NASA Logo

NTRS

NTRS - NASA Technical Reports Server

Collections About News Help Login
Back to Results
Effect of Wing Thickness and Sweep on the Oscillating Hinge-Moment and Flutter Characteristics of a Flap-Type Control at Transonic SpeedsFree-oscillation tests were made in the Langley high-speed 7- by 10-foot tunnel to determine the effects of wing thickness and wing sweep on the hinge-moment and flutter characteristics of a trailing-edge flap-type control. The untapered semispan wings had full-span aspect ratios of 5 and NACA 65A-series airfoil sections. Unswept wings having ratios of wing thickness to chord of 0.04, 0.06, 0.08, and 0.10 were investigated. The swept wings were 6 percent thick and had sweep angles of 30 deg and 45 deg. The full-span flap-type controls had a total chord of 50 percent of the wing chord and were hinged at the 0.765-wing-chord line. Tests were made at zero angle of attack over a Mach number range from 0.60 to 1.02, control oscillation amplitudes up to about 12 deg, and a range of control-reduced frequencies. Static hinge-moment data were also obtained. Results indicate that the control aerodynamic damping for the 4-percent-thick wing-control model was unstable in the Mach number range from 0.92 to 1.02 (maximum for these tests). Increasing the ratio of wing thickness to chord to 0.06, 0.08, and then to 0.10 had a stabilizing effect on the aerodynamic damping in this speed range so that the aerodynamic damping was stable for the 10-percent-thick model at all Mach numbers. The 6-percent-thick unswept-wing-control model generally had unstable aerodynamic damping in the Mach number range from 0.96 to 1.02. Increasing the wing sweep resulted in a general decrease in the stable aerodynamic damping at the lower Mach numbers and in the unstable aerodynamic damping at the higher Mach numbers. The one-degree-of-freedom control-surface flutter which occurred in the transonic Mach number range (0.92 to 1.02) for the 4-, 6-, and 8-percent-thick unswept-wing-control models could be eliminated by further increasing the ratio of thickness to chord to 0.10. Flutter could also be eliminated by increasing the wing sweep angle to either 30 deg or 45 deg. The magnitude of variation in spring moment derivative with Mach number at transonic speeds was decreased by either increasing the ratio of wing thickness to chord or increasing the wing sweep angle.
Document ID
19980232889
Acquisition Source
Headquarters
Document Type
Technical Memorandum (TM)
Authors
Moseley, William C., Jr. (NASA Langley Research Center Hampton, VA United States)
Gainer, Thomas G. (NASA Langley Research Center Hampton, VA United States)
Date Acquired
September 6, 2013
Publication Date
October 1, 1959
Subject Category
Aircraft Stability And Control
Report/Patent Number
NASA-TM-X-123
Distribution Limits
Public
Copyright
Work of the US Gov. Public Use Permitted.

Related Records

There are no records associated with this record.
No Preview Available