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Results for “CONTROLS, FLAP TYPE - COMPLETE WINGS”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Wing flap-type control effectiveness and effects of control hinge gap seals for a supercritical wing

Wind-tunnel tests were made to investigate the control effectiveness of wing trailing-edge flap-type aerodynamic controls for a supercritical wing. The tests determined the effects of spanwise flap control location, two different wing transition grit location patterns, magnitude of angular control deflection, control hinge moments and control hinge gap seals. Results show that the inboard flap position deflections caused interference rearward at the vertical tail, and that all flap positions for deflections in one direction only caused interference at the horizontal tail at a higher than design Mach number. Changes in wing transition grit pattern caused significant changes to static longitudinal characteristics and longitudinal control effectiveness.

Anglin, E. L.↗
NASA NTRS1951

Theoretical characteristics of two-dimensional supersonic control surfaces

The "Busemann second-order-approximation theory" for the pressure distribution over a two-dimensional airfoil in supersonic flow was used to determine some of the aerodynamic characteristics of uncambered symmetrical parabolic and double-wedge airfoils with leading-edge and trailing-edge flaps. The characteristics presented and discussed in this paper are: flap effectiveness factor, rate of change of hinge-moment coefficient with flap deflection, rate of change of the pitching-moment coefficient with flap deflection, rate of change of the pitching-moment coefficient about the mid chord with flap deflection, and the location of the center of pressure of the airfoil-flap combination.

MACH NUMBER EFFECTS - WING SECTIONS↗
NASA NTRS1951

An Investigation of Flow Characteristics at Mach Number 4.04 over 6- and 9-percent-thick Symmetrical Circular-arc Airfoils Having 30-percent-chord Trailing-edge Flaps

Section data obtained from wind-tunnel tests at Mach number 4.04 and Reynolds numbers of about 5.0 times 10 to the 6th power and 8.4 times 10 to the 6th power are presented for 6- and 9-percent-thick symmetrical circular-arc airfoils with 30 percent chord trailing-edge flaps. Pressure distributions, Schlieren photographs, and force and moment coefficients are included.

FLOW, VISCOUS↗
NASA NTRS1948

The effectiveness at high speeds of a 10-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section

This report contains the results of a high-speed wind-tunnel investigation of the effectiveness of a 10-percent-chord plain flap on the NACA 65-210 airfoil section. The results include an indication of the lift-producing characteristics and the effectiveness of the 10-percent-chord flap. From a comparison of the characteristics of the 10-percent-chord flap with those of a 20-percent-chord flap it was concluded that, although a reduction in flap-chord ratio from 0.20 to 0.10 lessens the severity of the effectiveness loss at supercritical speeds, the 20-percent-chord flap is more effective throughout the entire range of Mach numbers from 0.3 to 0.875.

CONTROLS, FLAP TYPE - WING SECTIONS↗
NASA NTRS1959

Effect of Wing Thickness and Sweep on the Oscillating Hinge-Moment and Flutter Characteristics of a Flap-Type Control at Transonic Speeds

Free-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.

Moseley, William C., Jr.↗
NASA NTRS1947

Investigation of a Spoiler-Type Lateral Control System on a Wing with Full-Span Flaps in the Langley 19-Foot Pressure Tunnel

Tests of a partial-span model of a large bomber-type air1ane were conducted to determine the. aerodynamic characteristics of the wing equipped with full-span flaps and a retractable spoiler end aileron lateral control system. The arrangement consisted of (1) a double slotted flap extending over aproximate1y 86 percent of the wing semispan, (2) a 20-percent constant-percentage-chord aileron extending from the outboard end of the flap to the wing tip, and (3) a retractable spoiler, located at the 65-percent wing-chord station and extending from approximately 63 percent of the wing semispan to the wing tip. In addition, tests were made of a wing vent (of 1 and 2 percent of the wing chord located directly behind the spoiler), perforations in the spoiler, a blot or cut-out along the lower edge of the spoiler and spoilers of various spans. With full-span flaps deflected and with the 2-percent vent open or closed the initial stalling of the wing occurred at the tips, but with the vents closed there probably would be no appreciable loss in lateral control until maximum lift was reached. The l-percent vent increased the rolling effectiveness of the spoiler at small spoi1er deflections, particularly at high angles of attack with flaps deflected. With flaps deflected the 2-percent vent caused a large reduction in both the wing lift and rolling effectiveness of the spoiler at large angles of attack. However, at small angle of attack the 2-percent vent increased the rolling effectiveness of the spoiler at small spoiler deflections. The simultaneous operation of the spoiler and vent (in contrast to a vent fixed in the wing) would result in a large increase in the effectiveness of the spoiler and would avoid any loss in wing lift as in a fixed vent arrangement. The tests of the spoiler modifications revealed that (1) the spoiler perforations reduced the rolling-moment and yawing-moment coefficients but caused the spoiler hinge-moment coefficients to become more positive; (2) the spoiler slot had no notable effect on the rolling-moment and yawing-moment characteristics but produced a positive increase in the spoiler hinge-moment coefficients at large spoiler deflections; (3) the effects produced by the individual modifications were additive when the various modifications were combined. In general, progressively decreasing the spoiler span by removing the segments from the inboard end of the spoiler caused a decrease in rolling effectiveness approximately proportional to the span of the segment.

WING FLAPS, SLOTTED & AILERONS↗