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An investigation has been made of a complete airplane model having a wing with the quarter-chord line swept back 40 degrees, aspect ratio 2.50, and taper ratio 0.42 to determine its low-speed stability and control characteristics. The longitudinal stability investigation included stabilizer and tail-off tests with different wing dihedral angles (Gamma = 0 degrees and Gamma = -10 degrees) over an angle-of-attack range for the cruising and landing configurations and tests. with a high horizontal-tail location (Gamma = -10 degrees) for the cruising configuration. Tests were made of the wing alone and to determine the effect of wing end plates in pitch. Lateral stability characteristics were determined for the airplane with different geometric wing dihedrals, with end plates, and with several dorsal modifications. Tests were made with ailerons and spoilers to determine control characteristics.
A test was conducted in the NASA Ames 7- by 10-Foot Wind Tunnel to derive aerodynamic spring and damping estimates for free-pitching tips on a semispan wing. The test model was a rectangular planform semispan wing with wing tips that had a single rigid-body pitch degree of freedom with respect to the inboard wing. A number of different tip planform geometries were tested, incorporating a range of quarter-chord sweep angles and taper ratios. The wing-tip dynamic response characteristics were measured at several wing angles of attack and tunnel dynamic pressures. The tip oscillations were initiated by releasing the tips from prescribed angles of attack. A new method to isolate Coulomb damping from aerodynamic damping from these tip-motion time histories is developed and applied. Correlations were performed between the experimentally derived wing-tip aerodynamic spring and damping values and predictions from a semiempirical analysis based on steady-state tip aerodynamic loads.
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An experimental investigation has been conducted to determine the effect of wing leading edge sweep and wing translation on the aerodynamic characteristics of a wing body configuration at a free stream Mach number of about 6 and Reynolds number (based on body length) of 17.9 x 10 to the 6th power. Seven wings with leading edge sweep angles from -20 deg to 60 deg were tested on a common body over an angle of attack range from -12 deg to 10 deg. All wings had a common span, aspect ratio, taper ratio, planform area, and thickness ratio. Wings were translated longitudinally on the body to make tests possible with the total and exposed mean aerodynamic chords located at a fixed body station. Aerodynamic forces were found to be independent of wing sweep and translation, and pitching moments were constant when the exposed wing mean aerodynamic chord was located at a fixed body station. Thus, the Hypersonic Isolation Principle was verified. Theory applied with tangent wedge pressures on the wing and tangent cone pressures on the body provided excellent predictions of aerodynamic force coefficients but poor estimates of moment coefficients.
A method is presented for estimating the lift due to flap deflection at low angles of attack in incompressible flow. In this method provision is made for the use of incremental section-lift data for estimating the effectiveness of high-lift flaps. The method is applicable to swept wings of any aspect ratio or taper ratio. The present method differs from other current methods mainly in its ease of application and its more general application. Also included is a simplified method of estimating the lift-curve slope throughout the subsonic speed range.
The viscous transonic flow around a low-aspect-ratio wing has been computed using an implicit, three-dimensional, 'thin-layer' Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. The geometry chosen for this study is a low-aspect-ratio wing with large sweep, twist, taper, and camber. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. Using this grid, the flow around the wing was computed for a free-stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three-dimensional flow separation off the wing surface are observed. Particle path lines indicate that the three-dimensional flow separation on the wing surface is part of the roots of the tip-vortex formation. The lifting of the tip vortex before the wing trailing edge is clearly observed by following the trajectory of particles released around the wing tip.
An experimental investigation was performed to obtain the aerodynamic loading of a series of wingtips. The wingtips varied in taper ratio (1.0, 0.6, and 0.3) and in sweep angle (0, 20, 30, 35, and 45 deg). The wingtips were tested separately (wingtip only configuration) and while mounted on a semi-span wing. The quarter-chord sweep angle of the wing was varied from 30 deg forward sweep, to unswept, to 30 deg. aft sweep. Wingtip loading was obtained from a six component strainguage balance mounted in the wingtip. The effect of wing sweep and wingtip taper and wingtip sweep on the wingtip lift curve slope, drag polar, and pitching moment behavior is discussed. The obtained wingtip data was used to define blade tip platforms, suitable for use in the Free-Tip Rotor configuration. The Free-Tip Rotor has blade tips, which are allowed to weathervane into the blade tip's relative wind. A passive control moment is applied to the blade tip's pitch axis so as to cause blade tip motion around a prescribed null point providing a finite blade tip lift.
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