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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 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

Lift enhancement by an externally trapped vortex

A theoretical study is made of the performance capabilities of a lift concept that utilizes a spanwise vortex over the upper surface of the wing. The vortex is generated by a vertical flap near the leading edge of the wing and maintained by suction through orifices in endplates at the wingtip. The analysis approximates the three-dimensional flow field with a two-dimensional configuration that is mapped by conformal transformation into the flow about a circle. Theoretical solutions for a range of flap and orifice configurations predict that section lift coefficients up to around 10 can be achieved. It is concluded that such a lift concept is applicable to STOL aircraft if the vortex can be adequately stabilized and if the endplate suction can be generated efficiently.

Rossow, V. J.↗
NASA NTRS1987

Lift distribution and velocity field measurements for a three-dimensional, steady blade/vortex interaction

A wind tunnel experiment simulating a steady three-dimensional helicopter rotor blade/vortex interaction is reported. The experimental configuration consisted of a vertical semispan vortex-generating wing, mounted upstream of a horizontal semispan rotor blade airfoil. A three-dimensional laser velocimeter was used to measure the velocity field in the region of the blade. Sectional lift coefficients were calculated by integrating the velocity field to obtain the bound vorticity. Total lift values, obtained by using an internal strain-gauge balance, verified the laser velocimeter data. Parametric variations of vortex strength, rotor blade angle of attack, and vortex position relative to the rotor blade were explored. These data are reported (with attention to experimental limitations) to provide a dataset for the validation of analytical work.

Dunagan, Stephen E.↗
NASA NTRS1992

Numerical simulation of STOL operations using thrust-vectoring

The flow field about a delta wing equipped with thrust reverser jets in slow speed flight near the ground has been computed. Results include the prediction of the flow about the delta wing at four fixed heights above the ground, and a simulated landing, in which the delta wing descends towards the ground. Comparison of computed and experimental lift coefficients indicates that the simulations can capture at least the qualitative trends in lift-loss encountered by thrust-vectoring aircraft operating in ground effect.

Chawla, Kalpana↗
NASA NTRS1993

Numerical simulation of a powered-lift landing

The flow field about a delta wing equipped with thrust reverse jets in slow speed flight near the ground has been computed. Results include the prediction of the flow about the delta wing at four fixed heights above the ground, and simulated landing, in which the delta wing descends towards the ground. Comparison of computed and experimental lift coefficients indicates that the simulations can capture at least the qualitative trends in lift-loss encountered by thrust-vectoring aircraft operating in ground effect.

Chawla, Kalpana↗
NASA NTRS1961

Effects at Mach Numbers of 1.61 and 2.01 of Camber and Twist on the Aerodynamic Characteristics of Three Swept Wings Having the Same Planform

An investigation has been made at Mach numbers of 1.61 and 2.01 to determine the aerodynamic characteristics of three wings having a sweepback of 50 deg at the quarter-chord line, a taper ratio of 0.20, an NACA 65A005 thickness distribution, and an aspect ratio of 3.5. One wing was flat, one had at each spanwise station an a = 0 mean line modified to have a maximum height of 4-percent chord, and one had a linear variation of twist with 6 deg of washout at the tip. Tests were made with natural and fixed transition at Reynolds numbers ranging from 1.2 x 10(exp 6) to 3.6 x 10(exp 6) through an angle-of-attack range of -20 deg to 20 deg. When compared with the flat wing, the effect of the linear variation of twist with 6 deg of washout at the tip was to increase the lift-drag ratio when the leading edge was subsonic; but little increase in lift-drag ratio was obtained when the leading edge was supersonic. Pitching moment was increased and gave a positive trim point without greatly affecting the rate of change of pitching moment with lift coefficient. For the cambered wing the high minimum drag resulted in comparatively low lift-drag ratios. In addition, the pitching moments were decreased so that a negative trim point was obtained.

Landrum, Emma Jean↗
NASA NTRS1975

Effect of wing planform and canard location and geometry on the longitudinal aerodynamic characteristics of a close-coupled canard wing model at subsonic speeds

A generalized wind-tunnel model with canard and wing planforms typical of highly maneuverable aircraft was tested in the Langley 7- by 10-foot high-speed tunnel at a Mach number of 0.30 to determine the effect of canard location, canard size, wing sweep, and canard strake on canard-wing interference to high angles of attack. The major results of this investigation may be summarized as follows: the high-canard configuration (excluding the canard strake and canard flap), for both the 60 deg and 44 deg swept leading-edge wings, produced the highest maximum lift coefficient and the most linear pitching-moment curves; substantially larger gains in the canard lift and total lift were obtained by adding a strake to the canard located below the wing chord plane rather than by adding a strake to the canard located above the wing chord plane.

Gloss, B. B.↗
NASA NTRS1978

Three-dimensional canard-wing shape optimization in aircraft cruise and maneuver environments

This paper demonstrates a numerical technique for canard-wing shape optimization at two operating conditions. For purposes of simplicity, a mean surface wing paneling code is employed for the aerodynamic calculations. The optimization procedures are based on the method of feasible directions. The shape functions for describing the thickness, camber, and twist are based on polynomial representations. The primary design requirements imposed restrictions on the canard and wing volumes and on the lift coefficients at the operating conditions. Results indicate that significant improvements in minimum drag and lift-to-drag ratio are possible with reasonable aircraft geometries. Calculations were done for supersonic speeds with Mach numbers ranging from 1 to 6. Planforms were mainly of a delta shape with aspect ratio of 1.

De Silva, B. M. E.↗
NASA NTRS1995

Experimental Study of Lift-Enhancing Tabs on a Two-Element Airfoil

The results of a wind-tunnel test are presented for a two-dimensional NASA 63(sub 2)-215 Mod B airfoil with a 30% chord single-slotted flap. The use of lift-enhancing tabs (similar to Gurney flaps) on the lower surface near the trailing edge of both elements was investigated on four nap configurations. A combination of vortex generators on the flap and lift-enhancing tabs was also investigated. Measurements of surface-pressure distributions and wake profiles were used to determine the aerodynamic performance of each configuration. By reducing flow separation on the flap, a lift-enhancing tab at the main-element trailing edge increased the maximum lift by 10.3% for the 42-deg flap case. The tab had a lesser effect at a moderate flap deflection (32 deg) and adversely affected the performance at the smallest flap deflection (22 deg). A tab located near the flap trailing edge produced an additional lift increment for all flap deflections. The application of vortex generators to the flap eliminated lift-curve hysteresis and reduced flow separation on two configurations with large flap deflections (greater than 40 deg). A maximum-lift coefficient of 3.32 (17% above the optimum baseline) was achieved with the combination of lift-enhancing tabs on both elements and vortex generators on the flap.

Storms, Bruce L.↗
NASA NTRS1954

Static Longitudinal and Lateral Stability and Control Characteristics of a 1/15-Scale Model of the Grumman F9F-9 Airplane at a Mach Number of 1.41

An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the static stability and control and drag characteristics of a l/l5-scale model of the Grunman F9F-9 airplane. The effects of alternate fuselage shapes, wing camber, wing fences, and fuselage dive brakes on the aerodynamic characteristics were also investigated. These tests were made at a Reynolds number of 1.96 x l0 (exp 6) based on the wing mean aerodynamic chord of 0.545 foot. The basic configuration had a static margin of stability of 38.4 percent of the mean aerodynamic chord and a minimum drag coefficient of 0.049. For the maximum horizontal tail deflection investigated (-l0 deg), the maximum trim lift coefficient was 0.338. The basic configuration had positive static lateral stability at zero angle of attack and positive directional control throughout the angle-of-attack range investigated up to ll deg.

Palazzo, Edward B.↗
NASA NTRS2016

Computational Design of a Krueger Flap Targeting Conventional Slat Aerodynamics

In this study, we demonstrate the design of a Krueger flap as a substitute for a conventional slat in a high-lift system. This notional design, with the objective of matching equivalent-mission performance on aircraft approach, was required for a comparative aeroacoustic study with computational and experimental components. We generated a family of high-lift systems with Krueger flaps based on a set of design parameters. Then, we evaluated the high-lift systems using steady 2D RANS simulations to find a good match for the conventional slat, based on total lift coefficients in free-air. Finally, we evaluated the mean aerodynamics of the high-lift systems with Krueger flap and conventional slat as they were installed in an open-jet wind tunnel flow. The surface pressures predicted with the simulations agreed well with experimental results.

Computational↗
NASA NTRS1945

High-Speed Wind-Tunnel Tests of a 0.3-Scale Model of the P-47D Airplane

Presented in this report are the results of high-speed wind-tunnel tests of a 0.3-scale model of the P-47D airplane. Included are the aerodynamic characteristics of the model, pressure distributions on the wing and horizontal tail, elevator hinge moments, and tuft photographs showing the flow over the wing and horizontal tail. The results indicate that, at level-flight lift coefficients, the P-47D airplane will not exhibit a diving tendency below a Mach number of 0.78 at an altitude of 20,000 feet or below a Mach number of 0.72 at an altitude of 40,000 feet. The Mach number at which the lift, drag, and pitching-moment coefficients diverged rapidly from the low speed values is about C.15 higher than the critical Mach number. This is an unusually large margin.

William T Hamilton↗
NASA NTRS1991

Aerodynamic characteristics of scissor-wing geometries

A scissor-wing configuration, consisting of two independently sweeping-wing surfaces, is compared with an equivalent fixed-wing geometry baseline over a wide Mach number range. The scissor-wing configuration is shown to have a higher total lift-to-drag ratio than the baseline in the subsonic region primarily due to the slightly higher aspect ratio of the unswept scissor wing. In the transonic region, the scissor wing is shown to have a higher lift-to-drag ratio than the baseline for values of lift coefficient greater than 0.35. It is also shown that, through the use of wing decalage, the lift of the two independent scissor wings can be equalized. In the supersonic regime, the zero lift wave drag of the scissor-wing at maximum sweep is shown to be 50 and 28 percent less than the zero lift wave drag of the baseline at Mach numbers 1.5 and 3.0, respectively. In addition, a pivot-wing configuration is introduced and compared with the scissor wing. The pivot-wing configuration is shown to have a slightly higher total lift-to-drag ratio than the scissor wing in the supersonic region due to the decreased zero lift wave drag of the pivot-wing configuration.

Selberg, Bruce P.↗
NASA NTRS1928

Air Force Tests of Sperry Messenger Model with Six Sets of Wings

The purpose of this test was to compare six well-known airfoils, the R.A.F 15, U.S.A. 5, U.S.A. 27, U.S.A. 35-B, Clark Y, and Gottingen 387, fitted to the Sperry Messenger model, at full scale Reynolds number as obtained in the variable density wind tunnel of the National Advisory Committee for Aeronautics; and to determine the scale effect on the model equipped with all the details of the actual airplane. The results show a large decrease in minimum drag coefficient upon increasing the Reynolds number from about one-twentieth scale to full scale. Maximum lift coefficient was increased with increasing scale for all the airfoils except the Gottingen 387, for which it was slightly decreased. A comparison is made between the results of these tests and those obtained from tests made in this tunnel on airfoils alone. (author)

Shoemaker, James M↗
NASA NTRS2008

Wind Tunnel Testing of Powered Lift, All-Wing STOL Model

Short take-off and landing (STOL) systems can offer significant capabilities to warfighters and, for civil operators thriving on maximizing efficiencies they can improve airspace use while containing noise within airport environments. In order to provide data for next generation systems, a wind tunnel test of an all-wing cruise efficient, short take-off and landing (CE STOL) configuration was conducted in the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) 14- by 22-foot Subsonic Wind Tunnel. The test s purpose was to mature the aerodynamic aspects of an integrated powered lift system within an advanced mobility configuration capable of CE STOL. The full-span model made use of steady flap blowing and a lifting centerbody to achieve high lift coefficients. The test occurred during April through June of 2007 and included objectives for advancing the state-of-the-art of powered lift testing through gathering force and moment data, on-body pressure data, and off-body flow field measurements during automatically controlled blowing conditions. Data were obtained for variations in model configuration, angles of attack and sideslip, blowing coefficient, and height above ground. The database produced by this effort is being used to advance design techniques and computational tools for developing systems with integrated powered lift technologies.

Collins, Scott W.↗
NASA NTRS1987

Navier-Stokes computations of vortical flows over low aspect ratio wings

An upwind-biased finite-volume algorithm is applied to the low-speed flow over a low aspect ratio delta wing from zero to forty degrees angle of attack. The differencing is second-order accurate spatially, and a multigrid algorithm is used to promote convergence to the steady state. The results compare well with the detailed experiments of Hummel (1983) and others for a Re(L) of 0.95 x 10 to the 6th. The predicted maximum lift coefficient of 1.10 at thirty-five degrees angle of attack agrees closely with the measured maximum lift of 1.06 at thirty-three degrees. At forty degrees angle of attack, a bubble type of vortex breakdown is evident in the computations, extending from 0.6 of the root chord to just downstream of the trailing edge.

Thomas, J. L.↗
NASA NTRS1989

An Euler analysis of a High-Speed Civil Transport (HSCT) concept at Mach 3

A marching Euler solver, GEM3D, was used to predict the Mach 3 flow field for the wing and body of a High-Speed Civil Transport concept. The analysis focused on a typical cruise lift coefficient of 0.1 at alpha = 3 deg. The Euler solution indicated that embedded shocks formed on the upper surface of the inboard wing panel and at the leading-edge of the outboard wing panel, due to its supersonic leading edge condition. According to a simple static-pressure criterion, the embedded wing upper-surface shocks were sufficiently strong to separate a turbulent boundary layer. Comparison of aerodynamic coefficients from the Euler solver with those from linear theory shows that the linear theory estimates of lift and drag are optimistic, which would lead to optimistic estimates of cruise range.

Pittman, James L.↗
NASA NTRS1952

Method for calculating lift distributions for unswept wings with flaps or ailerons by use of nonlinear section lift data

A method is presented which allows the use of nonlinear section lift data in the calculation of the spanwise lift distribution of unswept wings with flaps or ailerons. This method is based upon lifting line theory and is an extension to the method described in NACA rep. 865. The mathematical treatment of the discontinuity in absolute angle of attack at the end of the flap or aileron involves the use of a correction factor which accounts for the inability of a limited trigonometric series to represent adequately the spanwise lift distribution. A treatment of the apparent discontinuity in maximum section lift coefficient is also described. Simplified computing forms containing detailed examples are given for both symmetrical and asymmetrical lift distributions. A few comparisons of calculated characteristics with those obtained experimentally are also presented.

Sivells, James C↗
NASA NTRS2003

Effects of Various Fillet Shapes on a 76/40 Double Delta Wing from Mach 0.18 to 0.7

The effects of linear, diamond, and parabolic fillets on a double delta wing were investigated in the NASA Langley 7 x 10 ft High Speed Tunnel from Mach 0.18 to 0.7 and angles of attack from 4 deg. to 42 deg. Force and moment, pneumatic pressures, pressure sensitive paint, and vapor screen flow visualization measurements were used to characterize the flow field and to determine longitudinal forces and moments. The fillets increased lift coefficient and reduced induced drag without significantly affecting pitching moment. Pressure sensitive paint showed the increase in lift is caused by an increase in suction and broadening of the vortex suction footprint. Vapor screen results showed the mixing and coalescing of the strake fillet and wing vortices causes the footprint to broaden.

Gonzalez, Hugo A.↗