Wind Tunnel Investigation At Low Speed Of A Twisted And Cambered Wing Swept Back 63 Degrees With Vortex Generators And Fences

A low-speed investigation has been conducted in the Langley two-dimensional low-turbulence tunnel to study a sweptback wing-root air-inlet configuration believed suitable for transonic-speed jet-powered airplanes. The test configurations consisted of a basic model with an NACA 64-008 wing with quarter-chord sweepback of 45 degrees mounted in the midwing position on a fuselage of fineness ratio 6.7, and an inlet model which had a triangular-shaped sweptback inlet installed in the wing root.

The experimental balance and pressure data obtained from tests of an untwisted variable-sweep wing with an outboard pivot exhibited nonlinearities in both total normalforce-coefficient and pitching-moment-coefficient curves for all sweep angles and fuselage conditions. These total effects have been traced back through the section data and pressure distributions to find the causes. The causes of these nonlinearities were found to be (1) separation of flow on the outer panel and (2) a leading-edge vortex flow on the inner panel. A fuselage added to the wing had little effect on the aerodynamic characteristics. Predictions of wing loadings and performance characteristics were made by using the modified Multhopp method and a comparison with experimental results indicated reasonable agreement. The changes in span loading as a result of leading-edge shed-vortex formation and flow separation at the higher angles of attack caused an increase in the induced-drag parameter.

A wind-tunnel investigation at low speed was made to determine the lateral control characteristics of a 45 degree sweptback untapered semispan wing of aspect ratio 1.59 equipped with various 25-percent-chord plain unsealed ailerons. Variations of the lateral-control data with aileron span and spanwise location were determined through a large aileron-deflection range for various angles of attack up to about the angle of maximum lift.

An investigation was performed in the Langley high-speed 7- by 10-foot tunnel in order to determine the rolling derivatives for swept-wing-body configurations at angles of attack from 0 degrees to 13 degrees and at high subsonic Mach numbers. The wings had sweep angles of 3.6 degrees, 32.6 degrees, 45 degrees, and 60 degrees at the quarter-chord line, an aspect ratio of 4, a taper ratio of 0.6, and an NACA 65A006 airfoil section parallel to the free stream. The results indicate a reduction in the damping-in-roll derivative at the higher test angles of attack. Of the wings tested, instability of the damping-in-roll derivative was experienced over the largest ranges of angle of attack and Mach number for the 32.6 sweptback wing.

Tunnel-wall corrections for the induced upwash velocity for a swept wing completely spanning a rectangular wind tunnel are included in the appendix.