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At least 1,243 records · Page 69

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NASA NTRS1972

Sideslip of wing-body combinations

A small-disturbance theory is developed for predicting the aerodynamics of an airplane in sideslip. Second-order terms involving the interaction between sideslip angle and angle of attack, sideslip angle and wing camber, etc., are retained. It is found that the second-order terms can produce the dominant sideslip effects when the dihedral of the lifting surfaces is small. Numerical implementation of the theory requires a solution procedure capable of producing accurate velocity gradients in the first-order solution.

Rubbert, P. E.↗

NASA NTRS1976

Investigation of two bifurcated-duct inlet systems from Mach 0 to 2.0 over a wide range of angles of attack

A 15.354 percent/scale lightweight fighter type inlet/forebody was tested over a Mach number range of 0 to 2.0. Model configurations consisted of side mounted normal shock and fixed overhead ramp type inlets. Each configuration consisted of two inlets ducted (bifurcated) to supply a single engine face. The normal shock inlet variables included a boundary layer splitter bleed system, alternate boundary layer splitter plates, alternate upper and lower cowl lip shapes, and a blow-in-door (auxiliary inlet) in one lower lip. The only variable of the fixed overhead ramp inlet was the boundary layer bleed flow. Reynolds numbers ranged from 7.6 x 1 million to 19.5 x 1 million/m. Angle of attack ranged from -10 to 35 deg and angle of sideslip from -8 to 8 deg. Test measurements included engine face total pressure recovery, steady state distortion, dynamic distortion, and surface static pressures on the forebody and inlet surfaces.

Latham, E. A.↗

NASA NTRS1977

Experimental evaluation of nacelle-airframe interference forces and pressures at Mach numbers of 0.9 to 1.4

Detailed interference force-and-pressure data were obtained on a representative supersonic transport wing-body-nacelle combination at Mach numbers of 0.9 to 1.4. The basic model consisted of a delta wing-body aerodynamic model with a length of 158.0 cm (62.2 in.) and a wingspan of 103.6 cm (40.8 in.) and four independently supported nacelles positioned beneath the model. The experimental program was conducted in the Ames 11- by 11-Foot Wind Tunnel at a constant unit Reynolds number. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass-flow ratio. Under the most favorable conditions, the net interference drag was equal to 50 percent the drag of four isolated nacelles at M = 1.4, 75 percent at M = 1.15, and 144 percent at M = 0.90. The overall interference effects were found to be rather constant over the operating angle-of-attack range of the configuration. The effects of mass-flow ratio on the interference pressure distributions were limited to the lip region of the nacelle and the local wing surface in the immediate vicinity of the nacelle lip. The net change in the measured interference forces resulting from variations in the nacelle mass-flow ratio were found to be quite small.

Bencze, D. P.↗

NASA NTRS1981

Large-scale wind tunnel tests of a sting-supported V/STOL fighter model at high angles of attack

A new sting model support has been developed for the NASA/Ames 40- by 80-Foot Wind Tunnel. This addition to the facility permits testing of relatively large models to large angles of attack or angles of yaw depending on model orientation. An initial test on the sting is described. This test used a 0.4-scale powered V/STOL model designed for testing at angles of attack to 90 deg and greater. A method for correcting wake blockage was developed and applied to the force and moment data. Samples of this data and results of surface-pressure measurements are presented.

Stoll, F.↗

NASA NTRS1982

An experimental and theoretical investigation of thick wings at various sweep angles in and out of ground effect

The effects of sweep and aspect ratio on the longitudinal aerodynamics of a wing in and out of ground effect are analyzed. Experimental data were obtained in the Langley 4 by 7 Meter Tunnel for a wing with various sweep angles, aspect ratios, and flap deflections both in and out of ground effect. Theoretical predictions of the out of ground effect lift coefficients and flap effectiveness and the in ground effect lift coefficients are compared with the experimental results. As expected, the lift curve slope and flap effectiveness are reduced when the aspect ratio is reduced or the sweep angle is increased both in and out of ground effect. In ground effect, the lift and flap effectiveness are increased above a wing height to span ratio of 0.15. However, with flap deflections less than or equal to 10 deg and an angle of attack near 0 deg lift is markedly decreased at very low heights above the ground plane. This trend is not predicted by planar theoretical models but is predicted by a surface panel method where thickness effects are included.

Paulson, J. W., Jr.↗

NASA NTRS1983

Transonic pressure distributions on a rectangular supercritical wing oscillating in pitch

Steady and unsteady aerodynamic data were measured on a rectangular wing with a 12 percent thick supercritical airfoil mounted in the NASA Langley Transonic Dynamics Tunnel. The wing was oscillated in pitch to generate the unsteady aerodynamic data. The purpose of the wind-tunnel test was to measure data for use in the development and assessment of transonic analytical codes. The effects on the wing pressure distributions of Mach number, mean angle of attack, and oscillation frequency and amplitude were measured. Results from the newly-developed XTRAN3S program (a nonlinear transonic small disturbance code) and from the RHOIV program (a linear lifting surface kernel function code) were compared to measured data for a Mach number of 0.7 and for oscillation frequencies ranging from 0 to 20 Hz. The XTRAN3S steady and unsteady results agreed fairly well with the measured results. The RHOIV unsteady-result agreement was fair but, of course, did not predict shock effects.

Ricketts, R. H.↗

NASA NTRS1985

Reflection plane tests of a wind turbine blade tip section with ailerons

Tests were conducted in the NASA Langley 30 by 60 foot Wind Tunnel on a full scale 7.31 m (24 ft) long tip section of a wind turbine rotor blade. The blade tip section was built with ailerons on the trailing edge. The ailerons, which spanned a length of 6.1 m (20 ft), were designed so that two types could be evaluated: the plain and the balanced. The ailerons were hinged on the suction surface at the 0.62 X chord station behind the leading edge. The purpose of the tests was to measure the aerodynamic characteristics of the blade section for: an angle of attack range from 0 deg to 90 deg aileron deflections from 0 deg to -90 deg, and Reynolds numbers of 0.79 and 1.5 x 10 to the 6th power. These data were then used to determine which aileron configuration had the most desirable rotor control and aerodynamic braking characteristics. Tests were also run to determine the effects of vortex generators, leading edge roughness, and the gaps between the aileron sections on the lift, drag, and chordwise force coefficients of the blade tip section.

Savino, J. M.↗

NASA NTRS1987

Numerical analysis of a NACA0012 airfoil with leading edge ice accretions

Analysis of a NACA0012 airfoil with leading edge ice has been performed using a Navier-Stokes code coupled with a grid generation code. The computed results were compared to experimental information obtained for an airfoil with a well defined artificial ice shape. The computations were performed at angles of attack ranging from zero to ten degrees. This range is sufficient to show the development of the separation bubble aft of the ice shape on both the upper and lower surfaces. Velocity profile plots in the separation bubble are examined in order to determine if recirculation patterns are predicted properly and if separation and reattachment points are found within the resolution of the experimental information. Also, the massive separation near the point of stall is examined in order to more accurately evaluate the lift coefficient curve in that region. Lift, drag, and moment coefficients are computed and compared to experiment.

Potapczuk, Mark G.↗

NASA NTRS1989

Flow visualization studies of the Mach number effects on the dynamic stall of an oscillating airfoil

Compressibility effects on the dynamic stall of a NACA 0012 airfoil undergoing sinusoidal oscillatory motion were studied using a stroboscopic schlieren system. Schlieren pictures and some quantitative data derived from them are presented and show the influence of free-stream Mach number and reduced frequency on the dynamic-stall vortex. This study shows that a dynamic stall vortex always forms near the leading edge and convects on the airfoil upper surface at approximately 0.3 times the free stream velocity for all cases studied. The results also demonstrate that initiation of the dynamic stall vortex is delayed to higher angles of attack with increased reduced frequency, but that dynamic stall occurs at lower angles of incidence with increasing Mach numbers.

Chandrasekhara, M. S.↗

NASA NTRS1989

Numerical study of the effect of tangential leading edge blowing on delta wing vortical flow

A numerical simulation of tangential blowing along the leading edge of a delta wing is analyzed as a means of controlling the position and strength of the leading-edge vortices. The computation is done by numerical solutions of the three-dimensional thin-layer Navier-Stokes equations. Numerical results are shown to compare favorably with experimental measurements. It is found that the use of tangential leading-edge blowing at low to moderate angles of attack tends to reduce the pressure peaks associated with leading-edge vortices and to increase the suction peak around the leading edge, such that the integrated value of the surface pressure remains about the same.

Yeh, David T.↗

NASA NTRS1993

Viscous/Inviscid Interaction On An Oscillating Airfoil

Viscous/inviscid-interaction procedure developed for computation of two-dimensional flow of air about airfoil, angle of attack of which oscillates at high amplitude. Purpose of procedure to study flow effects in dynamic stall. Bulk of flow treated as inviscid, while effects of viscosity treated only on surface of airfoil and in boundary layer. Inviscid part of flow simulated via time-marching, unsteady-subsonic-flow panel method incorporating appropriate distributions of unsteady surface and wake singularities.

Carr, L. W.↗

NASA NTRS1921

The Pressure Distribution over the Horizontal Tail Surfaces of an Airplane II

This investigation was undertaken to determine whether the results obtained upon model tail surfaces can be used to accurately predict loads upon the full-sized tail; and also to find the distribution of load when large elevator angles are used, as the loads from such angles can not be obtained readily in free flight. The method consisted in using a metal horizontal tail surface inside of which small air passages, connecting with a series of holes in the surface, led the pressure off from the tail in rubber tubes. In this way the pressure at each of these holes was measured by a manometer at several angles of attack and several to the loading under similar conditions in the full-sized airplane and the manner of distribution is quite similar in the two cases when there is no slip stream.

Norton, F H↗

NASA NTRS1995

Flame-Resistant Composite Materials For Structural Members

Matrix-fiber composite materials developed for structural members occasionally exposed to hot, corrosive gases. Integral ceramic fabric surface layer essential for resistance to flames and chemicals. Endures high temperature, impedes flame from penetrating to interior, inhibits diffusion of oxygen to interior where it degrades matrix resin, resists attack by chemicals, helps resist erosion, and provides additional strength. In original intended application, composite members replace steel structural members of rocket-launching structures that deteriorate under combined influences of atmosphere, spilled propellants, and rocket exhaust. Composites also attractive for other applications in which corrosion- and fire-resistant structural members needed.

Spears, Richard K.↗

NASA NTRS1996

The modelling of symmetric airfoil vortex generators

An experimental study is conducted to determine the dependence of vortex generator geometry and impinging flow conditions on shed vortex circulation and crossplane peak vorticity for one type of vortex generator. The vortex generator is a symmetric airfoil having a NACA 0012 cross-sectional profile. The geometry and flow parameters varied include angle-of-attack alfa, chordlength c, span h, and Mach number M. The vortex generators are mounted either in isolation or in a symmetric counter-rotating array configuration on the inside surface of a straight pipe. The turbulent boundary layer thickness to pipe radius ratio is delta/R = 0. 17. Circulation and peak vorticity data are derived from crossplane velocity measurements conducted at or about 1 chord downstream of the vortex generator trailing edge. Shed vortex circulation is observed to be proportional to M, alfa, and h/delta. With these parameters held constant, circulation is observed to fall off in monotonic fashion with increasing airfoil aspect ratio AR. Shed vortex peak vorticity is also observed to be proportional to M, alfa, and h/delta. Unlike circulation, however, peak vorticity is observed to increase with increasing aspect ratio, reaching a peak value at AR approx. 2.0 before falling off.

Reichert, B. A.↗

NASA NTRS1999

Turbulence Model Comparisons for a High-Speed Aircraft

Four turbulence models are described and evaluated for transonic flows over the High-Speed Research/industry baseline configuration known as Reference H by using the thin-layer, upwind, Navier-Stokes solver known as CFL3D. The turbulence models studied are the equilibrium model of Baldwin-Lomax (B-L) with the Degani-Schiff (D-S) modifications, the one-equation Baldwin-Barth (B-B) model, the one-equation Spalart-Allmaras (S-A) model, and Menter's two-equation Shear Stress Transport (SST) model. The flow conditions, which correspond to tests performed in the National Transonic Facility (NTF) at Langley Research Center, are a Mach number of 0.90 and a Reynolds number of 30 x 10 (exp. 6) based on mean aerodynamic chord for angles of attack of 1 deg., 5 deg., and 10 deg. The effects of grid topology and the representation of the actual wind tunnel model geometry are also investigated. Computed forces and surface pressures compare reasonably well with the experimental data for all four turbulence models.

Rivers, Melissa B.↗

NASA NTRS2000

Additional Testing of the DHC-6 Twin Otter Tailplane Iced Airfoil Section in the Ohio State University 7x10 Low Speed Wind Tunnel

The need for fundamental research in Ice Contaminated Tailplane Stall (ICTS) was established through three international conferences sponsored by the FAA. A joint NASA/FAA Tailplane Icing Program was formed in 1994 with the Ohio State University playing a critical role for wind tunnel and analytical research. Two entries of a full-scale 2-dimensional tailplane airfoil model of a DHC-6 Twin Otter were made in The Ohio State University 7x10 ft wind tunnel. This report describes the second test entry that examined additional ice shapes and roughness, as well as airfoil section differences. The addition data obtained in this test fortified the original database of aerodynamic coefficients that permit a detailed analysis of flight test results with an OSU-developed analytical program. The testing encompassed a full range of angles of attack and elevator deflections at flight Reynolds number conditions. Aerodynamic coefficients, C(L), C(M), and C(He), were obtained by integrating static pressure coefficient, C(P), values obtained from surface taps. Comparisons of clean and iced airfoil results show a significant decrease in the tailplane aeroperformance (decreased C(Lmax), decreased stall angle, increased C(He)) for all ice shapes with the grit having the lease affect and the LEWICE shape having the greatest affect. All results were consistent with observed tailplane stall phenomena and constitute an effective set of data for comprehensive analysis of ICTS.

Gregorek, Gerald↗

NASA NTRS2002

Boundary Layer Transition on Slender Cones in Conventional and Low Disturbance Mach 6 Wind Tunnels

An experimental investigation was conducted on a 5-degree half-angle cone and a 5-degree half-angle flared cone in a conventional Mach 6 wind tunnel to examine the effects of facility noise on boundary layer transition. The influence of tunnel noise was inferred by comparing transition onset locations determined from the present test to that previously obtained in a Mach 6 low disturbance quiet tunnel. Together, the two sets of experiments are believed to represent the first direct comparison of transition onset between a conventional and a low disturbance wind tunnel using a common test model and transition detection technique. In the present conventional hypersonic tunnel experiment, separate measurements of heat transfer and adiabatic wall temperatures were obtained on the conical models at small angles of attack over a range of Reynolds numbers, which resulted in laminar, transitional, and turbulent flow. Smooth model turbulent heating distributions are compared to that obtained with transition forced via discrete surface roughness. The model nosetip radius was varied to examine the effects of bluntness on transition onset. Despite wall to total temperature differences between the transient heating measurements and the adiabatic wall temperature measurement, the two methods for determining sharp cone transition onset generally yielded equivalent locations. In the 'noisy' mode of the hypersonic low disturbance tunnel, transition onset occurred earlier than that measured in the conventional hypersonic tunnel, suggesting higher levels of freestream acoustic radiation relative to the conventional tunnel. At comparable freestream conditions, the transition onset Reynolds number under low disturbance conditions was a factor of 1.3 greater than that measured on flared cone in the LaRC conventional hypersonic tunnel and a factor of 1.6 greater that the flared cone run in the low disturbance tunnel run 'noisy'. Navier-Stokes mean flow computations and linear stability analysis were conducted to assess the experimental results and have indicated N factors associated with sharp flared cone transition onset to be approximately a factor of 2 lower than that inferred from the corresponding low disturbance tunnel measurements.

Horvath, Thomas J.↗

NASA NTRS2003

Automated CFD Database Generation for a 2nd Generation Glide-Back-Booster

A new software tool, AeroDB, is used to compute thousands of Euler and Navier-Stokes solutions for a 2nd generation glide-back booster in one week. The solution process exploits a common job-submission grid environment using 13 computers located at 4 different geographical sites. Process automation and web-based access to the database greatly reduces the user workload, removing much of the tedium and tendency for user input errors. The database consists of forces, moments, and solution files obtained by varying the Mach number, angle of attack, and sideslip angle. The forces and moments compare well with experimental data. Stability derivatives are also computed using a monotone cubic spline procedure. Flow visualization and three-dimensional surface plots are used to interpret and characterize the nature of computed flow fields.

Chaderjian, Neal M.↗