Search NASA⌕ Search

SEARCH · Search NASA

Results for “PITCHING MOMENTS”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1954

Effect of a Leading-edge Flap upon the Lift, Drag, and Pitching Moment of an Airplane Employing a Thin, Unswept Wing

The effects of deflecting full-span, constant-chord, leading-edge flaps, having either round or sharp leading edges, upon the lift, drag,. and pitching moment characteristics of a model of an interceptor-type aircraft have been determined experimentally at subsonic and supersonic speeds. Results indicate that the variations of lift with angle of attack and of pitching moment with lift were unaffected by either the shape of the flap leading edge or flap deflection. Deflection of the flaps having either a round or sharp leading edge increased the drag at zero lift at both subsonic and supersonic speeds. In spite of the increase in the drag at zero lift, however, deflection of the flaps increased the maximum lift-drag ratio at subsonic speeds and had no deleterious effect at supersonic speeds.

Heitmeyer, John C.↗
NASA NTRS1987

Predicted pitching moment characteristics of X-29A aircraft

The predicted pitching moment characteristics of the X-29A aircraft are presented for angles of attack from 0 to 20 deg. and Mach numbers of 0.2, 0.6, 0.9, 1.2, and 1.5 for altitudes of sea level, 4572 m (15,000 ft), 9144 m (30,000 ft), and 12,192 m (40,000 ft). These data are for both rigid and flexible aircraft and for the full range of control-surface positions. The characteristics were extracted from a nonlinear, symmetric, flexibilized wind tunnel data base.

Budd, Gerald D.↗
NASA NTRS1959

An Investigation of the Pitching-Moment Contribution of a High Horizontal Tail on an Unswept-Wing and Body Combination at Mach Numbers from 0.80 to 1.40

An investigation has been conducted to determine the effects of a high positioned horizontal tail on a wing-body configuration having a thin unswept wing of aspect ratio 3.09. Lift and pitching-moment coefficients were obtained for Mach numbers from 0.80 to 1.40 at Reynolds numbers of 1.0 and 1.5 million and for angles of attack to 20 deg. An experimental study of the pitching-moment contribution of the horizontal tail indicated that the marked destabilizing effect of the horizontal tail at high angles of attack for Mach numbers of 0.80 to 1.00 was associated with the formation of completely separated flow on the upper surface of the wing. Computations of the interference effects of the wing-body combination on the tail for Mach numbers of 0.80 and 0.94 and high angles of attack confirmed this conclusion. For a Mach number of 1.40, and high angles of attack, computations disclosed that the destabilizing effect primarily resulted from the trailing vortices of the wing. Two modifications to the basic wing plan form, which consisted of chord extensions, were generally unsuccessful in reducing the destabilizing contributions of the horizontal tail at high angles of attack.

Lippmann, Garth W.↗
NASA NTRS1949

Investigation of a Thin Straight Wing of Aspect Ratio 4 by the NACA Wing-flow Method : Lift and Pitching-moment Characteristics of the Wing Alone

This report presents measurements of the lift and pitching-moment characteristics of a straight wing of aspect ratio 4, taper ratio 0.5, having a symmetrical double-wedge airfoil section with a maximum thickness of 4.4-percent chord. The tests were conducted in the Mach number range 0.51 to 1.20, Reynolds numbers 380,000 to 660,000, by the NACA wing-flow method. The results are compared with theory and with wind-tunnel tests of a similar model. It is indicated that in the Mach number range 0.82 to 1.00 the model surface, profile, and test Reynolds number all would be very important considerations in any attempt to study or predict full-scale characteristics from small-model tests.

MACH NUMBER EFFECTS - COMPLETE WINGS↗
NASA NTRS1970

Static and dynamic pitching moment measurements on a family of elliptic cones at Mach number 11 in helium

Static and dynamic pitching moment measurements were made on a family of constant volume elliptic cones about two fixed axes of oscillation in the NAE helium hypersonic wind tunnel at a Mach number of 11 and at Reynolds numbers based on model length of up to 14 million. Viscous effects on the stability derivatives were investigated by varying the Reynolds number for certain models by a factor as large as 10. The models investigated comprised a 7.75 deg circular cone, elliptic cones of axis ratios 3 and 6, and an elliptic cone with conical protuberances.

Orlik-Rueckermann, K. J.↗
NASA NTRS1978

Closed-form equations for the lift, drag, and pitching-moment coefficients of airfoil sections in subsonic flow

Closed-form equations for the lift, drag, and pitching moment coefficients of two dimensional airfoil sections in steady subsonic flow were obtained from published theoretical and experimental results. A turbulent boundary layer was assumed to exist on the airfoil surfaces. The effects of section angle of attack, Mach number, Reynolds number, and the specific airfoil type were considered. The equations were applicable through an angle of attack range of -180 deg to +180 deg; however, above about + or - 20 deg, the section characteristics were assumed to be functions only of angle of attack. A computer program is presented which evaluates the equations for a range of Mach numbers and angles of attack. Calculated results for the NACA 23012 airfoil section were compared with experimental data.

Smith, R. L.↗
NASA NTRS1951

A semi-empirical method for calculating the pitching moment of bodies of revolution at low Mach numbers

A semiempirical method, in which potential theory is arbitrarily combined with an approximate viscous theory, for calculating the aerodynamic pitching moments for bodies of revolution is presented. The method can also be used for calculating the lift and drag forces. The calculated and experimental force and moment characteristics of 15 bodies of revolution are compared.

FLOW, VISCOUS↗
NASA NTRS1972

Numerical method for design of minimum-drag supersonic wing camber with constraints on pitching moment and surface deformation

A numerical method, based on linearized theory, for designing minimum-drag supersonic wing camber surfaces of arbitrary planform for a given lift, with options for constraining the pitching moment and/or the surface deformation at the trailing edge of the root chord and for selecting any desired combination of eight specified wing-loading distributions to be employed in the optimization procedure is presented. Two examples are given to illustrate applications of the method. The results indicate that relatively small drag penalties are incurred in designing wings to be self-trimming and to have a reasonable camber surface.

Sorrells, R. B.↗
NASA NTRS1959

Lift, Drag, and Pitching Moments of an Arrow Wing Having 80 Degree of Sweepback at Mach Numbers from 2.48 to 3.51 and Reynolds Numbers up to 11.0 Million

Measurements were made of the lift, drag, and pitching moments on an arrow wing (taper ratio of zero) having an aspect ratio of 1.4 and a leading-edge sweepback of 80 (degrees). The wing was designed to have a subsonic leading-edge and a Clark-Y airfoil with a thickness ratio of 12 percent of the chord perpendicular to the wing leading edge. The wing was tested both with and without the wing tips bent upward in an attempt to alleviate possible flow separation in the vicinity of the wing tips. Small jets of air were used to fix transition near the wing leading edge. Force results are presented for Mach numbers of 2.48, 2.75, 3.04, 3.28, and 3.51 at Reynolds numbers of 3.5 and 9.0 million and for a Mach number of 3.04 at a Reynolds number of 11.0 million. The measured aerodynamic characteristics are compared with those estimated by linear theory. The maximum lift-drag ratio measured was much less than that predicted. This difference is attributed to lack of full leading-edge thrust and to the experimental lift-curve slope being about 20 percent below the theoretical value.

Hopkins, Edward J.↗
NASA NTRS2016

Parametric Investigation of the Effect of Hub Pitching Moment on Blade Vortex Interaction (BVI) Noise of an Isolated Rotor

At the most fundamental level, main rotor loading noise is caused by the harmonically-varying aerodynamic loads (acoustic pressures) exerted by the rotating blades on the air. Rotorcraft main rotor noise is therefore, in principle, a function of rotor control inputs, and thus the forces and moments required to achieve steady, or "trim", flight equilibrium. In certain flight conditions, the ensuing aerodynamic loading on the rotor(s) can result in highly obtrusive harmonic noise. The effect of the propulsive force, or X-force, on Blade-Vortex Interaction (BVI) noise is well documented. This paper presents an acoustics parametric sensitivity analysis of the effect of varying rotor aerodynamic pitch hub trim moments on BVI noise radiated by an S-70 helicopter main rotor. Results show that changing the hub pitching moment for an isolated rotor, trimmed in nominal 80 knot, 6 and 12 deg descent, flight conditions, alters the miss distance between the blades and the vortex in ways that have varied and noticeable effects on the BVI radiated-noise directionality. Peak BVI noise level is however not significantly altered. The application of hub pitching moment allows the attitude of the fuselage to be controlled; for example, to compensate for the uncomfortable change in fuselage pitch attitude introduced by a fuselage-mounted X-force controller.

Parametric Investigation↗
NASA NTRS1959

An Experimental Investigation of the Effect of a Canard Control on the Lift, Drag, and Pitching Moment of an Aspect-Ratio 2.0 Triangular Wing Incorporating a Form of Conical Camber

The results of an experimental investigation to determine the effect of a canard control on the lift, drag, and pitching-moment characteristics of an aspect-ratio-2.0 triangular wing incorporating a form of conical camber are presented. The canard had a triangular plan form of aspect ratio 2.0 and was mounted in the extended chord plane of the wing. The ratio of the area of the exposed canard panels to the total wing area was 6.9 percent, and the ratio of the total areas was 12.9 percent. Data were obtained at Mach numbers from 0.70 to 2.22 through an angle-of-attack range from -6 deg to +18 deg with the canard on, and with the canard off. To provide a basis for comparison, the canard was also tested with a symmetrical wing having the same plan form, aspect ratio, and thickness distribution as the cambered wing. The results of the investigation showed that at the high subsonic speeds the gain in maximum lift-drag ratio achieved by camber was considerably reduced by the addition of a canard. At the supersonic speeds, the addition of the canard did not change the effect of camber on the maximum lift-drag ratios.

Menees, Gene P.↗
NASA NTRS1951

Lift, Drag, and Pitching Moment of Low-aspect-ratio Wings at Subsonic and Supersonic Speeds : Twisted and Cambered Triangular Wing of Aspect Ratio 2 with NACA 0003-63 Thickness Distribution

This report presents the results of an investigation to ascertain the lift, drag, and pitching moment of a wing-body combination having a triangular wing of aspect ratio 2 with NACA 0003-63 thickness distribution in streamwise planes and twisted and cambered for a nearly elliptical span load distribution. Results are shown for Mach numbers from 0.60 to 0.90 and from 1.30 to 1.70 at Reynolds numbers of 3.0 million and 7.5 million.

WINGS, COMPLETE - DESIGN VARIABLES↗