Showing papers on "Wing root published in 1976"

Flutter of Asymmetrically Swept Wings

Abstract: Two formulations of the oblique wing flutter problem are presented; one formulation allows only simple wing bending deformations and rigid body roll as degrees of freedom, while the second formulation includes a more complex bending-torsional deformation together with the roll freedom. Flutter is found to occur in two basic modes. The first mode is associated with wing bending-aircraft roll coupling and occurs at low values of reduced frequency. The second instability mode closely resembles a classical bending-torsion wing flutter event. This latter mode occurs at much higher reduced frequencies than the first. The occurrence of the bending-roll coupling mode is shown to lead to lower flutter speeds while the bending-torsion mode is associated with higher flutter speeds. The ratio of the wing mass moment of inertia in roll to the fuselage roll moment of inertia is found to be a major factor in the determination of which of the two instabilities is critical.

Wing and reduced airspeed system for jet aircraft

Abstract: A thin, high performance swept wing of the tapered type with an improved leading edge characterized by (1) camber that increases from a minimum near the wing root to a maximum near the wing tip, and (2) substantially a constant leading edge radius extending substantially across the wing span which defines a "blunt" contour. The wing in combination with a T-tail aircraft with a stick shaker/pusher activated by a rate of change of angle of attack sensor and optionally a strake between the leading edge and a wing tip tank which intrinsically combine to define a system that enhances aircraft performance by reducing minimum airspeed without impairing aircraft performance at high subsonic Mach (M) numbers.

A computer program to calculate the longitudinal aerodynamic characteristics of wing-flap configurations with externally blown flaps

Abstract: A vortex lattice lifting-surface method is used to model the wing and multiple flaps. Each lifting surface may be of arbitrary planform having camber and twist, and the multiple-slotted trailing-edge flap system may consist of up to ten flaps with different spans and deflection angles. The engine wakes model consists of a series of closely spaced vortex rings with circular or elliptic cross sections. The rings are normal to a wake centerline which is free to move vertically and laterally to accommodate the local flow field beneath the wing and flaps. The two potential flow models are used in an iterative fashion to calculate the wing-flap loading distribution including the influence of the waves from up to two turbofan engines on the semispan. The method is limited to the condition where the flow and geometry of the configurations are symmetric about the vertical plane containing the wing root chord. The calculation procedure starts with arbitrarily positioned wake centerlines and the iterative calculation continues until the total configuration loading converges within a prescribed tolerance. Program results include total configuration forces and moments, individual lifting-surface load distributions, including pressure distributions, individual flap hinge moments, and flow field calculation at arbitrary field points.

An experimental investigation of supersonic flow past a wedge-cylinder configuration

Abstract: An experimental investigation of supersonic flow past double-wedge configurations was conducted. Over the range of geometries tested, it was found that, while theoretical solutions both for a Type V pattern and for a Type VI pattern could be generated for a particular flow condition (as defined by the geometry and the free-stream conditions), the weaker, Type VI pattern was observed experimentally. More rigorous flow-field solutions were developed for the flow along the wing leading-edge. Solutions were developed for the three-dimensional flow in the plane of symmetry of a swept cylinder (which represented the wing leading-edge) which was mounted on a wedge (which generated the "bow" shock wave). A numerical code was developed using integral techniques to calculate the flow in the shock layer upstream of the interaction region (i.e., near the wing root). Heat transfer rates were calculated for various free stream conditions. The present investigation was undertaken to examine the effects of crossflow on the resultant flow-field and to verify the flow model used in theoretical calculations.

A Parametric Analysis of Winglet Effects.

Abstract: : A two part study was undertaken. In the first part, winglet effects were examined on a variety of wing planforms. A nonplaner lifting surface computer program was used to calculate aerodynamic coefficients including lift, induced drag, wing pitching moment, and wing root bending moment coefficients. Typical cruise flight conditions were examined and the coefficients calculated for the wing alone were compared to those obtained for the wing with winglets installed. The calculations were then repeated as the wing aspect ratio, sweep angle, and dihedral angle were varied. Winglet size and orientation remained constant. The percentage induced drag reduction was found to be the greatest on the wing with the highest aspect ratio and wing sweep angle (27.5% for an aspect ratio 7, 45 degree swept wing). Percentage induced drag reduction increased slightly with increasing positive dihedral angle. The greatest incremental drag reduction occurred with the lowest aspect ratio, highest swept wing. The second study examined the effects of winglet cant angle on induced drag reduction. Winglet cant angle on two separate wings was varied from -3 degrees to +3 degrees in an attempt to find an optimum value. The model used in this analysis ignored viscous effects and indicated that positive cant angles (leading edge inboard yielded the greatest reductions in induced drag. (Author)