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At least 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

Effects of a modified leading edge on noise and boundary-layer transition in a rod-wall sound shield at Mach 5

A version of a rod wall sound shield was tested in the Mach 5 pilot quiet tunnel over a range of unit Reynolds numbers from 10 to 35 million per meter. The model was modified by inclining the leading edge plates to produce an initial 2 deg expansion to ascertain the sensitivity of boundary layer transition to leading edge disturbances. Rod surface pitot pressures, mean free stream pitot pressures, and static pressures on the rods and plenum walls were measured. Hot-wire measurements were also made in the model and nozzle free stream at a unit Reynolds number of 15 million per meter. The surface pitot pressures indicated that transition was much farther forward than for the previous tests due to the leading edge modification and minor fabrication flaws in the model. Early boundary layer transition on the rods was confirmed by hot-wire measurements which showed much higher noise levels in the free stream shield flow when compared with results from previous tests. Mean pitot pressure surveys within the shielded region inside the model indicated that there was an overexpansion and recompression that would limit the streamwise length of undisturbed flow to about 13 cm along the centerline.

Creel, T. R., Jr.↗
NASA NTRS1980

Numerical lifting line theory applied to drooped leading-edge wings below and above stall

A numerical iterative solution to the classical Prandtl lifting-line theory, suitably modified for poststall behavior, is used to study the aerodynamic characteristics of straight rectangular finite wings with and without leading-edge droop. This study is prompted by the use of such leading-edge modifications to inhibit stall/spins in light general aviation aircraft. The results indicate that lifting-line solutions at high angle of attack can be obtained that agree with experimental data to within 20%, and much closer for many cases. Therefore, such solutions give reasonable preliminary engineering results for both drooped and undrooped wings in the poststall region. However, as predicted by von Karman, the lifting-line solutions are not unique when sectional negative lift slopes are encountered. In addition, the present numerical results always yield symmetrical lift distributions along the span, in contrast to the asymmetrical solutions observed by Schairer in the late 1930's. Finally, a series of parametric tests at low angle of attack indicate that the effect of drooped leading edges on aircraft cruise performance is minimal.

Anderson, J. D., Jr.↗
NASA NTRS1981

Effectiveness of leading-edge vortex flaps on 60 and 75 degree delta wings

A series of wind tunnel tests were run on 60 and 75 deg sweep delta wings to examine the effectiveness of leading-edge vortex flaps. Tests results showed that leading-edge vortex flaps are effective in giving large increases in lift-to-drag ratio and decreases in drag over a wide range of angle of attack. Tests on inverted flaps on the 60 deg delta wing showed substantial increases in lift and drag and may indicate a possibility of using inverted flaps on delta wings in the landing portion of flight. The 60 deg data were compared with that for a 75 deg sweep delta wing confirming that leading-edge vortex flap effectiveness is stronger as sweep is increased. Pitching moment effects due to vortex flaps use were also examined.

Marchman, J. F., III↗
NASA NTRS1956

Use of Truncated Flapped Airfoils for Impingement and Icing Tests of Full-Scale Leading-Edge Sections

In an effort to increase the operational range of existing small icing tunnels, the use of truncated airfoil sections has been suggested. With truncated airfoils, large-scale or even full-scale wing-icing-protection systems could be evaluated. Therefore, experimental studies were conducted in the NACA Lewis laboratory icing 'tunnel with an NACA 651-212 airfoil section to determine the effect of truncating the airfoil chord on velocity distribution and impingement characteristics. A 6-foot-chord airfoil was cut successively at the 50- and 30-percent-chord stations to produce the truncated airfoil sections, which were equipped with trailing-edge flaps that were used to alter the flow field about the truncated sections. The study was conducted at geometric angles of attack of 00 and 40, an airspeed of about 156 knots, and volume-median droplet sizes of 11.5 and 18.6 microns. A dye-tracer technique was used in the impingement studies. With the trailing-edge flap on the truncated airfoil deflected so that the local velocity distribution in the impingement region was substantially the same as that for the full-chord airfoil, the local impingement rates and the limits of impingement for the truncated and full-chord airfoils were the same. In general, truncating the airfoils with flaps undeflected resulted in a subs'tantially altered velocity distribution and local impingement rates compared with those of the full-chord airfoil. The use of flapped truncated airfoils may permit impingement and icing studies to be conducted with full-scale leading-edge sections, ranging in size from tip to root sections.

vonGlahn, Uwe H.↗
NASA NTRS1981

Subsonic balance and pressure investigation of a 60-deg delta wing with leading-edge devices (data report)

The drag reduction potential of leading edge devices on a 60 degree delta wing at high lift was examined. Geometric variations of fences, chordwise slots, pylon type vortex generators, leading edge vortex flaps, and sharp leading edge extensions were tested individually and in specific combinations to improve high-alpha drag performance with a minimum of low-alpha drag penalty. The force, moment, and surface static pressure data for angles of attack up to 23 degrees, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter are documented.

Rao, D. M.↗
NASA NTRS1981

Helicopter rotor trailing edge noise

An experimental and theoretical study was conducted to assess the importance of trailing edge noise as a helicopter main rotor broadband noise source. The noise mechanism was isolated by testing a rotor blade segment in an open jet acoustic wind tunnel at close to full scale Reynolds numbers. Boundary layer data and acoustic data were used to develop scaling laws and assess a first principles trailing edge noise theory. Conclusions from the isolated blade study were analytically transformed to the rotating frame coordinate system to develop a generalized rotor noise prediction. Trailing edge noise was found to contribute significantly to the total helicopter noise spectrum at high frequencies.

Schlinker, R. H.↗
NASA NTRS1981

Interlaminar stress singularities at a straight free edge in composite laminates

A quasi-three-dimensional finite-element analysis was used to analyze the edge-stress problem in four-ply, composite laminates. The seven laminates that were considered belong to the laminate family where the outer ply angle is between 0 and 90 deg. Systematic convergence studies were made to explore the existence of stress singularities near the free edge. The present analysis appears to confirm the existence of stress singularities at the intersection of the interface and the free edge. The power of the stress singularity was the same for all seven laminates considered.

Raju, I. S.↗
NASA NTRS1982

Sharp edges of planetary rings

The ring systems of Saturn and Uranus exhibit several sharp edges, across which the optical depth drops from order unity to essentially zero. At least two and perhaps all of these features are associated with the location of orbital resonances between a satellite and the ring particles. It is remarkable that the optical depth varies on a distance scale which is much finer than that over which angular momentum can be transferred between a satellite and the ring material. The important features of this phenomenon are: (1) a perturbed band of width Delta a/a of about (satellite mass/planetary mass) to the 1/2 power adjacent to the edge within which the angular momentum transfer occurs, (2) streamlines perturbed such that the angular momentum luminosity decreases smoothly across the band to zero at the edge even though the optical depth remains constant, and (3) dynamical equilibrium requires a relation between the random velocity, the rate of deformation and the optical depth.

Borderies, N.↗
NASA NTRS1973

Wind-tunnel development of underwing leading-edge vortex generators on a NASA supercritical-wing research airplane configuration

A program for the wind tunnel development of underwing leading-edge vortex generators on an NASA supercritical-wing research airplane model was conducted in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.25 to 0.99. The effects on the longitudinal aerodynamic characteristics of vortex-generator wing semispan location, distance from wing leading edge, leading-edge sweep angle, toe-in angle, and span are presented. The effects of the vortex generators on the lateral-directional aerodynamic characteristics of the model for a sideslip angle of 5 deg are included.

Bartlett, D. W.↗
NASA NTRS1982

Study of supersonic wings employing the attainable leading-edge thrust concept

A theoretical study was made of supersonic wing geometries at Mach 1.8, using the attainable leading-edge thrust concept. The attainable thrust method offers a powerful means to improve overall aerodynamic efficiency by identifying wing leading-edge geometries that promote attached flow and by defining a local angle-of-attack range over which attached flow may be obtained. The concept applies to flat and to cambered wings, which leads to the consideration of drooped-wing leading edges for attached flow at high lift coefficients.

Middleton, W. D.↗
NASA NTRS1983

Mixed-mode strain-energy-release rate effects on edge delamination of composites

Unnotched graphite/epoxy laminates, designed to delaminate at the edges under static and cyclic tensile loads, were tested and analyzed. The specimen stacking sequences were chosen so that the total strain-energy-release rate, G, for edge delamination was identical for all three layups. However, each layup had different percentages of crack-opening and shear-mode strain-energy-release rates, G sub 1 and G sub 2, respectively. Results with composites made from T300 graphite fibers and 5208 epoxy, a brittle resin, indicated that only G sub 1 contributed to delamination onset under static loading. However, results with composites made from C6000 fibers and H205 epoxy, a tougher resin, indicated that the total F governed the onset of edge delaminations under cyclic loads. In addition, for both materials, the threshold level of G for delamination onset in fatigue was significantly less than the critical G sub c measured in static tests. Futhermore, although the C6000/H205 material had a much higher static G sub c than T300/5208, its fatigue resistance was only slightly better. A series of mixed-mode tests, like the ones in this study, may be needed to evaluate toughened-resin composites developed for highly strained composite structures subjected to cyclic loads.

Obrien, T. K.↗
NASA NTRS1983

Experimental study of wing leading-edge devices for improved maneuver performance of a supercritical maneuvering fighter configuration

Wind tunnel tests were conducted to examine the use of wing leading-edge devices for improved subsonic and transonic maneuver performance. These devices were tested on a fighter configuration which utilized supercritical-wing technology. The configuration had a leading-edge sweep of 45 deg and an aspect ratio of 3.28. The tests were conducted at Mach numbers of 0.60 and 0.85 with angles of attack from -0.5 deg to 22 deg. At both Mach numbers, sharp leading-edge flaps produced vortices which greatly altered the flow pattern on the wing and resulted in substantial reductions in drag at high lift. Underwing or pylon-type vortex generators also reduced drag at high lift. The vortex generators worked better at a Mach number of 0.60. The vortex generators gave the best overall results with zero toe-in angle and when mounted on either the outboard part of the wing or at both an outboard location and halfway out the semispan. Both the flaps and the vortex generators had a minor effect on the pitching moment. Fluorescent minitufts were found to be useful for flow visualization at transonic maneuver conditions.

Mann, M. J.↗
NASA NTRS1983

Wind-tunnel investigation of leading-edge thrust on arrow wings in supersonic flow

Six wing models were tested in the Langley Unitary Plan Wind Tunnel to identify and study leading-edge thrust at supersonic speeds. The tests were conducted at Mach numbers of 1.6, 1.8, 2.0, and 2.16, at a stagnation temperature of 125.0 F, and at Reynolds numbers per foot of 2.0 x 10 to the 6th power and 5.0 x 10 the 6th power. Test results showed that significant benefits from leading-edge thrust and nonlinear thickness effects can be obtained with very little airfoil bluntness, that these benefits were lost when the airfoil was severely blunted, and that such benefits seem to be found on wings with supersonic as well as subsonic leading edges.

Mack, R. J.↗
NASA NTRS1982

On the relationship of the plasmapause to the equatorward boundary of the auroral oval and to the inner edge of the plasma sheet

ISEE 1 observations of the plasmapause are compared with simultaneous observations of the electron plasma sheet and also the auroral oval observed in DMSP photographs. Only a limited amount of appropriate data was available for the comparisons: the plasmapause/plasma sheet inner edge comparisons were restricted to the early and late morning sectors, while there were two satisfactory comparisons of the plasmapause and the equatorward boundary of the auroral oval in the evening sector. However, these examples indicate that the plasmapause location often coincides to within a change in L of about 0.1-0.2 with both the plasma sheet inner boundary and the field line threading the equatorward boundary of the auroral oval. This co-location of the plasmapause and the plasma sheet inner edge may be due to shielding of the magnetospheric convection electric field by an Alfven layer located at the inner edge of the plasma sheet as discussed by Jaggi and Wolf (1973) and others.

Horwitz, J. L.↗
NASA NTRS1982

Edge softening of the Shuttle TPS strain isolation pad

Tensile tests and an analytical investigation were performed to characterize the edge softening behavior of the strain isolation pad (SIP) between the Orbiter skin and thermal protection system. The tensile tests were carried out with varying sizes of disk-shaped specimens bonded between aluminum disks. The specimens strength and stiffness were determined on the basis of specimen size, and an analytical model of the microstructural stress-strain characteristics was developed. Strength and stiffness were found to decrease near the free edges because through-the-thickness fibers located there were not anchored. No size dependence at maximum load was observed in specimens between 0.75-4.0 in. thick. In-plane and out-of-plane coupling in deformation was detected. The model gave accurate predictions of the tensile behavior of the SIP as a function of distance to a free edge.

Ransone, P. O.↗
NASA NTRS1983

Edge effects in composites by moire interferometry

The very high sensitivity of moire interferometry has permitted the present edge effect experiments to be conducted at a low average stress and strain level, assuring linear and elastic behavior in the composite material samples tested. Sensitivity corresponding to 2450 line/mm moire was achieved with a 0.408 micron/fringe. Simultaneous observations of the specimen face and edge displacement fields showed good fringe definition despite the 1-mm thickness of the specimens and the high gradients, and it is noted that the use of a carrier pattern and optical filtering was effective in even these conditions. Edge effects and dramatic displacement gradients were confirmed in angle-ply composite laminates.

Czarnek, R.↗
NASA NTRS1983

An investigation of wing leading-edge vortices at supersonic speeds

Studies at subsonic and transonic speeds of the fundamental vortex behavior on the leeward surface of wings have led to the design of several unique and novel leading-edge devices commonly referred to as 'vortex flaps'. The present investigation has the objective to provide some fundamental vortex-flow results obtained at supersonic speeds. Experimental studies were performed in which pressure data and several types of flow visualization data were obtained on the leeward surface of a series of flat delta-wing models to identify the various flow mechanisms which can occur and to determine the effect of leading-edge sweep, Mach number, and angle of attack on the vortex strength and location. The reported investigation forms part of a study which is to explore the use of wing leading-edge vortex technology as a supersonic wing-design tool. The obtained results indicate that the procedure of distributing the vortex force as a pressure variation about a vortex action line is a promising concept.

Miller, D. S.↗
NASA NTRS1983

A flight test of laminar flow control leading-edge systems

NASA's program for development of a laminar flow technology base for application to commercial transports has made significant progress since its inception in 1976. Current efforts are focused on development of practical reliable systems for the leading-edge region where the most difficult problems in applying laminar flow exist. Practical solutions to these problems will remove many concerns about the ultimate practicality of laminar flow. To address these issues, two contractors performed studies, conducted development tests, and designed and fabricated fully functional leading-edge test articles for installation on the NASA JetStar aircraft. Systems evaluation and performance testing will be conducted to thoroughly evaluate all system capabilities and characteristics. A simulated airline service flight test program will be performed to obtain the operational sensitivity, maintenance, and reliability data needed to establish that practical solutions exist for the difficult leading-edge area of a future commercial transport employing laminar flow control.

Fischer, M. C.↗