2161 results about "Leading edge" patented technology

An aircraft defining a longitudinal centerline and extending between a forward end and an aft end is provided. The aircraft comprises a fuselage extending between the forward end of the aircraft and the aft end of the aircraft; a wing assembly extending laterally outwardly with respect to the longitudinal centerline from a portion of the fuselage; an unducted turbofan engine including an unducted fan that defines a plurality of fan wakes; and a pylon fairing that connects the unducted turbofan engine to the wing assembly, the pylon fairing defining a pylon fairing dihedral, wherein the pylon fairing dihedral is in a direction opposite the incoming fan flow swirl and associated sheared profiles of fan wakes upon arrival at the pylon leading edge.

An aircraft defining a longitudinal centerline and extending between a forward end and an aft end is provided. The aircraft comprises a fuselage extending between the forward end of the aircraft and the aft end of the aircraft; a wing assembly extending laterally outwardly with respect to the longitudinal centerline from a portion of the fuselage; an unducted turbofan engine including an unducted fan defining a fan diameter; and a pylon fairing that connects the unducted turbofan engine to the wing assembly, the pylon fairing defining a leading edge and including a first inflection point along the leading edge, the first inflection point defining a first radius of curvature that is less than 5 times the fan diameter and greater than 0.1 times the fan diameter.

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

A sheath is provided. The sheath includes a leading edge portion having pressure and suction sides for respective association with pressure and suction sides of an airfoil and first and second wings respectively extending from the pressure and suction sides of the leading edge portion. The first wing includes a first elongate portion and a first trailing edge disposed at an end of and thinned relative to the first elongate portion. The second wing includes a second elongate portion and a second trailing edge disposed at an end of and thinned relative to the second elongate portion.

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

An airfoil is provided that includes a first end, a second end, a leading edge, a trailing edge, a pressure side and a suction side. The leading edge and the trailing edge are joined by the pressure side and the suction side to provide an exterior airfoil surface extending in a spanwise direction from the first end of the airfoil to the second end of the airfoil. The exterior airfoil surface are formed in conformance with a plurality of cross-section profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading edge and the trailing edge at the span location.

An assembly for an aircraft propulsion system includes a propulsor rotor and a guide vane structure. The propulsor rotor is configured to rotate about an axis. The guide vane structure includes a plurality of guide vanes arranged circumferentially about the axis. The guide vane structure is axially next to and downstream of the propulsor rotor. The guide vanes include a first guide vane. The first guide vane includes a camber line, a leading edge, a trailing edge, a leading edge section and a trailing edge section. The first guide vane extends longitudinally along the camber line from the leading edge to the trailing edge. The leading edge section forms the leading edge. The trailing edge section forms the trailing edge. The leading edge section is configured to move longitudinally along the camber line relative to the trailing edge section to change a camber of the first guide vane.

In an aircraft wing structure where space and weight considerations are a principal concern, there may be significant challenges in providing a moveable slat and / or a spoiler, particularly in combination with one another. A control device for an aircraft includes upper and lower flaps pivotably mountable adjacent to a slot in a wing behind a leading edge, wherein the control device is operable in a lift configuration to move the upper and lower flaps towards each other to abut opposing surfaces of the slot, to permit airflow along the slot.

Aircraft turbine engine including two coaxial annular walls defining between them a main annular duct for the flow of a primary airflow; —a rotor blading extending radially through the first duct; —an annular separator arranged downstream of the rotor blading and configured to divide the primary airflow into two to form the secondary air flows; —first variable guide vanes which are distributed around the shaft and each comprise a leading edge which is located upstream of the separator and trailing edges which are located in the secondary airflows; —and fixed guide vanes which are distributed around the shaft in the external airflow and downstream of the first variable vanes.

A work vehicle includes a boom coupled to a machine body, an arm coupled to the boom, a bucket coupled to the arm and including an edge portion to function as a leading edge when the bucket excavates earth, a rotation sensor to measure rotation of a to-be-measured object which is at least one of the boom, the arm, or the bucket, and a controller configured or programmed to control rotation of the to-be-measured object. The controller is configured or programmed to control rotation of the to-be-measured object based on an output value obtained by applying a correction function directly or indirectly to a difference between a target angular velocity for rotation of the to-be-measured object and an actual angular velocity of the to-be-measured object derived based on a measurement result from the rotation sensor, the correction function being based on the actual angular velocity.

An aircraft wing has a leading edge, a fixed wing portion and a leading-edge spoiler device. The spoiler device is hingedly mounted relative to the fixed wing portion adjacent the leading edge, moveable between retracted and deployed positions. In the retracted position a spoiler surface forms a continuous surface with a fixed wing portion upper aerodynamic surface, and in the deployed position the spoiler surface is rotated relative to the upper aerodynamic surface. In the retracted position the spoiler surface has a first portion forward of the hinge and a second portion aft of the hinge. A center of pressure of the local wing section acts on the spoiler surface providing a moment about the hinge in a first direction to bias the spoiler device to the retracted position and providing a moment about the hinge in a second, opposite direction, rotating the spoiler device to the deployed position.

The invention discloses a multi-degree-of-freedom morphing wing based on a zero poisson ratio superstructure, and belongs to the technical field of morphing wings. The wing comprises a front edge section, a wing body and a rear edge section which are sequentially connected, the front edge section comprises an inner side front edge plate, a middle front edge plate and an outer side front edge plate, the wing body comprises a plurality of zero-Poisson-ratio honeycomb structures, each zero-Poisson-ratio honeycomb structure is formed by arraying a plurality of zero-Poisson-ratio unit cell cubic structures, and the adjacent zero-Poisson-ratio unit cell cubic structures share a connecting rod; the rear edge section comprises an inner side rear edge plate, a middle rear edge plate and an outer side rear edge plate; the inner-side front edge plate and the inner-side rear edge plate, the middle front edge plate and the middle rear edge plate, and the outer-side front edge plate and the outer-side rear edge plate are respectively connected through zero-Poisson-ratio honeycomb structures which are arranged along the chord length direction of the wing; the two sides of the middle front edge plate are connected with the inner side front edge plate and the outer side front edge plate through zero-Poisson-ratio honeycomb structures arranged in the extension direction of the wing correspondingly, and the two sides of the middle rear edge plate are connected with the inner side rear edge plate and the outer side rear edge plate through zero-Poisson-ratio honeycomb structures arranged in the extension direction of the wing correspondingly. The wing has four deformation functions of variable elongation, chord length, dihedral angle and camber, so that the aircraft can adapt to different flight environments and tasks.

A modular blended wing body (BWB) aircraft includes common components containing a nose structure having a left aft edge and a right aft edge intersecting on a centerline of the BWB on one end and intersecting on a leading edge of the BWB on another end, and two wing structures attached to the nose structure, wherein the at least two rear structures includes a left-wing structure arranged on the left side of the centerline, wherein the left-wing structure includes a left forward edge, and a right-wing structure arranged on the right side of the centerline, wherein the right-wing structure includes a right forward edge, and wherein the left-wing structure and the right-wing structure are laterally symmetrical to the centerline, and wherein there is no clear demarcation between the at least two wing structures and the nose structure of the plurality of common components of the BWB aircraft.

A stator vane includes a platform that defines an opening. The stator vane further includes an airfoil that has a leading edge, a trailing edge, a suction side wall, and a pressure side wall. The airfoil extends radially between a base and a tip. At least one of the base or the tip includes a protrusion. The protrusion extends into the opening of the platform such that the platform surrounds the protrusion of the airfoil. The stator vane further includes a braze joint disposed between and fixedly coupling the platform and the protrusion of the airfoil. The stator vane further includes a cooling circuit defined in at least one of the protrusion or the platform to cool the braze joint.

A turbine engine includes a compressor section, a combustion section, and a turbine section, and an airfoil with an outer wall defining a pressure side and a suction side and extending between a leading edge and a trailing edge to define a mean camber line. A first thickness is defined between the pressure side and the suction side at a first location along the mean camber line.

The invention belongs to the technical field of three-dimensional design of aero-engine gas compressor composite swept-curved rotor blades, and particularly relates to an aero-engine gas compressor composite swept-curved rotor blade and an integral disc structure thereof, and the front edge of the rotor blade is sequentially provided with a blade root sweepback area A1, a sweepforward area A2 and a blade tip sweepforward area A3 from bottom to top; the blade root sweepback area A1, the sweepforward area A2 and the blade tip sweepforward area A3 account for 0-20%, 20-75% and 75-100% of the radial height range of the rotor blade respectively; the blade root sweepback area A1 and the blade tip sweepforward area A3 are provided with dihedral angles; a blade root sweepback area B1, a sweepback degree area B2 and a blade tip straight area B3 are sequentially arranged on the rear edge of the rotor blade from bottom to top; the blade root sweepback area B1, the sweepback passing area B2 and the blade tip straight area B3 account for 0-30%, 30%-80%, 75%-100% and 80%-100% of the radial height range of the rotor blade respectively; the blade root sweepback area B1 and the blade tip straight area B3 are provided with dihedral angles.

The invention provides a low-scattering carrier for an airfoil front edge RCS test, the low-scattering carrier comprises a front edge cover plate and a carrier main body which are detachably connected, the carrier main body and a carrier end face are of an integrally formed metal hollow structure, the front edge of the carrier main body is of a linear structure attached to an airfoil front edge wave-absorbing structure, and the front edge of the carrier main body is of a rectangular structure attached to the airfoil front edge wave-absorbing structure. The rear edge of the carrier body is of a linear structure parallel to the front edge. According to the low-scattering carrier, the drop-shaped design is adopted, the carrier end face and the carrier body are integrally formed, and the airfoil front edge and rear edge parallel structure is used, so that the influence of the edge on angular domain scattering assessment is reduced, and the vertical polarization and horizontal polarization electromagnetic scattering characteristics of the carrier are greatly reduced; and the stealth performance of the carrier in various radar environments is further improved. The low-scattering carrier provided by the invention has a good traveling wave suppression effect, and in an X wave band 0-50 degrees range, the RCS mean value of polarization in the horizontal direction can reach-45dBsm, and the RCS mean value of polarization in the vertical direction can reach-40dBsm.

The invention relates to a device comprising an asymmetrical wing profile (2) and an adjusting device (15), which asymmetrical wing profile (2) has a profile thickness (q) and a profile chord (p) extending between a wing leading edge (n) and a wing trailing edge (e), which asymmetrical wing profile (2) has, in at least one longitudinal portion (L1-Ln), a three-part variable wing profile (21) in which a front and a rear wing segment (211, 213) are joined to a central wing segment (212) in an articulated manner by means of two hinges (214) having axes of rotation (z) and are designed to permit a rotational movement about the associated axis of rotation (z) with respect to the central wing segment (212).

The invention relates to a propulsion unit which comprises: - an outer casing (102); - a hub; - a pusher propeller; and - a fixed flow straightener comprising at least one stator blade (114) with variable pitch about a pitch axis (Y'), each variable-pitch stator blade (114) having a pitch angle defined as the angle between any plane perpendicular to the main axis (X) and a chord line (706) connecting a leading edge of the stator blade (114) to a trailing edge of the stator blade (114), this chord line (706) being taken at 75% of an outer radius of the stator blade (114) with respect to the main axis (X). The leading edge of at least one stator blade (114) has a dihedral angle (D'BA') at 95% of the outer radius with an absolute value greater than 1°, preferably greater than 3°, more preferably greater than 10°, for at least one value of the pitch angle within the range [70°; 90°].

A wing defines a chord which runs from a leading edge to a trailing edge. The wing has a main portion and a flap which is movable relative to the main portion to change the aerodynamics of the wing, for instance to maximize lift, or to reduce drag at the expense of some loss of lift. The wing has a guide mechanism which guides movement of the flap. The guide mechanism uses a pair of runners which run along respective tracks. The tracks are positioned at different locations along the direction of the chord, but overlap one another in the direction of the chord.

A composite material combined layer wall plate structure suitable for combined connection of wings comprises a main box section skin and a plurality of reinforcing ribs which are integrally cured and formed, the front edge and the rear edge of the main box section skin are sequentially provided with recesses, and a front sunken edge for lap joint of the front edge skin and a rear sunken edge for lap joint of the rear edge skin are formed; a front beam reinforcing rib used for being connected with a front spar web is fixedly arranged at the bottom of the transition area of the front sunken edge and the main box section skin body, and a rear beam reinforcing rib used for being connected with a rear spar web is fixedly arranged at the bottom of the transition area of the rear sunken edge and the main box section skin body. The connecting mode of the wall plate and a related structure is simplified, the number of parts is reduced, the assembling efficiency can be improved, and dismounting, replacing and maintaining of an airplane in the using and maintaining process are facilitated.

An assembly is provided for an aircraft propulsion system. This assembly includes a nacelle inlet structure which extends axially along and circumferentially around a centerline. The nacelle inlet structure includes an inlet lip, an inner barrel, an outer barrel and a plurality of structure segments. The inlet lip forms a leading edge of the nacelle inlet structure. The inner barrel projects axially aft away from the inlet lip. The outer barrel projects axially aft away from the inlet lip. The outer barrel is radially outboard of and axially overlaps the inner barrel. Each of the structure segments includes an exterior skin, a mount and an electric heater configured to heat the exterior skin. The exterior skin forms a respective circumferential section of the inlet lip and a respective circumferential section of the outer barrel. The mount is bonded to the exterior skin and mechanically attached to the inner barrel.

An aircraft wing includes a wing root having a leading edge, a trailing edge, and a distal end. The wing also includes a wing tip and a wing-tip hinge that pivotally couples the wing tip to the wing root and defines a wing-tip axis that is perpendicular to a folding plane. The wing tip is pivotable about the wing-tip axis, within the folding plane, between, and inclusive of, an extended position and a retracted position, relative to the wing root. When in the extended position, the leading edge of the wing tip extends from the leading edge of the wing root in a spanwise direction of the wing root. When in the retracted position, the leading edge of the wing tip extends from the leading edge of the wing root in a chordwise direction of the wing root.

A leading-edge device for an aerodynamic component is disclosed having a first skin element having an inner surface and an aerodynamic outer surface, and further includes a second skin element having an inner surface and an aerodynamic outer surface. The leading-edge device includes a first coupling element protruding from the inner surface of the first skin element and a second coupling element protruding from the inner surface of the second skin element. The first and second coupling elements are coupled to each other in a coupling region. The first skin element and the first coupling element are formed as a single-piece structure.

The invention provides a displacement vortex generator and a position regulation and control method, and belongs to the field of active flow control equipment. The VG comprises a belt, a belt guide rail, a belt bearing, a sliding block, a supporting column, a triangular VG body, a guide rail supporting block, a limiting groove, a body connecting plate and a motor. The belt is arranged on the belt guide rail and is driven by the motor; the sliding block is arranged on the belt, is in an inverted concave shape and is connected with the body connecting plate through the supporting column, and the VG bodies are fixed to the two ends of the body connecting plate. The belt guide rail is arranged on the guide rail supporting block, and the sliding block can penetrate through the limiting groove. The motor is controlled to drive the belt to drive the sliding block and the VG body to move, and accurate adjustment of the position of the vortex generator body is achieved. The position of the vortex generator is automatically adjusted according to changes of the incoming flow wind speed and the attack angle, the optimal pneumatic effect is kept under different working conditions, the pneumatic performance and the power generation efficiency of the wind turbine blade are remarkably improved, and the negative influence of the front edge roughness can be effectively weakened.

The invention discloses a method for quickly generating a compressor blade profile containing local geometric deviation. The method comprises the following steps: determining a mean camber line point and thickness distribution according to geometric data of a suction surface and a pressure surface; determining parameterized expression of the mean camber line geometry according to the camber line points, and obtaining parameter expression of a thickness distribution curve according to the thickness distribution; generating pressure surface geometric points and suction surface geometric points containing local geometric deviation based on parameterized expression of mean camber line geometry and parameterized expression of the thickness distribution curve; a blade profile leading edge curve and a blade profile trailing edge curve are generated; and integrating the curve data to generate the blade profile containing the local geometric deviation. The blade obtained on the basis of the method contains the local geometric deviation, control over the machining deviation of the blade is facilitated, and the method has great significance in determining the influence of the typical geometric deviation of the blade on the aerodynamic characteristics of the gas compressor through simulation calculation.

A propeller. The blades of the propeller are characterized by a blade center axis which corresponds to a generator line of a cone to which the blade, or blade thrust surface, conforms. The propeller hub is fixed to the blade at the root, in line with the blade center axis, such that the hub axis and blade center axis lie in the same plane, and the leading edge of the blade is positioned forward (referring to the upstream direction of movement caused by the propeller) of the trailing edge of the blade.

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

An aircraft includes a wing and a leading-edge aerodynamic surface coupled to the wing. An actuator is disposed within the wing and coupled to the leading-edge aerodynamic surface via a linkage system. Two or more hinge systems are coupled to the leading-edge aerodynamic surface and configured to rotate about a fixed axis internal to the wing to move the leading-edge aerodynamic surface between a retracted position and a drooped position.

The invention discloses a hybrid honeycomb flexible skin of a variable sweepback wing, which belongs to the field of flexible structure design of a morphing aircraft, and comprises a frame, a deformable support structure and a flexible surface layer, the frame is divided into a parallelogram area and a trapezoidal area, and the corresponding deformable supporting structures are also divided into two types; the first rod, the second rod, the third rod and the fourth rod are rods with the same length, and the fifth rod is a nested structure rod with the variable length. The four rods are flexible skin front edges, and the five rods are flexible skin rear edges; the AC section of the first rod is as long as the BD section of the second rod; the AC section of the first rod, the BD section of the second rod, the fourth rod and the third rod form the edge of a parallelogram area; the CE section of the first rod, the DF section of the second rod, the third rod and the fifth rod form the edge of the trapezoidal area. According to the method, the flexible skin is divided into areas, different deformable supporting structures are used for filling according to the deformation characteristics of all the areas, and the bidirectional corrugated honeycombs meet the requirement for edge length change in the sweepback changing process.