*James W. Wiggins*

*naca-tn-4185*

*Jan 1958*

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. In general, the variation of the damping-in-roll derivative with sweep angle at zero angle of attack was only in fair agreement with the predicted variation, inasmuch as the 32.6 degree sweptback wing showed more damping in roll at zero angle of attack in the Mach number range from 0.85 to 0.93 than any of the other plan forms. The predicted variation of the derivative at zero angle of attack with Mach number was in good agreement with the experimental trend to the critical Mach number. Contrary to predictions based on potential-flow theory, the yawing moment due to rolling was positive and the lateral force due to rolling was negative at the higher angles of attack throughout the range of Mach number for all configurations of the investigation. Presented herein is a method of estimating yawing moment due to rolling and lateral force due to rolling through the angle-of-attack range. The method is shown to be applicable over large ranges of leading-edge radii, wing thickness, and Mach number. The results indicate a loss of wing-tip suction within the ranges of Mach number and angle of attack investigated.

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