Showing papers on "Leading edge published in 1982"

Experimental Studies of Separation on a Two-Dimensional Airfoil at Low Reynolds Numbers

Abstract: The laminar separation, transition, and turbulent reattachment near the leading edge of a two-dimensional NACA 663 -018 airfoil were investigated using a low-speed, smoke visualization wind tunnel. Lift and drag force measurements were made using an external strain gage balance for a chord Reynolds number range of 40,GOO400,000. An extensive flow visualization study was performed and correlated with the force measurements. Experiments were also conducted with distributed surface roughness at the leading edge and external acoustic excitation to influence the development of the airfoil boundary layer. This study delineates the effects of angle of attack and chord Reynolds number on the separation characteristics and airfoil performance. Nomenclature c = model chord cd = section profile drag coefficient (uncorrected) cf = section lift coefficient (uncorrected) Cp = pressure coefficient / = acoustic frequency, Hz R = reattachment location Rc = Reynolds number based on chord length, U^ civ S = separation location T = location of approximate end of transition £/«, = freestream velocity x/c = nondimensional distance along chord a = angle of attack v - kinematic viscosity

Blunt fin-induced shock wave/turbulent boundary-layer interaction

Abstract: This paper presents results from an experimental study of blunt fin-induced shock wave/turbulent boundarylayer interaction Semi-infinite fin models with hemicylindrical, unswept leading edges were tested in Mach 3, high Reynolds number, turbulent boundary layers All tests were made under approximately adiabatic wall conditions The program had two fundamental objectives The first was to examine the spanwise development of the disturbed flowfield and to determine its dependence on the configuration geometry and incoming flow conditions To achieve this, streamwise surface pressure distributions were measured in the region extending from the centerline to 110 fin diameters outboard The second objective was to determine the vertical extent of the interaction on the fin This was carried out using a fin model whose leading edge and side face were instrumented with pressure taps The results show that, on the test surface near the fin and on the fin itself, the leading-edge diameter plays a dominant role in determining the interaction's scale and characteristics

Vortex–leading-edge interaction

Abstract: Visualization of successive vortices impinging upon the leading edge of a wedge reveals patterns of deformation of each incident vortex; for certain offsets of the edge with respect to the incident vortex there is pronounced vortex shedding from the leading edge, whereby the shed vortex has a vorticity orientation opposite to that of the incident vortex. Simultaneous consideration of this visualization interaction and the force induced on the wedge gives the relation between the nature of the interaction mechanism and the relative magnitude and phase of the force exerted on the wedge. The amplitude of the induced force is found to be a strong function of the transverse offset of the leading edge with respect to the incident vortex and the degree of vorticity shedding from the leading edge. Application of Stuart's vortex model to the incident vortices provides a means for approximating the phase and relative amplitude of the induced force as a function of the transverse offset of the leading edge.

Water-Tunnel Studies of Leading-Edge Vortices

Abstract: Flow visualization studies have been made at Northrop in a hydrodynamic facility of leading-edge vortex flows. Vortex core trajectory and stability characteristics have been obtained on wing planforms suitable for subsonic-transonic and supersonic cruise fighter designs. The applicability of results obtained at low Reynolds number in water to higher Reynolds number vortex flow phenomena in air is addressed. Comparisons of watertunnel vortex positions and burst locations are made with flow visualization results obtained in air. Where appropriate, correlation of water-tunnel vortex flow behavior is made with trends observed in subsonic windtunnel data.

Oscillations of an unstable mixing layer impinging upon an edge

Abstract: The central features of lihear and nonlinear disturbance growth in the unstable shear layer, mechanisms of impingement of the resultant vortices on the edge, induced force on the wedge, and upstream influence in the form of induced velocity fluctuations at separation are examined by simultaneous visualization, velocity, and force-measurement techniques.The nature of the vortex–wedge interaction, and the associated force on the wedge, are directly related to the induced velocity at the upstream separation edge, thereby providing the essential ‘feedback’ for the self-sustained oscillation. Velocity fluctuations at the upper and lower sides of the separation edge tend to be π out of phase, a condition that is maintained along the outer boundaries of the downstream shear layer. Moreover, the phase between velocity fluctuations at separation and impingement satisfies the relation 2nπ, where n is an integer.The shear layer downstream of the separation edge initially forms an asymmetric wake, which evolves into large-scale vortices, all of which have a circulation appropriate to the high-speed side. The disturbance amplification associated with the high-speed side dominates from the separation edge onwards, precluding development of instabilities associated with the low-speed side.Regardless of the initial amplitude of the disturbance induced at the separation edge, the same saturation amplitude is attained in the downstream (nonlinear) region of the shear layer, underscoring the fact that variations in force amplitude at the wedge are dominated by the type of vortex–edge interaction mechanism. The sensitivity of this interaction to small offsets between the vortex centre and the leading edge entails that jumps in frequency of oscillation are also associated with jumps in the force amplitude.

Variable camber leading edge assembly for an airfoil

Abstract: A variable camber leading edge device having a movable nose section and an upper flexible panel extending rearwardly from the nose section. To move the nose section between its upper cruise position and a high lift downwardly deflected position, there is attached to the nose section a generally arcuate cam track, having its forward and rear portions curving upwardly. The cam track is constrained to move in a generally arcuate path by means of two spaced pair of rollers, and it is driven by a pinion gear engaging an upwardly facing middle portion of the track.

Volumetric drop detector

Abstract: A device for measuring the volume of flow in a liquid conveying apparatus wherein the length and velocity of a liquid column are measured. Drops of the liquid are formed and directed into a length of measurement tubing having a known internal diameter. As the liquid column is passed through the measurement tubing its length is determined by two spaced apart liquid sensors. The actual length is determined by calculating the time elapsed for the leading edge of the liquid column to interrupt the two liquid sensors. The time for the trailing edge to pass the first sensor is proportioned to the length of the column and the quotient of the two values will give the actual length. Accordingly, accurate volume determinations can be made irrespective of velocity change such as due to viscosity, surface tension or back pressure. The device of this invention is especially adapted to be used in conjunction with an intravenous administration set.

Concerning Dynamic Stall

Abstract: The unsteady breakdown or stall of streamlined flow near the rounded leading edge of an aerofoil, as the small angle of attack is raised above its critical stall value, is studied theoretically for large Reynolds number motion. The unsteady developments take place first over a relatively slow time scale but then the corresponding solution breaks down with a singularity, forcing a switch to a faster and more nonlinear process. The latter involves a very pronounced local bulge appearing in the flow displacement, accompanied by reversed flow at the aerofoil surface, and comparisons with experimental observations of dynamic stall are noted.

Asymptotic theory of short separation regions on the leading edge of a slender airfoil

Abstract: The two-dimensional flow of a viscous incompressible fluid near the leading edge of a slender airfoil is considered. An asymptotic theory of this flow is constructed on the basis of an analysis of the Navier—Stokes equations at large Reynolds numbers by means of matched asymptotic expansions. A central feature of the theory is the region of interaction of the boundary layer and the exterior inviscid flow; such a region appears on the surface of the airfoil in a definite range of angles of attack. The boundary-value problem for this region is reduced to an integrodifferential equation for the distribution of the friction. This equation has been solved numerically. As a result, closed separation regions are constructed, and the angle of attack at which separation occurs is found.

Monocoque type rotor blade

Abstract: The present rotor blade is constructed as a so-called monocoque or shell type blade for use in large fans and windmill rotors. For this purpose the blade is made of fiber reinforced synthetic material with a differnt orientation of the fibers in different sections of the blade. The transition between the blade proper and the connecting end of the blade is formed by a blade root section which has a connecting end of circular or elliptical cross-section and a wing facing end corresponding in cross-section to the blade cross-section. The blade root section has three zones. In the main leading edge zone including the leading edge and adjacent sides of the blade root section the reinforcing fibers extend in parallel to one another. The fibers also extend unidirectionally in a rear zone forming the trailing edge of the blade root section and in this trailing edge zone the fibers also extend in parallel to one another but at an angle relative to the fibers in the leading edge zone. An intermediate zone is located between the leading and trailing edge zones. The fibers in the intermediate zone are arranged in a cross-over relationship relative to one another. The connecting end is formed only by unidirectionally extending fibers.

Cooled turbine blade

Abstract: The blade has an aerofoil body 10 having a leading edge surface 11 which is cooled by air passing through a helical first passage 12 having first portions passing close to said leading edge surface and alternating with second portions passing through a more nearly central part of the blade section remote from said leading edge. A spanwise but straight second passage 15 extends through the blade in a position within the helical passage and closer to the second than the first portions thereof. Heat abstracted from the leading edge by air flow in said first portions is transferred by the flow to the second portions and from there through the blade material to the flow in the straight passage.

Wing leading edge slat

Abstract: A chordwise extensible wing leading edge system incorporating a rigid contoured panel member (22) which is exposed during forward extension movement of the leading edge slat (21) for closing in the proximity area between said slat and the relatively stationary leading edge portion of the wing (20), as a function of the forward movement of the slat (21). A further embodiment utilizes a second panel member (32) so that both the upper and the lower surface contour of the wing airfoil envelope are aerodynamically continuous to complete the contour of the intervening space between the extended position of the slat and the stationary portion of the wing; whereby, this combination results in approaching the merits of wing cambering devices capable of operation at relatively high airspeeds as opposed to landing or take-off airspeeds.

Effect of thickness on airfoil surface noise

Abstract: Noise emission from very small chord and very large chord airfoils was measured with eleven 0.63 cm microphones placed along a horizontal semicircle (4.57 m radius) that was centered at the leading edge of the test airfoil. The noise signals were analyzed by an automated spectrum analyzer which yielded 1/3-octave band sound pressure level spectra for each microphone, and the data were corrected to remove the effects of atmospheric attenuation and jet noise. It is found that the effect of thickness is large and must be accounted for in any fundamental airfoil noise theory that attempts to describe the noise emitted from real airfoils. Incident mean velocity gradients and compressibility must also be taken into account. The effect of thickness increases with frequency, with thick airfoils being quieter than thin ones.

Axial flow fan

Abstract: An axial flow fan has a plurality of impeller blades spaced around a hub assembly. The leading edge of each blade overlaps completely the trailing edge of the preceding blade, and the angular spacing between the radial center lines of the blades is unequal, preferably varying in a sinusoidal pattern, so that tonal noise of the fan is effectively attenuated.

Three-Dimensional Flow Field in the Tip Region of a Compressor Rotor Passage—Part II: Turbulence Properties

Abstract: The turbulence properties in the annulus wall region of an axial flow compressor rotor was measured using a triaxial, hot-wire probe rotating with the rotor. The flow was surveyed across the entire passage at five axial locations (leading edge, 1/4 chord, 1/2 chord, 3/4 chord, and the trailing edge location) and at six radial locations in a low-speed compressor rotor. The data derived include all three components of turbulence intensity and three components of turbulence stress. A comprehensive interpretation of the data with emphasis on features related to rotation, leakage flow, annulus wall boundary layer, and blade boundary layer interactions is included. All the components of turbulent intensities and stresses are found to be high in the leakage-flow mixing region. The radial component of intensities and stresses is found to be much higher than the corresponding streamwise components. The turbulent spectra clearly reveal the decay process of the inlet-guide-vane wake within the rotor passage.

Gas turbine blade with chamber for circulation of cooling fluid and process for its manufacture

Abstract: A gas turbine blade in which are provided inter alia two chambers for the circulation of cooling fluid, lying in proximity to the leading and trailing edges of the blade respectively, two conduits disposed respectively between the first chamber and the leading edge and between the second chamber and the trailing edge, and apertures connecting each conduit to the chamber nearest to it. The blade is produced by diffusion-brazing two half-blades together. The chambers are formed in part in one, and in part in the other, half blade. In order to produce the conduits and apertures by means of grooves engraved in the facing surface of the two half blades, the interface of the two half-blades divides said conduits and apertures in two.

Variable camber aircraft wing tip

Abstract: A variable wing tip (18) for securing to a variable camber leading edge portion (12) of an aircraft wing (10). The wing tip has a normal fore and aft camber in an upper position and is adapted to have an increased camber curvature in a deflected down position. In the down position, the wing tip has a compound curvature, is shortened in the fore and aft direction ad is generally thickened in depth in a substantial portion of the wing tip. By operation of an actuator (96) in the leading edge (21), a multiple of generally parallel scissor plates (30,32) are pivoted apart. The plates extend in an inboard outboard direction and are spaced in the fore and aft direction. The scissor plates are pairs of upper and lower plates with respect to the upper and lower surface of the wing tip, the upper plates (30) being adapted to be pivotally secured inboard to a complementary outer portion of the wing leading edge to secure the upper portion of the wing tip thereto, and the lower plates (32) being pivotally and slidably secured inboard within the wing tip and being adapted to be secured to a complementary outer portion of the wing leading edge to secure the lower portion of the wing tip thereto. Each respective pair of upper and lower plates are pivotally secured together outboard (40) adjacent and outboard leading edge (42) of the wing tip. Respective upper and lower plates are adapted to be pivoted apart inwardly of and on the outboard pivots (40) when the camber is increased in the down position so as to increase the depth of and shorten the fore and aft direction of the wing tip. An elastomer (84) is bonded to the plates, permitting the plates to be pivoted apart. An elastomeric wing tip configuration (88), is secured to the wing leading edge and to a forwardly facing edge of the wing. The elastomeric configuration has the normal camber of the wing tip in the upper position and the increased camber and compound curvature of the wing tip in the deflected position. It forms continuous surfaces in the varied camber positions with the wing.

The Effect of Bending-Torsion Coupling on Fan and Compressor Blade Flutter

Abstract: A study of the effects of bending-torsion interaction of the flutter boundaries of turbomachinery blading is presented. The blades are modeled as equivalent sections, and the equations of motion allow for the general case of structural, inertial and aerodynamic coupling, in the presence of structural damping. Two different speed regimes are investigated: incompressible flow, and supersonic flow with a subsonic leading edge locus. Flutter boundaries are presented for cascade design parameters representative of current technology fan rotors. These results illustrate that bending-torsion interaction has a pronounced effect on the flutter boundaries for both speed regimes, although the mode frequencies show no appreciable tendency to coalesce as flutter is approached. Several cases of bending branch instability were observed, without incorporating the effects of finite mean lift or strong shocks in the analysis.

Comparison of sharp and blunt fin-induced shock wave/turbulent boundary-layer interaction

Abstract: In this paper, results are presented from an experimental study of fin-induced shock wave/turbulent boundary-layer interaction. Semi-infinite fin models, with sharp and hemicylindrically blunted leading edges, were tested at Mach 3 in two high Reynolds number, adiabatic wall, turbulent boundary layers. Detailed streamwise surface pressure distributions were measured at several spanwise stations for angles of attack from 0 to 12 deg. The objective was to examine the spanwise development of the interaction region and determine its dependence on leading edge geometry and the incoming-flow parameters. The results show that for blunt fins there is a region of the interaction in the vicinity of the centerline where the scale and characteristics of the disturbed flowfield are controlled primarily by the leading edge diameter. Outboard of this inner, or "leading-edge dominated," region the interaction properties and spanwise development are essentially the same as if the leading edge were sharp. Thus there is an outer region of the blunt fin-induced flowfield in which the properties are effectively "independent of leading edge blunting." Nomenclature D = blunt fin leading-edge diameter h - fin height Lu = upstream influence measured relative to the freestream shock wave M = Mach number P = static pressure P2 = wall pressure downstream of interaction Re = Reynolds number X = coordinate in the plane of the test surface and aligned with the tunnel axis, with X= 0 at the fin leading edge Xs = coordinate in the X direction measured relative to the freestream shock wave Y = coordinate normal to the X axis in the plane of the test surface, with Y= 0 at the fin leading edge

Device in the delivery of sheet-fed rotary printing machine for inhibiting curl formation in the leading edge of delivered sheet

Abstract: A delivery system of a sheet-fed rotary printing machine includes a device for inhibiting curl formation in a leading edge of a sheet, comprising blowing units disposed at a location of the delivery system whereat, as viewed in travel direction of the sheet, a forward region of a stack of delivered sheets is disposed. The blowing units are in vicinity of a location at which the leading edge of the sheet being delivered travels downwardly. The blowing units have respective means for directing a controlled air jet, with one directional component in the travel direction of the sheet and another directional component towards the middle of the sheet, against the underside of the downwardly traveling sheet in vicinity of the leading edge of the sheet tending to curl.

Numerical simulation of wing-fuselage interference

Abstract: T wing-fuselage problem is investigated by means of the Navier-Stokes equations incorporated with an approximate turbulence model. The numerical solution yields a reasonable global agreement with experimental data in static and impact pressure distributions. However, in order to better describe the flowfield structure near the leading edge of the wing, an alternative choice of coordinate system is required.

Geometry of subducted plates and island arcs viewed as a buckling problem

Abstract: During subduction, a plate9s leading edge is forced to squeeze into a smaller total width than it occupied at the surface, yet the plate resists change of dimension in its own plane. A possible outcome is that the plate buckles as it descends: at a depth of 500 km, the amplitude of buckling could be as much as one-tenth the wavelength, and dip angles for a plate descending at 45° on average could range from 35° to 55°. Observed plate geometries, though not ideally regular, are compatible with these suggestions. Before being subducted a lithosphere plate is convex upward at all points, whereas, to buckle, it must become convex downward in a series of plunging synclines. The “pop-through” is most likely to occur after the leading edge has descended about 100 km. At the moment of pop-through, the subduction hinge line changes from being smooth to being a series of arcs and cusps; back-arc spreading associated with the change can be 150 km or more in the arc embayments but should be zero opposite the cusps, where a compressive regime is sustained. In the pop-through event, elastic strain energy is released, and the associated back-arc spreading should be a comparatively sudden event, on a geologic time scale.

Non-linear digital filter

Abstract: A filter for a pulse form of signal which substantially reduces sidebands, yet provides an output signal which is substantially free of jitter and at the same time has no inter-symbol interference. The filter comprises a circuit for detecting the pulse type of signal and for substituting .pi. radians of a sine wave signal with a positive slope upon detection of a positive going leading edge of the pulse type signal and .pi. radians of a sine wave signal with a negative slope upon detection of a negative going trailing edge of the pulse type signal, at an output terminal.

Numerical experiments on the leading-edge flowfield

Abstract: A numerical procedure has been developed to analyze the compressible inviscid flowfield in the leading-edge region of an airfoil, using one of the possible variations of the A scheme: a finite-difference time-dependent method which emphasizes the role of the domain of dependence of each computed point in a transient. The use of artificial permeable boundaries allows the computational domain to be confined to a small region surrounding the nose of the airfoil. A general procedure is described for treating the boundaries without violating the basic principles of the A scheme and maintaining accuracy at points next to the boundaries.

Method and device for decreasing the flow resistance on wings particularly aerofoils and blades of turbomachines exposed to gas flux such as air

Abstract: A device and method for reducing the flow resistance of air foils utilizes one or more flow compressors on the leading edge of the foil. The gas flowing in front of the leading edge is compressed and is expanded on the upper surface of the wing so that through induction of the expanded gas, gas from the area of the lower surface of the wing is admixed to increase the flow on the upper surface of the wing. The direction of the admixed gases is such that the velocity of the gas flowing out at the trailing edge of the wing is identical over the entire wing length. The separation point of the flow is displaced towards the trailing edge of the wind and, in the direction of the tip of the wing, the pressure of the gas flowing over the upper surface of the wing is increased.

Conditions for spreading as single molecules or as a thin planar drop

Abstract: Two angles can be associated with a liquid drop on a solid surface: the angle θ 0 at the leading edge of the drop, and the thermodynamic wetting angle θ that occurs at some distance from the leading edge. Since these two angles differ ( θ 0 > θ ), small amounts of the liquid can spread on the surface in either of two ways: (a) as single molecules, or (b) as a very thin extended planar drop with a rapidly changing profile near the leading edge. Spreading as single molecules occurs when values larger than unity result from the equation, for cos θ 0 , while spreading in a planar configuration occurs when the equations lead to cos θ > 1 and cos θ 0

Improved Potential Gradient Method to Calculate Airloads on Oscillating Supersonic Interfering Surfaces

Abstract: sectional normal force coefficient in AGARD notation, Eq. (7) = sectional pitching moment about quarter-chord point coefficient in AGARD notation, Eq. (8), C*mi is moment about midchord = variation of the potential jump behind the leading edge of an element on which a constant potential gradient in streamwise direction is assumed (Fig. Ib) = amplitude of the normal displacement of the oscillating wing = reduced frequency = ul/ U supersonic pressure dipole kernel = reference length = Mach number = normal directions of receiving wings = total number of unknowns = dynamic pressure = /2pU generalized aerodynamic coefficients , generalized forces, Eq. (6)

Effects of wing-leading-edge modifications on a full-scale, low-wing general aviation airplane: Wind-tunnel investigation of high-angle-of-attack aerodynamic characteristics. [conducted in Langley 30- by 60-foot tunnel

Abstract: Wing-leading-edge modifications included leading-edge droop and slat configurations having full-span, partial-span, or segmented arrangements. Other devices included wing-chord extensions, fences, and leading-edge stall strips. Good correlation was apparent between the results of wind-tunnel data and the results of flight tests, on the basis of autorotational stability criterion, for a wide range of wing-leading-edge modifications.

Interferometric study of combined forced and natural convection

Abstract: When a metal is electrodeposited in laminar flow onto a planar cathode facing upward in a channel, mass transfer rates are enhanced downstream beginning at some distance from the leading edge, because of the contribution of free convection to the fluid motion. The position and extent of this secondary flow has been clearly demonstrated by double beam interferometry, employed for the measurement of concentration profiles in the cathodic boundary layer. A narrow range of critical Rayleigh numbers, , has been associated with the onset of natural convection. This result provides a convenient basis for the prediction of the transition from forced convection control to combined forced and natural convection control of the mass transfer process.