Showing papers on "Leading edge published in 1994"

The effects of wing rotation on unsteady aerodynamic performance at low reynolds numbers

Abstract: The downstroke-to-upstroke transition of many insects is characterized by rapid wing rotation. The aerodynamic consequences of these rapid changes in angle of attack have been investigated using a mechanical model dynamically scaled to the Reynolds number appropriate for the flight of small insects such as Drosophila. Several kinematic parameters of the wing flip were examined, including the speed and axis of rotation, as well as the duration and angle of attack during the wing stroke preceding rotation. Alteration of these kinematic parameters altered force generation during the subsequent stroke in a variety of ways. 1. When the rotational axis was close to the trailing edge, the model wing could capture vorticity generated during rotation and greatly increase aerodynamic performance. This vortex capture was most clearly manifested by the generation of lift at an angle of attack of 0°;. Lift at a 0°; angle of attack was also generated following rotation about the leading edge, but only if the downstroke angle was large enough to generate a von Karman street. The lift may be due to an alteration in the effective angle of attack caused by the inter-vortex stream in the downstroke wake. 2. The maximum lift attained (over all angles of attack) was substantially elevated if the wing translated backwards through a wake generated by the previous stroke. Transient lift coefficient values of nearly 4 were obtained when the wing translated back through a von Karman street generated at a 76.5°; angle of attack. This effect might also be explained by the influence of the inter-vortex stream, which contributes a small component to fluid velocity in the direction of translation. 3. The growth of lift with angle of attack was significantly elevated following a 7.5 chord stroke with a 76.5°; angle of attack, although it was relatively constant under all other kinematic conditions. 4. The results also indicate the discrepancies between transient and time-averaged measures of performance that arise when unsteady mechanisms are responsible for force generation. Although the influence of wing rotation was strong during the first few chords of translation, averaging the performance over as little as 6.5 chords of motion greatly attenuated the effects of rotation. 5. Together, these modeling results suggest that the unsteady mechanisms generated by simple wing flips could provide an important source for the production of aerodynamic forces in insect flight. Furthermore, the extreme sensitivity to small variations in almost all kinematic parameters could provide a foundation for understanding the aerodynamic mechanisms underlying active flight control.

Suppression of flow-induced pressure oscillations in cavities

Abstract: Experimental methods for suppressing flow-induced pressure oscillations in a shallow cavity resulting from tangential flows over the cavity are described. The effects of manipulating the shear layer over the cavity leading edge are examined. Static and oscillating fences and steady and pulsating flow injection at the leading edge are studied for their effect on cavity sound pressure levels. Both subsonic and supersonic flow conditions are considered. Of the methods tested, static fences at the leading edge were found to provide the most suppression. Suppression was dependent on the frequency mode and the flow Mach number. Nomenclature a = speed of sound D = cavity depth / = frequency k = vortex convection velocity to freestream velocity ratio L = cavity length M = Mach number m = frequency mode number P = pressure P0 = stagnation pressure Re = Reynolds number, Uxlv S = Strouhal number S * = modified Strouhal number U = freestream velocity x = distance from nozzle exit Z = cavity span, lateral dimension a = phase delay parameter y = ratio of specific heats 8 = boundary-layer thickness v = kinematic viscosity

Observations of the Sea-Breeze Front during CaPE. Part I: Single-Doppler, Satellite, and Cloud Photogrammetry Analysis

Abstract: An analysis of two sea-breeze events on 6 August (an onshore flow event) and 12 August (an offshore flow event) 1991 is presented using single-Doppler observations, satellite images, and cloud pictures collected during the Convection and Precipitation/Electrification (CaPE) Experiment. Documentation of the alongfrontal variability at the leading edge of the sea-breeze circulation is presented for the first time. The horizontal structure of the front was strongly modulated by the near-perpendicular intersections of horizontal convective rolls developing in the ambient air out ahead of the sea breeze on 12 August. These intersection points also appeared to be preferred locations for cloud development along the front. Horizontal convective rolls were also documented on 6 August; however, their orientation was nearly parallel to the sea-breeze front. As a result, extended sections of these rolls appeared to have merged with the front as it propagated inland rather than having distinct intersection po...

Two-Dimensional Unsteady Leading-Edge Separation on a Pitching Airfoil

Abstract: The initial stages of two-dimension al unsteady leading-edge boundary-layer separation of laminar subsonic flow over a pitching NACA-0012 airfoil have been studied numerically at Reynolds number (based on airfoil chord length) Rec = 104, Mach number Mx = 0.2, and nondimensional pitch rate H£ = 0.2. Computations have been performed using two separate algorithms for the compressible laminar Navier-Stokes equations. The first method, denoted the structured grid algorithm, utilizes a structured, boundary-fitted C grid and employs the implicit approximate-factorization algorithm of Beam and Warming. The second method, denoted the unstructured grid algorithm, utilizes an unstructured grid of triangles and employs the flux-difference splitting method of Roe and a discrete representation of Gauss' theorem for the in viscid and viscous terms, respectively. Both algorithms are second-order accurate in space and time and have been extensively validated through comparison with analytical and previous numerical results for a variety of problems. The results show the emergence of a primary clockwise-rotating recirculating region near the leading edge which can be traced to a pair of critical points (a center and a saddle) that appear within the flowfield, followed by a secondary counterclockwise-rotating recirculating region and a tertiary clockwise-rotating recirculating region. The primary and secondary recirculating regions interact with each other to give rise to the unsteady separation ("breakaway") of the boundary layer.

Cooled airfoils for a gas turbine engine

Abstract: An internally cooled airfoil for a gas turbine engine includes a leading edge region, a mid-chord region, a transition region, and a trailing edge region. The mid-chord region has a double wall configuration with the two outer walls and two inner walls. The trailing edge region has a conventional single wall configuration. The transition region has a three wall configuration designed to provide a gradual transition from the four wall configuration to the two wall configuration and to minimize high stresses therein.

Unsteady vortex structure over a delta wing

Abstract: The structure of the shear layer which emanates from the leading edge of a 76-deg sweep delta wing and forms the primary vortex is investigated numerically. The flow conditions are Mv_ = 0.2, Re = 50,000 and angle of attack of 20.5 deg. Computational results are obtained using a Beam-Warming-t ype algorithm. The existence of a Kelvin-Helmholtz-type instability of the shear layer which emanates from the leading edge of the delta wing is demonstrated. A description is provided of the three-dimensional, unsteady behavior of the smallscale vortices associated with this instability. The numerical results are compared qualitatively with experimental flow visualizations exhibiting a similar behavior.

Flow directing assembly for the compression section of a rotary machine

Abstract: An inner and/or outer cylindrical wall limiting the working fluid flow path of an axial compressor radially has an ondulating contour. At the intersection with the leading edge of an airfoil the wall shows a convex contour (54) followed by a concave contour (58) in the region of the airfoils maximum thickness while at the intersection with the trailing edge of the airfoil the contour (56) is convex again. The airfoil can either be a rotor blade or a stator vane.

Leading edge shape for flat plate boundary layer studies

Abstract: In experimental boundary layer studies, a flat plate with some shaped nose piece is generally used; this is often prone to flow separation at the junction. By analysing the development of a laminar boundary layer on a two-parameter family of nose shapes, it is found that a cubic super-ellipse of axis ratio 6 or higher is a reasonable optimum shape for avoiding separation on or due to such nose-pieces.

Gas turbine vane with enhanced cooling

Abstract: A gas turbine stationary vane having an airfoil portion and inner (36) and outer (38) shrouds. Five serpentine radially extending cooling air passages (51-55) are formed in the vane airfoil. The first passage (51) is disposed adjacent the leading edge (40) of the airfoil and the second passage is disposed adjacent the trailing edge. A first portion (80) of the cooling air enters the first passage (51), from which it flows sequentially to the second, third, fourth and fifth passages. Additional cooling air (83) enters the third passage directly, thereby bypassing the first and second passages and preventing over heating of the cooling air by the time it reaches the fifth passage. A radial tube (45) extends through the second passage and directs cooling air through the airfoil, with essentially no rise in temperature, to an interstage cavity (70) for disc (42, 43) cooling. Fins project (60-64) into each of the passages and serve to increase the effectiveness and flow rate of the cooling air. The fins in the first and fifth passages are angled so as to direct the cooling air toward the leading (40) and trailing (41) edges, respectively. In addition, the fins in the second through fifth passages are angled to retard flow separation as the cooling air turns 180° from one passage to the next.

Leading edge slat/wing combination

Abstract: A leading edge slat/wing combination (100) where the slat (104) is mounted to a pair of circularly curved carrier tracks (122) positioned within the outer surface contour of the fixed wing (102). In moving from the cruise configuration to the take-off and climb configuration and thus to the fully deployed landing configuration, the slat (104) has a fixed angular orientation relative to the carrier track (122). Also, in two embodiments, the slat, in the take-off and climb configuration, forms a gap or slot with the fixed wing. The arrangement of the present invention is produced by a design technique of identifying design envelopes (209, 210, 215) for certain elements of the slat (104) and fixed wing (102), and then properly matching the configuration and movement of the components within these design envelopes to arrive at an optimized configuration where the curved carrier tracks (122) can be used while being fixedly attached to the slat (104).

Unsteady Wake Over a Linear Turbine Blade Cascade With Air and CO2 Film Injection: Part II—Effect on Film Effectiveness and Heat Transfer Distributions

Abstract: The effect of unsteady wake flow and air (D.R. = 0.97) or CO 2 (D.R. = 1.48) film injection on blade film effectiveness and heat transfer distributions was experimentally determined. A spoked wheel type wake generator produced the unsteady wake. Experiments were performed on a five-airfoil linear cascade in a low-speed wind tunnel at the chord Reynolds number of 3 × 10 5 for the no wake case and at the wake Strouhal numbers of 0.1 and 0.3. A model turbine blade with several rows of film holes on its leading edge, and pressure and suction surfaces (− 0.2 < X/C < 0.4) was used. Results show that the blowing ratios of 1.2 and 0.8 provide the best film effectiveness over most of the blade surface for CO 2 and air injections, respectively

Unsteady Wake Over a Linear Turbine Blade Cascade With Air and CO2 Film Injection: Part I—Effect on Heat Transfer Coefficients

Abstract: The effect of unsteady wake flow and air (D.R. = 1.0) or CO 2 (D.R. = 1.52) film injection on blade heat transfer coefficients was experimentally determined. a spoked wheel-type wake generator produced the unsteady wake. Experiments were performed on a five-airfoil linear cascade in a low-speed wind tunnel at the chord Reynolds number of 3 × 10 5 for the no-wake case and at the wake Strouhal numbers of 0.1 and 0.3. Results from a blade with three rows of film holes in the leading edge region and two rows each on the pressure and suction surfaces show that the Nusselt numbers are much higher than those for the blade without film holes. On a large portion of the blade, the Nusselt numbers «without wake but with film injection» are much higher than for «with wake but no film holes»

Suppression of dynamic-stall vortices over pitching airfoils by leading-edge suction

Abstract: Experiments to control the dynamic-stall vortex (DSV) over the suction surface of a two-dimensional NACA 0012 airfoil, undergoing a hold-pitch-hold motion, are described. Measurements were performed over a range of Reynolds number (3.0×10 4 ≤Re c ≤1.18×10 5 ) and pitch rate (0.072≤α + ≤0.31), using leading-edge suction duing a prescribed period of the airfoil motion. This strategy to manage the DSV, using controlled leading-edge suction, was developed from a study of the mechanisms responsible for the evolution of the vortex. The results indicate that formation of the DSV can be suppressed by removing an appropriate amount of the reverse-flowing fluid to prevent its accumulation in the near-leading-edge region, thereby preventing lift up of the shear layer

Turbine shroud segment with serpentine cooling channels

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 and a trailing edge serpentine channel. Both serpentine channels include ducts that extend to the serpentine channels from a point inward of adjacent retaining hooks. Cooling fluid flowing onto the outward surface of the segments flows through the ducts and along the serpentine channels to cool the leading and trailing edges of the segments.

Steam turbine blade

Abstract: A steam turbine blade having an airfoil portion and root portion by which the blade is affixed to a rotor. The geometry of the blade airfoil is configured to minimize energy loss through the row of blades and reduce the weight of the airfoil. The airfoil has a leading edge and a trailing edge defining a chord therebetween. The chord is reduced linearly from the base of the airfoil to 50% of the airfoil height. However, the chord remains essentially constant from 50% of the airfoil height to the airfoil tip. The root is fir tree shaped and has four sets of tangs and grooves that are configured to minimize the stresses in the root.

Turbine blade having tip slot

Abstract: A gas turbine engine rotor blade includes an airfoil having first and second sidewalls joined together at leading and trailing edges. First and second tip walls extend from adjacent the leading edge along the respective first and second sidewalls to adjacent the trailing edge and are spaced apart to define a tip cavity therebetween. A first notch is disposed in the first tip wall adjacent to the leading edge for channeling into the tip cavity a portion of combustion gases flowable over the airfoil for reducing the heating effect of the gases on the blade tip. In a preferred embodiment, a second notch is disposed adjacent to the trailing edge for promoting flow through the tip cavity from the first notch.

A nonlinear boundary element method for the analysis of unsteady propeller sheet cavitation

Abstract: The unsteady flow around a cavitating marine propeller is treated in nonlinear theory by employing a low-order potential-based boundary element method and a time-marching scheme. The kinematic and dynamic boundary conditions, which are fully three- dimensional and time-dependent, are satisfied on the propeller surface beneath the cavity and on the portion of the blade wake surface which is overlapped by the cavity. The formulation and algorithm are developed to treat arbitrary cavity planforms in an efficient and robust manner. The results from the numerical method are shown to converge quickly with number of panels and number of time steps per propeller revolution. The produced cavity shapes are validated and shown to satisfy the imposed dynamic boundary condition within acceptable accuracy. Computed cavity planforms are compared to those from linear theory and linear theory with leading edge corrections.

Reduction of blade-vortex interaction noise through porous leading edge

Abstract: The effect of the porous leading edge of an airfoil on the blade-vortex interaction noise, which dominates the far-field acoustic spectrum of the helicopter, is investigated The thin-layer Navier-Stokes equations are solved with a high-order upwind-biased scheme and a multizonal grid system The Baldwin-Lomax turbulence model is modified for considering transpiration on the surface The amplitudes of the propagating acoustic wave in the near field are calculated directly from the computation The porosity effect on the surface is modeled in two ways: (1) imposition of prescribed transpiration velocity distribution and (2) calculation of transpiration velocity distribution by Darcy's law Results show leading-edge transpiration can suppress pressure fluctuations at the leading edge during blade-vortex interaction and consequently reduce the amplitude of propagating noise by 30% at a maximum in the near field

Composite airfoil leading edge protection

Abstract: Methods for improving the erosion resistance of composite airfoils are disclosed as are the resultant structures. Wire mesh materials are coated with an erosion-resistant coating, formed to the shape of the airfoil leading edge, and molded into the leading edge during airfoil fabrication.

System and method for determining airfoil characteristics from coordinate measuring machine probe center data

Abstract: The present invention discloses a system and method for determining airfoil characteristics from coordinate measuring machine probe center data. A coordinate measuring machine measures external coordinates along an airfoil and sends the coordinates to a computer. The computer sorts and orders the coordinates into a plurality of probe center points along a plurality of sections along the airfoil. The probe center points are triangulated to detect and eliminate errant points. Then the probe center points are correlated to a nominal part of the airfoil stored in the memory of the computer until a specified plane of interest is obtained. A discrete inset operation is performed on the correlated data for the plane of interest until a maximum thickness is obtained. The inset points are then joined to form an airfoil meanline, which is used to determine characteristics such as leading edge thickness, trailing edge thickness, and chord length.

Air bearing slider having projections that provide minimal actual contact area

Abstract: A slider for providing an air bearing between a transducer and a moving recording surface includes a platform having a first surface with a leading edge opposite a trailing edge and a lateral axis between the leading and trailing edges. A plurality of projections extend from the first surface. At least one of the projections is disposed between the leading edge and the lateral axis and another of the projections is disposed between the trailing edge and the lateral axis. Each of the projections includes a second surface and a recess facing the leading edge. The recesses are bounded on their sides by shrouds that limit air bearing leakage. The recesses provide lift and stability to the slider when the recording surface is moving and always remain out of contact with the recording surface. To limit stiction, only the second surface of the projections is in apparent contact with the recording surface when the recording surface is stationary.

Gas turbine airfoil clocking

Abstract: The first stage of vanes (16) and second stage of vanes (24) each contain the same number of vanes. The second stage of vanes are located such that the wake flow (38) from the first stage of vanes falls on or near the leading edge, after passing through the stage of rotating blades.

Aerodynamic entrainment threshold: effects of boundary layer flow conditions

Abstract: Experimental data are presented demonstrating the influence of boundary layer flow conditions on aerodynamic entrainment of grains in the absence of intersaltation collisions. New methods are proposed for (1) the unambiguous determination of aerodynamic threshold for any grain population and (2) approximation of the probability density function (PDF) distributions of threshold shear velocity for aerodynamic entrainment. In wind tunnel experiments, the orderly spatial development of flow conditions within a developing boundary layer over the roughened surface of a flat plate constrains the aerodynamic threshold condition in terms of both mean and fluctuating values. Initial grain dislodgements and subsequent erosion from narrow strips of loose, finely fractionated ballotini were recorded photographically as wind speed was increased. Boundary layer parameters, including average threshold shear velocity (U*t), were calculated using the momentum integral method. Direct observations show that sporadic oscillation of grains preceded dislodgement. At slightly higher velocities most grains rolled over their neighbours before entering saltation. Initial entrainment in spatially semi-organized flurries of 50 or more grains was followed by quiescent periods at airflow velocities close to threshold. These observations provide strong circumstantial evidence linking both the nature and spatial pattern of initial grain motions to sweep events during the fluid bursting process. For each grain fraction, values of U*t were found to span an unexpectedly wide range and to decrease downwind from the leading edge of the plate as turbulence intensity increased. A probabilistic entrainment model is applied to the aerodynamic threshold condition so as to incorporate the effects of changing turbulent flow regimes over the plate. Analysis of strip erosion curves gives both an objective definition of the threshold condition and usable approximations of the PDF for U*t required by the model and for future stochastic treatment of the threshold condition.

The influence of airfoil thickness and Reynolds number on separation

Abstract: A theoretical description is sought for the mechanisms whereby the attached motion past a finite flat plate is gradually transformed into the quite different grossly separated motion past a bluff body, as the body's thickness increases from zero to a finite value at high Reynolds numbers. The theory is based largely on the free stream line inviscid approach allied to the viscous triple-deck requirement at separation, for laminar symmetric flow of an incompressible fluid. Some surprising and potentially significant phenomena are found to arise as the body thickness increases; and the geometries of the trailing and leading edges have a crucial effect on the occurrence and position of separation. The main properties found are: the match with the Kirchhoff theoryat one extreme; at the other extreme, the match with the triple-deck account for wedged trailing edges; in-between, a “cut-off” stage during which the wake dimensions and the drag abruptly increase; and, also in-between, a nonlinear stage in whichleading edge separation can occur, depending delicately on the precise shape of the leading edge. These properties arise from the strong viscous effect at separation. Again, whether the closure of the recirculatory flow region is blunt or not, the theory is consistent with the results of earlier work for the above extremes.Nonuniqueness/bifurcation of the separated flow solutions is also found to arise for certain profiles.

Integral airfoil total temperature sensor

Abstract: An air data sensing probe adapted for mounting to an aerodynamically-shaped airfoil or strut attached to an air vehicle. An inlet port located near the leading edge of the strut admits fluid, to a first cavity and then, in turn, to a second cavity, so that the total temperature of the fluid may be measured and a signal related thereto conveyed to suitable flight control gear. A first exhaust port located generally opposite the inlet port allows entrained particles to exit the probe and boundary layer fluid evacuation apertures formed through the strut across the interface between the two cavities permit only a substantially particle-free core sample of fluid on the temperature sensing element. The secondary cavity couples to the primary cavity at an angle so that inertial separation of entrained particles results. In a second embodiment, the air data sensing probe connects to a barrel-shaped probe head so that multiple parameters related to the fluid can be measured, such as total pressure, Pt, static pressure, Ps, and total temperature, Tt, of a fluid flowing relative to the air vehicle as well as angle-of-attack (AOA) of the air vehicle relative to the fluid.

Wing folding and free-flight kinematics in Coleoptera (Insecta): a comparative study

Abstract: Relative movements of the main wing areas around the major flexion lines are compared during wing folding at rest, and during the supinatory phase of the flight cycle, which involves considerable wing deformation. Folding of the wing apex at rest is achieved by a combination of movements around the median flexion line (the main longitudinal flexion line), the principal transverse fold, and a variety of smaller, oblique ‘tucking’ folds. During flight, wing tip deformation is strongly influenced by elastic forces involved in the normal wing folding and unfolding processes. Those beetles possessing an inwardly sprung wing apex display partial folding at supination, associated with the temporary relaxation of the forces opposing spring recoil. These beetles also show enhanced mobility about the median flexion line which facilitates leading edge supination. The presence of the principal transverse fold may help to concentrate ventral flexure towards the wing tip. The wings of beetles possessing an outwardly-sprung apex are much less affected by the presence of the flexion lines associated with folding. In these cases, enhanced supination of the leading edge, in the face of an overall increase in wing membrane stiffness, may be related to the presence of the highly-sclerotized pterostigma.

Engineered ceramic components for the leading edge of a helicopter rotor blade

Abstract: An engineered ceramic component (10) for the leading edge of a rotor blade provides enhanced erosion protection therefor. In one embodiment, the engineered ceramic component (10) includes a strain isolator member (19), an aerodynamic ceramic member (14), a first adhesive bond layer (16), and a second adhesive bond layer (18). The strain isolator member (19), which is operative to minimize strain transfer between the rotor blade infrastructure and the aerodynamic ceramic member (14), is configured so that inner mold line surface thereof is complementary to outer mold line surface of the rotor blade infrastructure. The aerodynamic ceramic member (14), which is operative to provide enhanced erosion protection for the respective leading edge of the rotor blade, is configured so that the outer mold line surface thereof defines the aerodynamic configuration of the respective leading edge and the inner mold line surface is complementary to the outer mold line surface of the strain isolator member (19). The aerodynamic ceramic member (14) is fabricated from an engineered ceramic material, which are defined in the present context as ceramic materials having a hardness greater than 1200kg/mm2 and an average flexural strength equal to or greater than 40ksi. The first adhesive bond layer (16) is operative to bond the strain isolator (19) member to the rotor blade infrastructure (140) and the second adhesive bond layer (18) is operative to bond the aerodynamic ceramic member (14) to the strain isolator member (19). In another embodiment, the engineered ceramic member (10) includes the aerodynamic ceramic member (14) described hereinabove and a thick adhesive layer. The thick adhesive layer is operative to minimize strain transfer between the rotor blade infrastructure and the aerodynamic ceramic member and to bond the aerodynamic ceramic member in combination with the rotor blade infrastructure.

Fan shroud aspirator

Abstract: An aspirator is mounted on a fan shroud surrounding a vehicle cooling fan. The aspirator has a housing which forms a duct which connects an opening to a port for connecting to a vehicle function. The aspirator is spaced apart from the fan and upstream from the fan and extends at an acute angle with respect to a tangent to a cylindrical wall of the shroud. The aspirator has a near edge which faces generally opposite to a direction of rotation of the fan blades and which is parallel to a leading edge of a fan blade when the leading edge is spaced apart from the near edge by a distance which is slightly larger than the width of the fan blade. The aspirator has a flange which projects parallel to a plane of rotation of the fan and generally in the rotation direction of the fan. The duct has a triangular cross sectional shape with an apex which projects in a direction which is upstream with respect flow of air moved by the fan.

Quantitative study of unsteady compressible flow on an oscillating airfoil

Abstract: Detailed interferometric measurements of the flow near the leading edge of an oscillating airfoil offer the first detailed experimental quantification of the locally compressible flow field that surrounds an oscillating airfoil at moderate subsonic Mach numbers. Interferograms obtained by a specially adapted real-time point-diffraction interferometry technique have revealed significant characteristics of this complex, and very rapidly varying, locally supersonic flow. Instantaneous pressure distributions determined from these interferograms document the effect of unsteadiness on the leading-edge flow environment.