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A means for attenuating the vortex created at aircraft wingtips which consists of a retractable planar surface transverse to the airstream and attached downstream of the wingtip which creates a positive pressure gradient just downstream from the wing is presented. The positive pressure forces a break up of the rotational air flow of the vortex.
A method was developed for determining the aerodynamic loads on the tip of an infinitely thin, swept, cambered semi-infinite wing at an angle of attack which is operating subsonically in an inviscid medium and is subjected to a sinusoidal gust. Under the assumption of linearized aerodynamics, the loads on the tip are obtained by superposition of the steady aerodynamic results for angle of attack and camber, and the unsteady results for the response to the sinusoidal gust. The near field disturbance pressures in the fluid surrounding the tip are obtained by assuming a dipole representation for the loading on the tip and calculating the pressures accordingly. The near field pressures are used to drive a reduced form of the Navier-Stokes equations which yield the tip vortex formation. The combined viscid-inviscid analysis is applied to determining the pressures and examining the vortex rollup in the vicinity of an unswept, uncambered wing moving steadily at a Mach number of 0.2 at an angle of attack of 0.1 rad. The viscous tip flow calculation shows features expected in the tip flow such as the qualitatively proper development of boundary layers on both the upper and lower airfoil surfaces. In addition, application of the viscous solution leads to the generation of a circular type flow pattern above the airfoil suction surface.
Vortex velocities were measured in the wakes of four configurations of a 0.61-m span model of a B-747 aircraft. The wakes were generated by towing the model underwater in a ship model basin. Tangential and axial velocity profiles were obtained with a scanning laser velocimeter as the wakes aged to 35 span lengths behind the model. A 45 deg deflection of two outboard flight spoilers with the model in the landing configuration resulted in a 36 percent reduction in wake maximum tangential velocity, altered velocity profiles, and erratic vortex trajectories. Deployment of the landing gear with the inboard flaps in the landing position and outboard flaps retracted had little effect on the flap vortices to 35 spans, but caused the wing tip vortices to have: (1) more diffuse velocity profiles; (2) a 27 percent reduction in maximum tangential velocity; and (3) a more rapid merger with the flap vortices.
A method is disclosed for controlling the dispersion of effluents in a flowing liquid. A vortex generator for creating a distinct recirculating vortical flow is disposed in a flowing liquid and effluents are discharged into the vortical flow. The effluents are entrained in the vortical flow and by selectively positioning the vortex generator the dispersion of the entrained effluents can be controlled.
Estimates of the effectiveness of a model following type control system in reducing the roll excursion due to a wake vortex encounter were obtained from single degree of freedom computations with inputs derived from the results of wind tunnel, flight, and simulation experiments. The analysis indicates that the control power commanded by the automatic system must be roughly equal to the vortex induced roll acceleration if effective limiting of the maximum bank angle is to be achieved.
The dynamic interaction of aircraft wake vortices was investigated using both inviscid and viscous models. For the viscous model, a computer code was developed using a second-order closure model of turbulent transport. The phenomenon of vortex merging which results in the rapid aging of a vortex wake was examined in detail. It was shown that the redistribution of vorticity during merging results from both convective and diffusive mechanisms.
A novel two-stage shear-modulated hydraulic vortex amplifier (U.S. patent 3,520,317) has been fabricated and put through preliminary steady-state testing at the 1000 psi supply pressure level with flows up to 15 gpm. The invention comprises a conventional fluidic vortex power stage and a shear-modulated pilot stage. In the absence of any mechanical moving parts, water may be used as the hydraulic medium thus opening the way to many underseas applications. At blocked load, a control input from 0 to 150 psi was required to achieve an output from 0 to 900 psi; at wide-open load, a control input of 0 to 120 psi was needed to achieve an output from 0 to 15 gpm. The power stage has been found unsuitable for the proportional control mode because of its nonlinear performance in the intermediate load range and because of strong pressure fluctuations (plus or minus 150 psi) in the intermediate control range. The addition of the shear-modulated pilot stage improves intermediate load linearity.
A Boeing 747 aircraft flew 54 passes at low level over ground-based sensors. Vortex velocities were measured by a laser-Doppler velocimeter, an array of monostatic acoustic sounders, and an array of propeller anemometers. Flow visualization of the wake was achieved using smoke and balloon tracers. Preliminary results were obtained on the initial downwash field, the time for merging of the multiple vortices, the velocity fields, vortex decay, and the effects of spoilers and differential flap settings on the dissipation and structure of vortices.
A Lagrangian vorticity method for numerical simulation of three-dimensional separated flows with vortex filaments, about solid bodies, is discussed. The method is an extension of an earlier one for three-dimensional rotational flows away from solid boundaries e.g., jets, vortex rings, and aircraft trailing vortices after initial rollup. In this method an harmonic contribution to the velocity field is computed at each time step to ensure tangency of the velocity field at the boundary. The mechanics of the boundary layer are approximated with sufficient accuracy so that the separation lines are located correctly on the surface of the body.
A description is given of experiments which have been carried out in a circular air jet facility consisting of two settling chambers in sequence. Sinusoidal perturbations in the exit profile are introduced at controlled frequencies and amplitudes with the aid of a loudspeaker attached to the wall of the first chamber. It was found that vortex pairing in circular jets can occur in two distinct modes, including the shear layer mode and the jet mode. Amplitude variations, the conditions for strong vortex pairing, and the spectral evolution downstream are illustrated with the aid of graphs.
Results obtained through application of potential-flow theory based on a suction flow analogy for incompressible flow are compared to attached-flow theory and to experimental data acquired for sharp-edged delta wings. The error incurred in predicting structural loads on the basis of attached-flow theory is investigated. Force and moment coefficients and load distributions are used in the comparisons. The root bending moment can be estimated by assuming vortex lift concentrated at the wing leading edge. While this may entail some error at large angles of attack (owing to inboard movement of the center of vortex lift), it may furnish an upper bound in evaluating results based on other theories.
The paper describes the facilities and test procedures used in a series of wind-tunnel and full-scale flight investigations of the effectiveness of flight spoilers currently existing on wide-bodied transport jet aircraft when used as trailing vortex hazard alleviation devices. Examples of the results of such studies include the variation of trailing wing rolling-moment coefficient with downstream distance behind a B-747 airplane model with various segments of its flight spoilers deflected 45 deg, and comparisons with models without spoilers deflected. It is concluded that the existing flight spoilers on the B-747 are effective as trailing vortex attenuators.
The vortex lattice method introduced by Lamar and Gloss (1975) was applied to the prediction of subsonic aerodynamic characteristics of hypersonic body-wing configurations. The reliability of the method was assessed through comparison of the calculated and observed aerodynamic performances of two National Hypersonic Flight Research Facility craft at Mach 0.2. The investigation indicated that a vortex lattice model involving 120 or more panel elements can give good results for the lift and induced drag coefficients of the craft, as well as for the pitching moment at angles of attack below 10 to 15 deg. Automated processes for calculating the local slopes of mean-camber surfaces may also render the method suitable for use in preliminary design phases.
The results of a recent wind tunnel test program to evaluate the effectiveness of the Tip Air Mass Injection (TAMI) system in modifying the blade tip vortex occurring during helicopter flight is described with attention to the effect of this modification on the impulsive noise. The measurement program is explained, and the correlation between experimental and predicted results is discussed. Topics considered include the effect of descent rate on noise pressure time histories, the effect of air mass injection on noise, and the analysis based on a dB(A) weighted approach. Impulsive noise generated by the interaction of a helicopter rotor blade and the concentrated tip vortex during forward flight descent is a primary contributor to acoustic annoyance as it draws early attention to the presence of the helicopter.
The impact of reducing the wake vortex minimum separation required behind heavy jets on terminal area operation rate was analyzed. The effect on arrival saturation and steady state average delay was determined for various percentages mix of heavy and large jet traffic samples operating under various precision of interarrival spacing. Benefits increase with percentage of heavy aircraft and with precision of control. These results demonstrate the payoff possible from research to reduce the severity of the trailing vortex by aerodynamic means.
A status report is presented on research directed at reducing the vortex disturbances of aircraft wakes. The objective of such a reduction is to minimize the hazard to smaller aircraft that might encounter these wakes. Inviscid modeling was used to study trailing vortices and viscous effects were investigated. Laser velocimeters were utilized in the measurement of aircraft wakes. Flight and wind tunnel tests were performed on scale and full model scale aircraft of various design. Parameters investigated included the effect of wing span, wing flaps, spoilers, splines and engine thrust on vortex attenuation. Results indicate that vortives may be alleviated through aerodynamic means.
To determine the feasibility of altering the formation and decay of aircraft trailing vortexes through aerodynamic means, the test capabilities of two wind tunnels and two towing basins were used. The facilities, common models, and measurement techniques that were employed in the evaluation of vortex minimization concepts are described.
The theory and use of a laser velocimeter that makes simultaneous measurements of vertical and longitudinal velocities while rapidly scanning a flow field laterally are described, and its direct application to trailing wake-vortex research is discussed. Pertinent measurements of aircraft wake-vortex velocity distributions obtained in a wind tunnel and water towing tank are presented. The utility of the velocimeter to quantitatively assess differences in wake velocity distributions due to wake dissipating devices and span loading changes on the wake-generating model is also demonstrated.