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Showing papers on "Leading edge published in 1995"


Journal ArticleDOI
Abstract: An experimental study was made of the effect of a periodic velocity perturbation on the separation bubble downstream of the sharp-edged blunt face of a circular cylinder aligned coaxially with the free stream. Velocity fluctuations were produced with an acoustic driver located within the cylinder and a small circumferential gap located immediately downstream of the fixed separation line to allow communication with the external flow. The flow could be considerably modified when forced at frequencies lower than the initial Kelvin-Helmholtz frequencies of the free shear layer, and with associated vortex wavelengths comparable to the bubble height. Reattachment length, bubble height, pressure at separation, and average pressure on the face were all reduced. The effects on the large-scale structures were studied on flow photographs obtained by the smoke-wire technique. The forcing increased the entrainment near the leading edge. It was concluded that the final vortex of the shear layer before reattachment is an important element of the flow structure. There are two different instabilities involved, the Kelvin-Helmholtz instability of the free shear layer and the “shedding” type instability of the entire bubble. A method of frequency scaling is proposed which correlates data for a variety of bubbles and supports an analogy with Karman vortex shedding.

194 citations


Journal ArticleDOI
TL;DR: In this article, a computational analysis was performed on a compressible flow oscillation due to shear layer instabilities over a cavity and pressure feedback in a cavity of length-to-depth ratio 3 at Mach 1.5 and 2.5.
Abstract: A computational analysis was performed on a compressible flow oscillation due to shear layer instabilities over a cavity and pressure feedback in a cavity of length-to-depth ratio 3 at Mach 1.5 and 2.5. The mass-averaged NavierStokes equations were solved. Turbulence closure was achieved using a k-u> model with compressibility corrections. Self-sustained oscillations were produced. Negative form drag coefficient was observed within an oscillatory cycle due to mass ejection from the cavity near the trailing edge and vortex production near the leading edge. The shock wave-expansion wave interaction patterns, modes of the oscillation, sound pressure level, and time-averaged surface pressure were compared with experimental results of previous investigations and good agreement was achieved, particularly the time-averaged pressure. The prediction showed a marked improvement over earlier analysis.

119 citations


Journal ArticleDOI
TL;DR: In this article, the performance degradation of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported, where a 0.025 mm (0.001 in) thick rough coating with a surface finish of 2.54-3.18 rms (100-125 rms μm) is applied to the rotor blades.
Abstract: The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54-3.18 rms μm (100-125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254-0.508 rms μm (10-20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms μm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier-Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

119 citations


Journal ArticleDOI
TL;DR: In this paper, a three-dimensional icing model was developed at ONERA to calculate ice accretion shapes for aerodynamic components that can not be predicted using conventional two-dimensional codes.
Abstract: A three-dimensional icing model has been developed at ONERA to calculate ice accretion shapes for aerodynamic components that can not be predicted using conventional two-dimensional codes. It is described, emphasizing the original parts with respect to the two-dimensional existing models. The model includes Euler inviscid flow calculation. Droplet trajectories are calculated in a three-dimensional grid. The remesh on the leading edge is adapted to follow aerodynamics singularities. The boundary layer is calculated using a mixing length formulation to model the wall roughness influence on convective heat transfer. Runback paths are integrated. The heat balance is calculated in a grid created along the runback paths. The domain of validity of the three-dimensional icing code is described; compared with the two-dimensional model this domain is wider, especially for high speeds. The three-dimensional model is shown to simulate well a uniform ice deposit on a three-dimensional rotor blade tip. Then, a comparison of the three- and two-dimensional codes on an infinite swept wing shows that the corrected two-dimensional code predicts the catch efficiency but not the ice shape. Finally, it is shown that the continuum flux hypothesis prevents the three-dimensional model from simulating correctly the "lobster tail" ice shape (nonuniform ice deposit).

104 citations


Patent
13 Oct 1995
TL;DR: In this article, a gas turbine vane has an inner shroud (26) that is cooled by a portion of the cooling air directed to a cavity between two adjacent rows of discs (55, 56).
Abstract: A gas turbine vane (17) having an inner shroud (26) that is cooled by a portion of the cooling air directed to a cavity (45) between two adjacent rows of discs (55, 56). A portion of the cooling air in the cavity flows through impingement plates (83, 84) and impinges on the inner (98) surface of the inner shroud (26). Another portion of the cooling air flows through a passage (88) in the leading edge (42) of the inner shroud that has a pin fin (89) array for enhanced cooling. The impingement plates form chambers that collect both the impingement air and the pin fin passage air and direct it through holes (92) in the trailing edge (43) of the inner shroud for cooling of the trailing edge. Longitudinal passages (93, 94) along the side of the inner shroud direct the cooling air from the pin fin passage to the trailing edge (43).

100 citations


Patent
20 Nov 1995
TL;DR: In this article, the shape memory alloy tendons are separately connected to a controlled source of electrical current such that tendons of the first and second flexible facesheets can be selectively heated in an antagonistic, slack-free relationship, to bring about a desired modification of the configuration of the control surface.
Abstract: A pliant, controllable contour control surface comprising a first flexible facesheet formed to a first initial contour of the control surface, and a second flexible facesheet formed to a second initial contour of the control surface. The first and second facesheets each have a set of prestrained shape memory alloy tendons embedded therein, extending from a leading edge to a trailing edge of the control surface. Each set of the shape memory alloy tendons is separately connected to a controlled source of electrical current such that tendons of the first and second flexible facesheets can be selectively heated in an antagonistic, slack-free relationship, to bring about a desired modification of the configuration of the control surface. A computer based control system is utilized for maintaining a constant temperature of the antagonists to establish conditions conducive to the stress induced transformation from austenite to martensite, accomplished by causing constant current to flow through the antagonists. Proportional/integral (PI) control is utilized in connection with the opposing shape memory tendons.

100 citations


Journal ArticleDOI
TL;DR: In this article, a model for the sound generated when a convected vortical or entropic gust encounters an airfoil at non-zero angle of attack is presented.
Abstract: A theoretical model is developed for the sound generated when a convected vortical or entropic gust encounters an airfoil at non-zero angle of attack. The theory is based on a linearization of the Euler equations about the steady subsonic flow past the airfoil. High-frequency gusts, whose wavelengths are short compared to the airfoil chord, but long compared to the displacement of the mean-flow stagnation point from the leading edge, are considered. The analysis utilizes singular-perturbation techniques and involves four asymptotic regions. Local regions, which scale on the gust wavelength, are present at the airfoil leading and trailing edges. Behind the airfoil a ‘transition’ region, which is similar to the transition zone between illuminated and shadow zones in optical problems, is present. In the outer region, far away from the airfoil edges and wake, the solution has a geometric-acoustics form. The primary sound generation is found to be concentrated in the local leading-edge region. The trailing edge plays a secondary role as a scatterer of the sound generated in the leading-edge region. Parametric calculations are presented which illustrate that moderate levels of airfoil steady loading can significantly affect the sound field produced by airfoil-gust interactions.

84 citations


Patent
20 Oct 1995
TL;DR: In this paper, a system for circulating heated gases within the circular leading edge of a jet engine housing to prevent ice build-up thereon, or to remove accumulated ice there is described.
Abstract: A system for circulating heated gases within the circular leading edge of a jet engine housing to prevent ice build-up thereon, or to remove accumulated ice thereform. Hot gases such as air from a hot, high pressure section of the jet-engine are directed through a conduit. The conduit enters the annular leading edge housing, usually from the aft side through a bulkhead, then turns about 90° to a direction tangential to the leading edge annulus. The hot gases exiting the tube entrain the cooler air in the housing, causing a much larger mass of air to swirl circularly around the annular housing. The entering hot gasses heat the mass of air to an intermediate, but still relatively hot, temperature. This large mass of circularly moving hot air is quite efficient in uniformly transferring heat to the skin of the leading edge without leaving any relatively cold areas and preventing the formation of ice thereon.

84 citations


Journal ArticleDOI
TL;DR: In this article, the spreading of an insoluble monolayer containing a fixed mass of surfaceactive material over the initially horizontal free surface of a viscous fluid layer is investigated.
Abstract: The unsteady spreading of an insoluble monolayer containing a fixed mass of surface-active material over the initially horizontal free surface of a viscous fluid layer is investigated. A how driving the spreading is induced by gradients in surface tension, which arise from the nonuniform surfactant distribution. Distinct phases in the flow's dynamics are distinguished by a time T = H-0(2)/v, where H-0 is the fluid depth and v its viscosity. For times t << T, i.e. before the lower boundary has any significant influence on the flow, a laminar sub-surface boundary-layer flow is generated. The effects of gravity, capillarity, surface diffusion or surface contamination may be weak enough for the flow to drive a substantial unsteady displacement of the free surface, upward behind the monolayer's leading edge and downward towards its centre. Similarity solutions are identified describing the spreading of a localized planar monolayer strip (which spreads like t(1/2)) or an axisymmetric drop (which spreads like t(3/8)); using the Prandtl transformation, the associated boundary-layer problems are solved numerically. Quasi-steady sub-layers are shown to exist at the centre and at the leading edge of the monolayer; that due to surface contamination, for example, may eventually grow to dominate the flow, in which case spreading proceeds like t(3/4). Once t = O(T), vorticity created at the free surface has diffused down to the lower boundary and the flow changes character, slowing appreciably. The dynamics of this stage are modelled by reducing the problem to a single nonlinear diffusion equation. For a spreading monolayer strip or drop, the transition from an inertia-dominated (boundary-layer) flow to a viscosity-dominated (thin-film) flow is predicted to be largely complete once t approximate to 85T.

77 citations


Patent
14 Mar 1995
TL;DR: In this article, the authors present a wind turbine blade with upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effect.
Abstract: Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.

75 citations


Patent
31 Jan 1995
TL;DR: In this article, the leading edge of a turbine shroud has a leading edge serpentine channel having a bend passage, which includes a purge hole to avoid separating flow in the bend passage.
Abstract: A turbine shroud segment for use in a gas turbine engine includes a serpentine channel along at least one axial edge of the segment. Various construction details are developed that disclose a channel for efficiently flowing cooling fluid through an axial edge of a shroud segment. In a particular embodiment, a turbine shroud segment includes a leading edge serpentine channel having a bend passage which includes a purge hole to avoid separating flow in the bend passage.

Patent
07 Jun 1995
TL;DR: In this paper, a negative pressure air bearing slider having at least one trailing edge pocket defined by a generally U-shaped rail open to the trailing edge of the slider, and at least two leading edge pocket defining by a Ushaped rail opened to the leading edge of a slider is described.
Abstract: Disclosed is a negative pressure air bearing slider having at least one trailing edge pocket defined by a generally U-shaped rail open to the trailing edge of the slider, and at least one leading edge pocket defined by a generally U-shaped rail open to the leading edge of the slider Also disclosed is a method for determining an optimal width for a leading edge pocket of predetermined length and depth to provide a flat fly height profile in combination with a trailing edge pocket of predetermined dimensions For a given trailing edge pocket configuration, width optimization is achieved by constructing a number of prototype sliders, each having a unique leading to trailing edge pocket width ratio; measuring the inner and outer diameter fly heights for each prototype to obtain ID and OD fly height profile curves; and determining a leading edge pocket width corresponding to the point of intersection of the curves

Patent
11 Dec 1995
TL;DR: In this article, a light spot is formed on a recording medium on which information is recorded so that a phase of light reflected by a small region as an information pit is different from other regions.
Abstract: In an optical information recording/reproducing apparatus, the reproduction of information is performed as following manner. A light spot is formed on a recording medium on which information is recorded so that a phase of light reflected by a small region as an information pit is different from a phase of light reflected by other regions. Upon scanning the recording medium with the light spot, a leading edge and a trailing edge of the information pit in the scanning direction are respectively detected. A detection signal for the leading edge of the information pit and a detection signal for the trailing edge of the information pit are synthesized to effect the reproduction of information.

Patent
26 Jan 1995
TL;DR: In this paper, a gas turbine engine with an airfoil portion with a leading and a trailing edge (60, 62) includes dual pressure source cooling, where the higher internal pressure in the leading edge ensures that the inward flow of products of combustion does not occur.
Abstract: An airfoil (42) for a gas turbine engine (10) having an airfoil portion (52) with a leading and a trailing edge (60, 62) includes dual pressure source cooling. The trailing edge (62) of the airfoil (42) includes an internal trailing edge passage (78) and the leading edge (60) includes an internal leading edge passage (64). Compressor bleed air at higher pressure is channeled through the leading edge passage (64) whereas compressor bleed air at lower pressure is channeled through the trailing edge passage (78). The higher internal pressure in the leading edge (60) ensures that the inward flow of products of combustion does not occur.

Journal ArticleDOI
TL;DR: In this article, a theoretical and experimental study is presented of the aeroelastic instability of the human soft palate, which can explain the occurrence of snoring, and an experimental apparatus which produces sounds very close to human snoring is described.

Patent
24 May 1995
TL;DR: A shingle is formed with a first, or trailing, side edge folded over toward the top surface of the shingle to catch moisture and direct it down to the top surfaces of the next lower shingle as mentioned in this paper.
Abstract: A shingle is formed with a first, or trailing, side edge folded over toward the top surface of the shingle to catch moisture and direct it down to the top surface of the next lower shingle. A second, or leading, side edge of the shingle is not folded. An "S" shaped fold, spaced apart from the trailing edge fold, receives the unfolded leading edge of an adjacent shingle. The trailing edge and S fold form a gutter under the leading edge of the adjacent shingle. The top edge of the shingle is folded-over toward the top surface of the shingle for engagement with the folded-under lower edge of the next higher shingle. The folded-over top edge extends to the right or leading edge of the shingle and slides under the left edge of the folded-over top edge of the next adjacent shingle to the fight side. The folded-under lower edge extends along the lower edge of the shingle but leaves a gap relative to the folded-under lower edge of the lateral adjacent shingle. In this way water is caught by the gutter and flows out the gap between folded-under lower edges of laterally adjacent shingles. One or more S-shaped folds are also made in the middle of the shingle to form a panel having the appearance of multiple shingles.

Journal ArticleDOI
TL;DR: In this article, the authors studied the dynamic stall process of an NACA 0012 airfoil undergoing a constant-rate pitching-up motion in a water towing tank facility.
Abstract: The dynamic stall process of an NACA 0012 airfoil undergoing a constant-rate pitching-up motion is studied experimentally in a water towing tank facility. This study focuses on the detailed measurement of the unsteady separated flow in the vicinity of the leading and trailing edges of the airfoil. The measurements are carried out using the particle image velocimetry technique. This technique provides the two-dimensional velocity and associated vorticity fields, at various instants in time, in the midspan of the airfoil. Near the leading edge, large vortical structures emerge as a consequence of van Dommelen and Shen type separation and a local vorticity accumulation. The interaction of these vortices with the reversing boundary-layer vorticity initiates a secondary flow separation and the formation of a secondary vortex. The mutual induction of this counter-rotating vortex pair eventually leads to the ejection process of the dynamic stall vortex from the leading-edge region. It is found that the trailing-edge flowfield only plays a secondary role on the dynamic stall process.

Patent
26 Dec 1995
TL;DR: A leading edge flap for supersonic transport airplanes is described in this article, where the upward curve of the leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex).
Abstract: A leading edge flap (16) for supersonic transport airplanes is disclosed. In its stowed position, the leading edge flap forms the lower surface of the wing leading edge up to the horizontal center of the leading edge radius. For low speed operation, the vortex leading edge flap moves forward and rotates down. The upward curve of the flap leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex). The rounded shape of the upper fixed leading edge provides the conditions for a controlled reattachment of the flow on the upper wing surface and therefore a stable vortex. The vortex generates lift and a nose-up pitching moment. This improves maximum lift at low speed, reduces attitude for a given lift coefficient and improves lift to drag ratio. The mechanism (27) to move the vortex flap consists of two spanwise supports (24) with two diverging straight tracks (64 and 68) each and a screw drive mechanism (62) in the center of the flap panel (29). The flap motion is essentially normal to the airloads and therefore requires only low actuation forces.

Patent
25 Sep 1995
TL;DR: In this paper, a metal solid section, composite or structural/syntactic foam segments, and metal solid spars are all attached together to define an airfoil portion.
Abstract: A lightweight, impact-resistant gas turbine blade, such as an aircraft engine fan blade, has a metal solid section, composite or structural/syntactic foam segments, and metal solid spars all attached together to define an airfoil portion. The solid section includes the leading edge, blade tip, and trailing edge. The segments together are bounded in part by the solid section near the leading edge, blade tip, and trailing edge. The solid spars separate and are attached to the segments.

Journal ArticleDOI
TL;DR: In this article, the characteristics of flow developments above 50-degrees sweep delta wings with different leading-edge profiles are shown by flow visualizations and velocity measurements, and it is noted that the flow angles associated with the separated shear layers vary with the leading edge profiles studied.
Abstract: The characteristics of flow developments above 50-deg sweep delta wings with different leading-edge profiles are shown by flow visualizations and velocity measurements. The Reynolds number based on freestream velocity and root chord is about 7 x 103. The leading-edge profiles studied include the shapes of square, round, windward surface beveling, leeward surface beveling, and wedge. Based on the velocity data obtained along the leading edges of the delta wings it is noted that the flow angles associated with the separated shear layers vary with the leading-edge profiles studied. This finding infers that varying the leading-edge profile has an impact on the initial development of the separated shear layer, consequently, the formation of leading-edge vortex. Furthermore, it is shown that the leading edge of windward beveling causes the largest leading-edge flow angle and produces the most organized leading-edge vortex.

Patent
13 Nov 1995
TL;DR: In this paper, the authors describe an actuator-controlled optimized emulation of the rotation of a fixed wing for hovering and vertical takeoff and landing in a fixed-wing aircraft with a ducted fan.
Abstract: The craft is for hovering flight, vertical takeoff and landing, and horizontal forward flight. It has a tail-sitting fuselage and a ducted fan mounted to the fuselage aft to provide propulsion in both (a) hovering and vertical flight and (b) horizontal forward flight. At each side is a floating wing, supported from the fuselage for passive rotation (or an actuator-controlled optimized emulation of such rotation) about a spanwise axis, to give lift in forward flight. The fuselage attitude varies between vertical in hovering and vertical flight, and generally horizontal in forward flight. Preferably the fuselage is not articulated; there is just one fan, the sole source of propulsion, rotating about only an axis parallel to the fuselage; and thrust-vectoring control vanes operate aft of the fan. Preferably at each side a small, nonrotating wing segment is fixed to the fuselage, and the floating wing defines--along its trailing portions--a corner notch or slot near the fuselage; forward portions of the fixed wing segment are within this notch. Preferably the spanwise axis is along a surface of the floating wing, and a long hinge supports that wing from the fixed wing segment, within the notch. During vertical and transitional flight characteristically the leading edge of the floating wing is down relative to the fuselage axis.

Journal Article
TL;DR: In this article, the authors consider the linearised stability characteristics of the thermal boundary layer induced by the uniform heating of a semi-infinite vertical surface embedded in a fluid - saturated porous medium.
Abstract: We consider the linearised stability characteristics of the thermal boundary layer induced by the uniform heating of a semi-infinite vertical surface embedded in a fluid - saturated porous medium. In this paper attention is restricted to two-dimensional disturbances far from the leading edge. This analysis complements and extends the direct numerical simulation of Rees [1993] which shows that the flow is stable at locations sufficiently close to the leading edge. In this asymptotic regime we also find that wave disturbances decay. However, the rate of decay decreases as the distance downstream of the leading edge increases.

Patent
06 Feb 1995
TL;DR: In this paper, a tripad air bearing slider is designed with two outer pads that extend from two or more tapered sections at the leading edge of the slider towards the trailing edge.
Abstract: A tripad air bearing slider useful in a disk drive is designed with two outer pads that extend from two or more tapered sections at the leading edge of the slider towards the trailing edge. The pads are configured and angled so that a desired lift force is obtained that acts in opposition to a force provided by a spring loaded flexure or load beam. The angles that define the shape of the three pads are formed by simple mechanical cutting, using a diamond cutting wheel, for example, or alternatively by ion milling or reactive ion etching.

Patent
22 Aug 1995
TL;DR: In this paper, a wear protector system for shielding the leading edge of an earthmoving implement such as a loading bucket includes an arrangement of shrouds which cover the leading edges between the laterally spaced digger teeth.
Abstract: A wear protector system for shielding the leading edge of an earthmoving implement such as a loading bucket includes an arrangement of shrouds which cover the leading edge between the laterally spaced digger teeth. Each shroud has a nose portion that wraps around the leading edge and a rearwardly extending tail portion having an abutment surface that engages an undercut abutment surface of an anchor block that is welded to the bucket. The complementary abutment surfaces retain the tail portion from moving upwardly away from the bucket surface, whereas the shroud is retained from movement forwardly out of engagement with the lip by use of a cotter pin extending transversely through a passage that is formed partially in the tail of the shroud and partially in the anchor.

Patent
10 Oct 1995
TL;DR: In this paper, a vane assembly for a gas turbine engine is provided, which includes a plurality of vanes, an inner vane support, a casing, and apparatus for maintaining a pressure difference.
Abstract: A vane assembly for a gas turbine engine is provided, which includes a plurality of vanes, an inner vane support, a casing, and apparatus for maintaining a pressure difference. Each vane has a leading edge, a trailing edge, an outer radial end, an inner radial end, and an internal cavity. The internal cavity includes a forward compartment adjacent the leading edge and an aft compartment adjacent the trailing edge. The casing, which includes an annulus, is positioned radially outside of the vanes. The vanes extend between the inner vane support and the casing. The apparatus for maintaining a pressure difference maintains a difference in the cooling air pressure within the forward and aft compartments of the vane cavity under operating conditions.

Patent
17 Jul 1995
TL;DR: In this paper, a windmill with a plurality of hydrodynamic, folding blades (12, 14, 16 and 18) secured to a hub (64) is described, where the blades are propelled by wind current such that the windmill rotates about a central axis of a shaft (20) on which a hub is journalled.
Abstract: Windmill (10) having a plurality of hydrodynamic, folding blades (12, 14, 16 and 18) secured to a hub (64). The blades are propelled by wind current such that the windmill rotates about a central axis of a shaft (20) on which a hub (64) is journalled. Each folding blade extends radially from the hub and includes a lower blade portion (24) attached to the hub and an upper blade portion (22) pivotally attached to the lower blade portion. The lower blade portion includes a hydrodynamic surface extending from a first edge(43) and terminating at a trailing edge (42). The upper blade portion is hydrodynamically contoured so as to form a nose cone (44). An apex of the nose cone forms a leading edge (40) of the upper blade portion that is opposite a trailing edge (41). The nose cone is weighted such that the trailing edge of the upper blade portion balances in a position slightly separate from the trailing edge of the lower blade portion when the fluid current fails to flow at a speed that exceeds a minimum threshold. However, when the fluid current impinges the trailing edge of the upper blade portion and the fluid current flows at a speed that exceeds a minimum threshold, the weighted nose cone of the upper blade portion causes the upper blade portion to pivot away from the lower blade portion so that the trailing edge of the upper blade portion is separated from the trailing edge of the lower blade portion at an acute angle. Accordingly, when the fluid current ceases to impinge the trailing edge of the upper blade portion, the weighted nose cone of the upper blade portion causes the upper blade portion to pivot toward the lower blade portion such that the trailing edge of the upper blade portion approaches the trailing edge of the lower blade portion.

Journal ArticleDOI
TL;DR: In this paper, the spread rate and temperature measurements in the gas and solid phases, and also recordings of the flame from ignition to extinction using two 16mm cameras, were gathered for two different oxygen levels and three different pressures.

Patent
06 Jun 1995
TL;DR: In this article, a planform and a high-lift airfoil are used for a vehicle engine-cooling fan assembly with a curved planform, where the planform has a first region adjacent the root of the blade with forward curvature, a second region with backward curvature and an intermediate region disposed between the first region and the second with substantially straight curvature.
Abstract: A blade for a vehicle engine-cooling fan assembly having a curved planform and a high-lift airfoil. The planform has a first region adjacent the root of the blade with forward curvature, a second region adjacent the tip of the blade with backward curvature, and an intermediate region disposed between the first region and the second region with substantially straight curvature. The airfoil has a leading edge; a rounded, bulbous nose section adjacent the leading edge; a trailing edge; a curved pressure surface extending smoothly and without discontinuity from the nose section to the trailing edge; a curved suction surface extending smoothly and without discontinuity from the nose section to the trailing edge; and a thin, highly cambered aft section formed adjacent the trailing edge and between the pressure surface and the suction surface. The nose section has a thickness which is greater than the thickness of the airfoil between the pressure surface and the suction surface and the nose section blends smoothly into the pressure surface and the suction surface.

Patent
22 Dec 1995
TL;DR: In this paper, a fuel injector includes outer and inner coaxial shells spaced radially apart to define a flow channel there between having an inlet and an outlet, and a strut extends radially outwardly from the inner and outer shells at leading edges thereof and is fixedly joined thereto.
Abstract: A fuel injector includes outer and inner coaxial shells spaced radially apart to define a flow channel therebetween having an inlet and an outlet. A strut extends radially outwardly from the inner and outer shells at leading edges thereof and is fixedly joined thereto. An annular lobed mixer is disposed coaxially in the channel and includes a leading edge, a trailing edge spaced from the channel outlet to define a mixing nozzle, and a plurality of circumferentially spaced apart lobes increasing in radial height from the leading to trailing edges of the mixer. The lobes defines with the outer and inner shells corresponding pluralities of outer and inner chutes for separately channeling respective portions of inlet air. The fuel is injected into the lobed mixer forming a fuel and air mixture in the mixing nozzle for discharge through the channel outlet into a combustor.

Proceedings ArticleDOI
TL;DR: In this article, a wind tunnel test facility equipped with a linear cascade of film cooled vane airfoils was used in the simultaneous determination of the local gas side heat transfer coefficients and the adiabatic film cooling effectiveness.
Abstract: A warm (315°C) wind tunnel test facility equipped with a linear cascade of film cooled vane airfoils was used in the simultaneous determination of the local gas side heat transfer coefficients and the adiabatic film cooling effectiveness. The test rig can be operated in either a steady-state or a transient mode. The steady-state operation provides adiabatic film cooling effectiveness values while the transient mode generates data for the determination of the local heat transfer coefficients from the temperature-time variations and of the film effectiveness from the steady wall temperatures within the same aerothermal environment. The linear cascade consists of five airfoils. The 14 percent cascade inlet free-stream turbulence intensity is generated by a perforated plate, positioned upstream of the airfoil leading edge. For the first transient tests, five cylinders having roughly the same blockage as the initial 20 percent axial chord of the airfoils were used. The cylinder stagnation point heat transfer coefficients compare well with values calculated from correlations. Static pressure distributions measured over an instrumented airfoil agree with inviscid predictions. Heat transfer coefficients and adiabatic film cooling effectiveness results were obtained with a smooth airfoil having three separate film injection locations, two along the suction side, and the third one covering the leading edge showerhead region. Near the film injection locations, the heat transfer coefficients increase with the blowing film. At the termination of the film cooled airfoil tests, the film holes were plugged and heat transfer tests were conducted with non-film cooled airfoils. These results agree with boundary layer code predictions.