted for a wing in the presence of a canard than was previously obtained.
Maps of helicity density and normalized helicity useful as graphical representations of important features of three-dimensional flow fields containg vortexes. Emphasize complicated and important parts of flow field, identify vortexes, differentiate between primary and secondary vortexes, indicate sense of swirling motion, locate free singular points, and trace vortex-core streamlines emanating from these points.
The trend towards slender wing aircraft for supersonic cruise and the early chronology of research directed towards their vortex-lift characteristics are briefly reviewed. An overview of the development of vortex-lift theoretical methods is presented, and some current computational and experimental challenges related to the viscous flow aspects of this vortex flow are discussed.
The onset of vortex breakdown in compressible flows is investigated analytically for the case in which the flow is axially symmetric, the vortex is isolated, its axis is parallel to the main flow, and the vortex radius is small compared to the breakdown length. The conservation equations for mass, momentum, and energy are formulated and solved numerically using a finite-difference scheme, as described by Krause (1985); numerical results are presented in graphs and briefly characterized.