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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Body-freedom flutter of a 1/2-scale forward-swept-wing model, an experimental and analytical study

The aeroelastic phenomenon known as body-freedom flutter (BFF), a dynamic instability involving aircraft-pitch and wing-bending motions which, though rarely experienced on conventional vehicles, is characteristic of forward swept wing (FSW) aircraft was investigated. Testing was conducted in the Langley transonic dynamics tunnel on a flying, cable-mounted, 1/2-scale model of a FSW configuration with and without relaxed static stability (RSS). The BFF instability boundaries were found to occur at significantly lower airspeeds than those associated with aeroelastic wing divergence on the same model. For those cases with RSS, a canard-based stability augmentation system (SAS) was incorporated in the model. This SAS was designed using aerodynamic data measured during a preliminary tunnel test in which the model was attached to a force balance. Data from the subsequent flutter test indicated that BFF speed was not dependent on open-loop static margin but, rather, on the equivalent closed-loop dynamics provided by the SAS. Servo-aeroelastic stability analyses of the flying model were performed using a computer code known as SEAL and predicted the onset of BFF reasonably well.

Chipman, R.↗
NASA NTRS1985

Experiments on the stability of crossflow vortices in swept-wing flows

Experiments are conducted which demonstrate the feasibility of creating crossflow vortices in a flat-plate flow. A swept leading edge and contoured wind-tunnel walls produce a strong pressure gradient and a typical swept-wing flow on the plate. Detailed three-dimensional measurements, made within the boundary layer using hot-wire anemometry, are supplemented with different flow-visualization techniques. Data are presented on the mean flow characteristics as well as on the behavior of the crossflow vortices. Comparisons are made with theoretical flow-field and stability calculations.

Saric, W. S.↗
NASA NTRS1986

Divergence study of a high-aspect ratio, forward-swept wing

An experimental wind-tunnel study to determine the divergence characteristics of a high-aspect ratio, forward-swept wing has been conducted in the NASA Langley Research Center (LaRC) Transonic Dynamics Tunnel (TDT). The rectangular wing used for this study had a panel aspect ratio of 9.16 (lambda = 0 deg.) and the sweep angle could be set at lambda = 0 deg., -15 deg., -30 deg., -45 deg., or -60 deg. A rectangular wing tip shape was tested at each of these sweep angles. In addition, a tip shape parallel to the freestream flow was tested for a wing sweep angle of lambda = -45 deg. The root of the wing was cantilever mounted to the wall of the wind tunnel. Divergence conditions were measured at M = 0.4 for each sweep angle and tip configuration tested. Subcritical response techniques were used to extrapolate to the divergence conditions during the wind-tunnel test. The primary objective of this test was to obtain data which could be used to verify for this configuration the divergence prediction capability of an aeroelastic analysis code. Subsonic lifting surface theory (kernel function) aerodynamics are utilized by this particular code. The analytical predictions of divergence were found to be significantly conservative at all forward sweep angles. At lambda = -45 deg., the analysis was 14 percent conservative. The effect of the two tip shapes on the divergence dynamic pressure was predicted accurately by the analysis. The divergence condition for the tip shape parallel to the flow occurred at a dynamic pressure 14 percent higher than the divergence condition with a rectangular tip shape.

Cole, S. R.↗
NASA NTRS1989

In-flight boundary-layer transition measurements on a swept wing

Flight tests were conducted at three different altitudes to detect transition on a smoothed test region of a swept-wing business jet wing using surface hot-film sensors and sublimating chemicals. Strong influence of sweep angle on transition location was observed when the aircraft was flown at some sideslip conditions to simulate changes in effective wing sweep angle. No effects of engine noise on transition were measured when different engine power settings were used. Flight instrumentation and ground data analysis techniques are described. Correlation was obtained between the hot-film sensor signals and sublimating chemicals for transition detection. Crossflow vortices were observed for one flight condition. Results of analyzed data for various flight-test conditions are presented.

Ahmed, Anwar↗
NASA NTRS1957

Sidewash in the Vicinity of Lifting Swept Wings at Supersonic Speeds

Equations and charts have been presented for the conical part of the sidewash above and below swept wings at small angles of attack. This flow-field component is necessary in making calculations of the wing-induced forces and moments on stores, missiles, and pylons. All supersonic Mach numbers and wing leading-edge sweep angles have been considered.

AERODYNAMIC LOADS, BODIES↗
NASA NTRS1957

An analysis of the effects of aeroelasticity on static longitudinal stability and control of a swept-wing airplane

A theoretical analysis has been made of the effects of aeroelasticity on the static longitudinal stability and elevator angle required for balance of an airplane. The analysis is based on the familiar stability equation expressing the contribution of wing and tail to longitudinal stability. Effects of wing, tail, and fuselage flexibility are considered. Calculated effects are shown for a swept-wing bomber of relatively high flexibility.

Skoog, Richard B↗
NASA NTRS1957

Experimental and Predicted Longitudinal and Lateral-Directional Response Characteristics of a Large Flexible 35 Degree Swept-Wing Airplane at an Altitude of 35,000 Feet

Measured and predicted dynamic response characteristics of a large flexible swept-wing airplane to control surface inputs are presented for flight conditions of 0.6 to 0.85 Mach number at an altitude of 35,000 feet. The report is divided into two parts. The first part deals with the response of the airplane to elevator control inputs with principal responses contained in a band of frequencies including the longitudinal short-period mode and several symmetrical structural modes. The second part deals with the response of the airplane to aileron and rudder control inputs with principal responses contained in a band of frequencies including the dutch roll mode, the rolling mode, and three antisymmetrical structural modes.

Cole, Henry A , Jr↗
NASA NTRS1961

Transonic Investigation of Aerodynamic Characteristics of a Swept-Wing Fighter-Airplane Model with Leading-Edge Droop in Combination with Outboard Chord-Extensions and Notches

An investigation of the effects of several wing leading-edge modifications on the aerodynamic characteristics of a 45 degree swept-wing fighter-airplane model has been conducted in the Langley 16-foot transonic tunnel at low and high lifting conditions at Mach numbers from 0.85 to 1.03. The investigation included the determination of the effect on longitudinal stability and performance characteristics of wing leading-edge and chord-extension droops of 60 and 20 degrees chord-extension overhangs of 0.075c and 0.15c (where c inboard end of the 0.075c chord-extension to depths of 0.075c and 0.l25c, and indention of the model fuselage to conform partially to the supersonic area rule for a Mach number of 1.20. Lift, drag, and pitching-moment data were obtained for configurations with the tail on and off. Comparisons of data obtained from the present model with data from a configuration with leading-edge slats are included. Generally, the model wing modifications provided only slight improvements of the airplane longitudinal stability characteristics, but did substantially reduce the airplane drag coefficients at moderate and high lifting conditions.

Whitcomb, Charles F.↗
NASA NTRS1959

Effects of Outboard Thickened and Blunted Leading Edges on the Wave Drag of a 45 Degree Swept-Wing and Body Combination

An investigation to evaluate the effects of thickened and blunted leading-edge modifications on the wave drag of a swept wing has been made at Mach numbers from 0.65 to 2.20 and at a Reynolds number of 2,580,000 based on the mean aerodynamic chord of the basic wing. Two leading-edge designs were investigated and they are referred to as the thickened and the blunted modifications although both sections had equally large leading-edge radii. The thickened leading edge was formed by increasing the thickness over the forward 40 percent of the basic wing section. The blunted modification was formed by reducing the wing chords about 1 percent and by increasing the section thickness slightly over the forward 6 percent of the basic section in a manner to keep the wing sweep and volume essentially equal to the respective values for the basic wing. The basic wing had an aspect ratio of 3, a leading-edge sweep of 45 deg., a taper ratio of 0.4, and NACA 64AO06 sections perpendicular to a line swept back 39.45 deg., the quarter-chord line of these sections. Test results indicated that the thickened modification resulted in an increase in zero-lift drag coefficient of from 0.0040 to 0.0060 over values for the basic model at Mach numbers at which the wing leading edge was sonic or supersonic. Although drag coefficients of both the basic and thickened models were reduced at all test Mach numbers by body indentations designed for the range of Mach numbers from 1.00 to 2.00, the greater drag of the thickened model relative to that of the basic model was not reduced. The blunted model, however, had less than one quarter of the drag penalty of the thickened model relative to the basic model at supersonic leading-edge conditions (M greater or equal to root-2).

Holdaway, George H.↗
NASA NTRS1953

A study of the use of leading-edge notches as a means for improving the low-speed pitching-moment characteristics of a thin 45-degree swept wing of aspect ratio 4

In an effort to develop a simple but effective means for controlling the stall progression of swept wings, a low-speed investigation has been undertaken primarily to determine the effectiveness of leading-edge notches. The investigation included consideration of notch size, location, and shape, as well as combined effects of notch—chord-extension and notch-fence arrangements. The wing used in the study had 45° quarter-chord sweep, aspect ratio 4, taper ratio 0.3, and NACA 65A006 airfoil sections streamwise. The results of the investigation indicated that although substantial improvements were obtained in the wing pitching-moment characteristics at moderate angles of attack, by use of outboard leading-edge notches, none of these notches produced desirable wing stability characteristics throughout the angle-of-attack range. The wing pitching-moment characteristics were relatively insensitive to notch size and to small changes in span-wise location. The use of a notch at the inboard end of a chord-extension increased the effectiveness of the chord-extension near stall and produced the same results as a chord-extension having an equivalent inboard overhang. Small localized notch—chord-extension combinations were found to produce the same pitching-moment characteristics as much larger chord-extension configurations. A leading-edge root cutout used in conjunction with a small outboard notch—chord-extension combination resulted in fairly desirable overall wing pitching-moment characteristics.

Joseph Weil↗
NASA NTRS1994

Crossflow-Vortex Breakdown on Swept Wings: Correlation of Nonlinear Physics

The spatial evolution of cross flow-vortex packets in a laminar boundary layer on a swept wing are computed by the direct numerical simulation of the incompressible Navier- Stokes equations. A wall-normal velocity distribution of steady suction and blowing at the wing surface is used to generate a strip of equally spaced and periodic disturbances along the span. Three simulations are conducted to study the effect of initial amplitude on the disturbance evolution, to determine the role of traveling cross ow modes in transition, and to devise a correlation function to guide theories of transition prediction. In each simulation, the vortex packets first enter a chordwise region of linear independent growth, then, the individual packets coalesce downstream and interact with adjacent packets, and, finally, the vortex packets nonlinearly interact to generate inflectional velocity profiles. As the initial amplitude of the disturbance is increased, the length of the evolution to breakdown decreases. For this pressure gradient, stationary modes dominate the disturbance evolution. A two-coeffcient function was devised to correlate the simulation results. The coefficients, combined with a single simulation result, provide sufficient information to generate the evolution pattern for disturbances of any initial amplitude.

Joslin, R. D.↗
NASA NTRS1987

Subsonic longitudinal and lateral-directional characteristics of a forward-swept-wing fighter configuration at angles of attack up to 47 deg

Subsonic lateral-direction and longitudinal characteristics of a forward-swept-wing fighter configuration were examined in wind-tunnel tests at Mach numbers of 0.2 and 0.5 for angles of attack from -7 to 47 deg. and over a sidelslip range of +/- 15 deg. The effects of a canard, strakes, vertical tail, and leading- and trailing-edge flaps are examined. The canard and strakes both reduce asymmetric moments and side forces at zero sideslip for angles of attack up to about 30 deg. The canard has a small influence on lateral-directional stability; however, strakes produce a substantial reduction in lateral stability for angles of attack greater than about 20 deg. The vertical tail improves directional stability for angles of attack up to 30 deg. Deflection of the leading-edge flap to 20 deg. at high angles of attack on the strake and canard configurations degrades lateral and directional stability. Deflection of the trailing-edge flap to 20 deg. on the canard configuration generally increases lateral and directional stability at high angles of attack. Leading- and trailing-edge flaps on the wing-body and canard configurations are effective for increased lift only for angles of attack up to about 40 deg. The leading-edge flap remains effective on the strake configuration over the entire angle-of-attack range tested.

Mann, Michael J.↗
NASA NTRS1955

The dynamic-response characteristics of a 35 degree swept-wing airplane as determined from flight measurements

The longitudinal and lateral-directional dynamic-response characteristics of a 35 degree swept-wing fighter-type airplane determined from flight measurements are presented and compared with predictions based on theoretical studies and wind-tunnel data. Flights were made at an altitude of 35,000 feet covering the Mach number range of 0.50 to 1.04. A limited amount of lateral-directional data were also obtained at 10,000 feet. The flight consisted essentially of recording transient responses to pilot-applied pulsed motions of each of the three primary control surfaces. These transient data were converted into frequency-response form by means of the Fourier transformation and compared with predicted responses calculated from the basic equations. Experimentally determined transfer functions were used for the evaluation of the stability derivatives that have the greatest effect on the dynamic response of the airplane. The values of these derivatives, in most cases, agreed favorably with predictions over the Mach number range of the test.

Triplett, William C↗
NASA NTRS1958

Blowing-Type Boundary-Layer Control as Applied to the Trailing-Edge Flaps of a 35 Degree Swept-Wing Airplane

A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying blowing-type boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effects of boundary-layer control on the handling qualities and operation of the airplane, particularly during landing and take-off. The wind-tunnel and flight tests indicated that blowing over the flaps produced large increases in flap lift increment, and significant increases in maximum lift. The use of blowing permitted reductions in the landing approach speeds of as much as 12 knots.

Kelly, Mark W↗
NASA NTRS1958

Area-Suction Boundary-Layer Control as Applied to the Trailing-Edge Flaps of a 35 Degree Swept-Wing Airplane

A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying area-suction boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effect of boundary-layer control in the handling qualities and operation of the airplane, particularly during landing. The wind-tunnel and flight tests indicated that area suction applied to the trailing-edge flaps produced significant increases in flap lift increment. Although the flap boundary-layer control reduced the stall speed only slightly, a reduction in minimum comfortable approach speed of about 12 knots was obtained.

Cook, Woodrow L↗
NASA NTRS1974

Flutter analysis of swept-wing subsonic aircraft with parameter studies of composite wings

A computer program is presented for the flutter analysis, including the effects of rigid-body roll, pitch, and plunge of swept-wing subsonic aircraft with a flexible fuselage and engines mounted on flexible pylons. The program utilizes a direct flutter solution in which the flutter determinant is derived by using finite differences, and the root locus branches of the determinant are searched for the lowest flutter speed. In addition, a preprocessing subroutine is included which evaluates the variable bending and twisting stiffness properties of the wing by using a laminated, balanced ply, filamentary composite plate theory. The program has been substantiated by comparisons with existing flutter solutions. The program has been applied to parameter studies which examine the effect of filament orientation upon the flutter behavior of wings belonging to the following three classes: wings having different angles of sweep, wings having different mass ratios, and wings having variable skin thicknesses. These studies demonstrated that the program can perform a complete parameter study in one computer run. The program is designed to detect abrupt changes in the lowest flutter speed and mode shape as the parameters are varied.

Housner, J. M.↗