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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

Effects of Large Wing-Tip Masses on Oscillatory Stability of Wing Bending Coupled with Airplane Pitch

An examination of oscillatory stability for a straight-winged airplane with large concentrated wing-tip masses was made using wing-bending and airplane-pitching degrees of freedom and considering only quasi-steady aerodynamic forces. It was found that instability caused by coupling of airplane pitching and wing bending occurred for large ratios of effective wing-tip mass to total airplane mass and for coupled wing-bending frequencies near or below the uncoupled pitching frequency. Boundaries for this instability are given in terms of two quantities: (1) the ratio of effective tip mass to airplane mass, which can be estimated, and (2) the ratio of the coupled bending frequency to the uncoupled pitch frequency, which can be measured in flight. These boundaries are presented for various values of several airplane parameters.

Higdon, Donald T.↗
NASA NTRS1977

An elementary analysis of the effect of sweep, Mach number, and lift coefficient on wing-structure weight

Results are presented from an elementary analysis of the effect of sweep angle on the idealized structural weight of swept wings, with cruise Mach number M and lift coefficient C sub L as parameters. The analysis indicates that sweep is unnecessary for cruise Mach numbers below about 0.80, whereas for the higher subsonic speeds, a well defined minimum-weight condition exists at a sweep angle in the neighborhood of 35 deg or 40 deg, depending on M and C sub L. The results further indicate that wing-structure weight increases sharply with Mach number in the high subsonic range, with Mach 0.85 wings weighing half again as much as Mach 0.75 wings. Weight is also shown to increase with cruise lift coefficient, but the effect is not strong for the usual range of design lift coefficients. Minimum wing-structure weight is found to occur at a ratio of thickness to normal chord of about 18 percent, but it is concluded that the thickness ratio for optimum wing design would probably lie in the range of 12 to 15 percent.

Kyser, A. C.↗
NASA NTRS2017

Whirl Flutter Stability and Its Influence on the Design of the Distributed Electric Propeller Aircraft X- 57

This paper studies the whirl flutter stability of the NASA experimental electric propulsion aircraft designated the X-57 Maxwell. whirl flutter stability is studied at two flight conditions: sea level at 2700 RPM to represent take-off and landing and 8000 feet at 2250 RPM to represent cruise. Two multibody dynamics analyses are used: CAMRAD II and Dymore. The CAMRAD II model is a semi-span X-57 model with a modal representation for the wing/pylon system. The Dymore model is a semi-span wing with a propeller composed of beam elements for the wing/pylon system that airloads can be applied to. The two multibody dynamics analyses were verified by comparing structural properties between each other and the NASTRAN analysis. For whirl flutter, three design revisions of the wing and pylon mount system are studied. The predicted frequencies and damping ratio of the wing modes show good agreements between the two analyses. Dymore tended to predict a slightly lower damping ratio as velocity increased for all three dynamic modes presented. Whirl flutter for the semi-span model was not present up to 500 knots for the latest design, well above the operating range of the X-57.

Hoover, Christian B.↗
NASA NTRS1990

The aerodynamic design of the oblique flying wing supersonic transport

The aerodynamic design of a supersonic oblique flying wing is strongly influenced by the requirement that passengers must be accommodated inside the wing. It was revealed that thick oblique wings of very high sweep angle can be efficient at supersonic speeds when transonic normal Mach numbers are allowed on the upper surface of the wing. The goals were motivated by the ability to design a maximum thickness, minimum size oblique flying wing. A 2-D Navier-Stokes solver was used to design airfoils up to 16 percent thickness with specified lift, drag and pitching moment. A new method was developed to calculate the required pressure distribution on the wing based on the airfoil loading, normal Mach number distribution and theoretical knowledge of the minimum drag of oblique configurations at supersonic speeds. The wing mean surface for this pressure distribution was calculated using an inverse potential flow solver. The lift to drag ratio of this wing was significantly higher than that of a comparable delta wing for cruise speeds up to Mach 2.

Vandervelden, Alexander J. M.↗
NASA NTRS1992

Aerodynamic Spring and Damping of Free-Pitching Tips on a Semispan Wing

A test was conducted in the NASA Ames 7- by 10-Foot Wind Tunnel to derive aerodynamic spring and damping estimates for free-pitching tips on a semispan wing. The test model was a rectangular planform semispan wing with wing tips that had a single rigid-body pitch degree of freedom with respect to the inboard wing. A number of different tip planform geometries were tested, incorporating a range of quarter-chord sweep angles and taper ratios. The wing-tip dynamic response characteristics were measured at several wing angles of attack and tunnel dynamic pressures. The tip oscillations were initiated by releasing the tips from prescribed angles of attack. A new method to isolate Coulomb damping from aerodynamic damping from these tip-motion time histories is developed and applied. Correlations were performed between the experimentally derived wing-tip aerodynamic spring and damping values and predictions from a semiempirical analysis based on steady-state tip aerodynamic loads.

Young, Larry A.↗
NASA NTRS1981

Experimental study of the interaction between the wing of a subsonic aircraft and a nacelle of a high by-pass ratio engine

The oncoming of a new generation of subsonic transport aircraft (with supercritical wing and high by-pass ratio turbofans) led to an experimental study of wing nacelle jet pylon interference in transonic flow. To this end, a test set-up was developed at the ONERA S3Ch wind tunnel. The nacelle models represent a turbofan by means of two compressed air jets. The scale is 1/18.5. The nacelles are fixed on a thrust balance measuring afterbody thrust and discharge coefficients. The wing is located between the sidewalls of the test section. Pressures are measured through 456 holes located on 8 airfoils. Drag coefficient of the wing is obtained by wake survey. The following parameters can vary (1) wing/nacelle position; (2) upstream Mach number (from 0.3 to 0.8); (3) jet pressure ratio; (4) with/without pylon and (5) type of nacelle. Wing nacelle interference can be studied by means of total thrust drag analysis as a functon of the various parameters. The test set-up is described and examples of results are presented.

Levart, P.↗
NASA NTRS1959

Approach and Landing Investigation at Lift-Drag Ratios of 2 to 4 Utilizing a Straight-Wing Fighter Airplane

A series of landings was performed with a straight-wing airplane to evaluate the effect of low lift-drag ratios on approach and landing characteristics. Landings with a peak lift-drag ratio as low as 3 were performed by altering the airplane configuration (extending speed brakes, flaps, and gear and reducing throttle setting). As lift-drag ratio was reduced, it was necessary either to make the landing pattern tighter or to increase initial altitude, or both. At the lowest lift-drag ratio the pilots believed a 270 deg overhead pattern was advisable because of the greater ease afforded in visually positioning the airplane. The values of the pertinent flare parameters increased with the reduction of lift-drag ratio. These parameters included time required for final flare; speed change during final flare; and altitude, glide slope, indicated airspeed, and vertical velocity at initiation of final flare. The pilots believed that the tolerable limit was reached with this airplane in the present configuration, and that if, because of a further reduction in lift-drag ratio, more severe approaches than those experienced in this program were attempted, additional aids would be required to determine the flare-initiation point.

Matranga, Gene J.↗
NASA NTRS1940

Design charts relating to the stalling of tapered wings

An aid in airplane design, charts have been prepared to show the effects of wing taper, thickness ratio, and Reynolds number on the spanwise location of the initial stalling point. Means of improving poor stalling characteristics resulting from certain combinations of the variables have also been considered; additional figures illustrate the influence of camber increase to the wing tips, washout, central sharp leading edges, and wing-tip slots on the stalling characteristics. Data are included from which the drag increases resulting from the use of these means can be computed. The application of the data to a specific problem is illustrated by an example.

Soule, H A↗
NASA NTRS1991

Interference effects of very high bypass ratio nacelle installations on a low-wing transport

A twin-engine, low-wing transport model, with a supercritical wing designed for a cruise Mach number of 0.77 and a lift coefficient of 0.55 was tested in the 16-Foot Transonic Tunnel at NASA Langley Research Center. The purpose of this test was to compare the wing/nacelle interference effects of superfans (BPR 18) with the interference effects of advanced turbofans (BPR 6). Flow-through nacelles were used. Forces and moments on the complete model were measured using a strain gage balance, and extensive surface static pressure measurements were made on the model's wing, nacelles, and pylons. Results of the investigation indicate that superfan nacelles can be installed with approximately the same drag penalty as conventional turbofan nacelles.

Ingraldi, Anthony M.↗
NASA NTRS1974

Static aeroelasticity and the flying wing, revisited

The roll dynamic behavior of some wing configurations is examined. Analysis shows that the static divergence instability which occurs for symmetrical or clamped sweptforward wings is modified by roll freedom. The instability found in the present highly idealized analysis is still of an aperiodic nature, but at a speed above the conventional divergence speed. The magnitude of this difference is seen to be a function of the ratio of the wing roll mass moment of inertia to the mass moment of inertia of the wing about the roll axis, and this difference may be quite significant.

Weisshaar, T. A.↗
NASA NTRS1976

Flutter of asymmetrically swept wings

Two formulations of the oblique wing flutter problem are presented; one formulation allows only simple wing bending deformations and rigid body roll as degrees of freedom, while the second formulation includes a more complex bending-torsional deformation together with the roll freedom. Flutter is found to occur in two basic modes. The first mode is associated with wing bending-aircraft roll coupling and occurs at low values of reduced frequency. The second instability mode closely resembles a classical bending-torsion wing flutter event. This latter mode occurs at much higher reduced frequencies than the first. The occurrence of the bending-roll coupling mode is shown to lead to lower flutter speeds while the bending-torsion mode is associated with higher flutter speeds. The ratio of the wing mass moment of inertia in roll to the fuselage roll moment of inertia is found to be a major factor in the determination of which of the two instabilities is critical.

Weisshaar, T. A.↗