839 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 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 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.

Provided are computer-implemented methods for autonomously controlling an aircraft with vertical take-off and landing capabilities and folding wings that includes controlling a plurality of thrust producing components of an aircraft to cause the aircraft to rise vertically when wings of the aircraft are in a first folded configuration, where when the wings of the aircraft are in the first folded configuration, a leading edge of each wing is oriented in a vertical direction setting motor controller gains based on the wings of the aircraft being in the first folded configuration, and causing the aircraft to align with a direction of airflow when the wings of the aircraft are in the first folded configuration, and controlling thrust producing components and control surfaces and internal articulation mechanisms of the aircraft to cause the aircraft to transition from folded wing configuration to unfolded wing configuration. Systems and computer program products are also provided.

A tonneau system for a vehicle including a cargo bed comprises a frame including a leading edge and a trailing edge. A cover is coupled to the frame and movable between a deployed position and a retracted position. A mechanism is coupled to the frame and operable to move the frame and the cover between a home position, an aerodynamic position, and an extended box position. When at the home position, the frame extends substantially parallel to the ground at a first elevation. The leading edge of the frame is positioned at a second elevation higher than the first elevation and the trailing edge remains at the first elevation when the frame is at the aerodynamic position. The leading edge and the trailing edge of the frame are positioned at the second elevation when at the extended box position.

An aircraft wing assembly includes a wing box including an upper surface, a lower surface, and a leading edge structure connecting the upper surface to the lower surface and secured to the upper surface at a connection point. The aircraft wing assembly includes a drive mechanism secured to the wing box and a high lift device mounted to the drive mechanism. The drive mechanism is configured to move the high lift device between a retracted position and an extended position in use, and wherein the high lift device is configured to overlap the connection point when in the retracted position.

A structure is proposed for traversing a fluid environment and reducing the interference of the fluid flow vector at a trailing edge of the structure. The structure comprises an elongate body having a root, a wingtip, a leading edge and the trailing edge. The structure has a plurality of flexible projections positioned along the trailing edge extending from a base to a tip. The plurality of flexible projections bend to conform to a curvature of the fluid flow vector at the trailing edge and have a first resonant frequency above an oscillating force generated by the fluid flow vector.

The invention relates to the technical field of aero-engine aerodynamic optimization and computer aided design, and discloses a response surface modeling-based aero-engine low-pressure turbine blade leading edge model numerical optimization design method and structure. According to the method, the wake injection angle is changed by adjusting the flow or the rotating speed of a rotor, accurate matching of Klebanoff stripes and large-scale vortex generation opportunities is achieved, stripe impact is promoted, and large-scale vortex fracture is accelerated; and meanwhile, by constructing a blade front edge modeling parameterization model and establishing a response surface agent model of a modeling design variable parameter combination and a boundary layer velocity mean square error, an optimal design variable parameter combination is determined in a constraint domain by taking maximization of the velocity mean square error in a boundary layer as a target. On the premise that the blade profile load is not changed, the momentum thickness and the total pressure loss of the boundary layer are effectively reduced, the design efficiency and the engineering applicability under the working condition of the low Reynolds number are improved, an additional flow control structure does not need to be introduced, and the structure is simple and easy to implement.

The invention relates to a centrifugal fan and a range hood. The centrifugal fan comprises a volute and an impeller. A current collector is arranged at the air inlet of the volute; the first disc and the second disc are oppositely arranged in the front-back direction; the blades are connected between the first disc and the second disc and distributed in the circumferential direction, the front edge of at least one blade is provided with a wave-shaped section, and the length H2 of the wave-shaped section in the axis direction of the impeller meets the conditions that 0.8 Dn < = D1 < = Dn, 0.5 < = Hz / Hq < = 2; wherein D1 is the maximum diameter of an inlet of the current collector, Dn is the minimum inner diameter of the impeller, Hz is the axial depth of the impeller, and Hq is the air inlet space depth of the front side or the rear side of the centrifugal fan in the machine shell. A plurality of parameters such as the axial length size of the wave-shaped section, the inlet size of the current collector, the inner diameter of the impeller, the depth direction size of the impeller on the single side and the depth of the air inlet space on the side are subjected to collaborative design synthesis and matching. The invention has the advantages that the pneumatic performance of the machine can be ensured while the airflow noise is obviously reduced.

The invention relates to the technical field of aircraft propulsion systems, in particular to a special-shaped air inlet lip for an on-wing integrated distributed propulsion short cabin unit. The lip is located at the starting part of the air inlet section and is defined by an upper lip, a lower lip, a left lip and a right lip; the inner surface of the upper lip is arc-shaped; the inner surface of the lower lip is linear and is smoothly fused with the upper surface of the wing; the inner surfaces of the left and right lips are curved surfaces with curved sweep characteristics, and the lower half parts of the left and right lips are in an inward-contracting shape; the length of the air inlet channel is defined as the axial distance from the foremost edge of the upper lip to the section of the inlet of the ducted fan and is not greater than the fan diameter D; a first rectification area and a second rectification area are arranged in the air inlet channel to optimize a flow field. Through the fusion design of the special-shaped lip and the wing and the compact air inlet channel, the requirement for compact integrated layout on the wing is met, meanwhile, the additional resistance of the nacelle is effectively reduced, the total pressure recovery coefficient not lower than 0.9 and the flow coefficient larger than 0.8 are guaranteed, and uniform and stable high-quality air inlet is provided for the ducted fan.

The invention discloses an aeroacoustic wind tunnel corner noise elimination flow deflector, a flow deflector group and a design method, and belongs to the technical field of aeroacoustic wind tunnel test equipment. The corner flow deflector aims at solving the technical problems that an existing corner flow deflector cannot give consideration to aerodynamic performance and noise elimination performance, and a special shape design method and a combination application scheme are lacked. The noise elimination flow deflector is of an integrated structure formed by sequentially connecting a semi-oval front edge, a front straight section, an arc section, a rear straight section and a wedge-shaped tail edge, and key parameters of the upper profile and the lower profile are designed in an associated mode. The flow deflector set is formed by combining and installing at least two noise elimination flow deflectors according to a specific rule, and a uniform-section airflow channel is formed. According to the design method, coupling optimization of the noise elimination amount and the aerodynamic performance is achieved through a parameterized design process in combination with the Picrov formula, and the design method has the advantages of being high in noise elimination efficiency, small in aerodynamic loss and good in airflow uniformity, is suitable for a symmetric corner aerodynamic acoustic wind tunnel and can be widely applied to aerodynamic noise tests in the fields of aviation, automobiles and the like.

The invention discloses a rotor blade tip gap design method based on cold and hot state transformation and an aero-engine. The design method comprises the steps that the design working condition of cold and hot state transformation of a fan rotor is determined; performing pneumatic calculation on fan components; designing the structure of the hot-state blade S1; reversely solving the hot-state blade S1 to obtain a cold-state blade S2; designing a cold-state blade structure; performing forward solution on the cold-state blade to obtain a hot-state blade S3; forward solving is carried out on the fan rotor cold-state blade S2 to obtain a corresponding hot-state blade S3; performing error analysis on the thermal-state blade S3 and the thermal-state blade S1; blade tip models of the cold-state blade S2 and the hot-state blade S1 are compared and analyzed, and the front edge radial deformation delta q and the tail edge radial deformation delta h of the fan rotor are obtained; and according to the designed thermal state gap value delta r, the front edge gap Sq and the tail edge gap Sh of the blade tip of the fan rotor blade are obtained.

An aeronautical propulsion unit includes a central engine, an upstream series and a downstream series of blades, the blades of at least one of these series being adapted to be driven in rotation about the central longitudinal axis by a turbine, and a nacelle which encloses the central engine. At least some of the blades of the two series are variable-pitch blades. At least one of the blade of the first series of blades has a trailing edge having serrations, and / or at least one of the blades of the second series of blades has a leading edge having serrations.

The invention relates to the technical field of cross-flow wind wheels, in particular to a cross-flow fan blade tail edge sawtooth structure and a cross-flow wind wheel applying the cross-flow fan blade tail edge sawtooth structure. The sawtooth structure comprises a fan blade body, and the peripheral contour of the fan blade body is defined by a fan blade root edge, a fan blade tail edge, a fan blade top edge and a fan blade front edge; the fan blade main body is provided with a pressure surface and a suction surface, the pressure surface and the suction surface are oppositely arranged, the pressure surface is concavely arranged towards the suction surface, so that the suction surface is convexly arranged, and the section shape of the pressure surface, the section shape of the suction surface, the fan blade root edge and the fan blade top edge are arranged in an arc shape; a plurality of notches are formed in the tail edge of the fan blade in an inwards-concave mode, and the adjacent notches are different. By means of the innovative structural design, noise is effectively reduced, and meanwhile it is guaranteed that the core aerodynamic performance of the fan blade is stable.

The invention discloses a preparation method of a bionic vibration reduction structure of an aircraft wing leading edge, relates to the technical field of aircraft wings, and completes establishment of a bionic kangaroo jumping structure, a bionic origami honeycomb structure and a bionic woodpecker spiral structure based on three-dimensional modeling software. A bionic kangaroo jumping structure, a bionic origami honeycomb structure and a bionic woodpecker spiral structure are regionally filled into a traditional aircraft wing entity model in Magics software, and based on a laser powder bed melting technology, nickel-titanium alloy powder is used for printing and forming the filled traditional aircraft wing entity model; and preparing a PU-MXene mixed solution, and plating the PU-MXene mixed solution on the surface of the printed and formed solid model of the wing of the traditional aircraft through an electrochemical deposition method. The effects of damping the leading edge of the aircraft wing, improving the rigidity, absorbing sound, reducing noise and being super-hydrophobic are achieved.

The present invention provides an extreme STOL airplane comprising: a fuselage and a wing having port and starboard wing sections. Each wing section has a body with top and bottom surfaces, leading and trailing edges, and a span, i.e., the distance between a wing section tip and root connected to the fuselage, and flap coves. Aerodynamic elements integrated into the wing sections include Fowler flaps, having leading and trailing edges and a span, flap tracks, wherein the flap tracks are external to the wing section body, extend aftward beyond the trailing edge of the wing section, and are configured to enable the Fowler flaps to rotate or deflect and to translate or extend and retract, Frise ailerons, wherein the Frise ailerons are located outboard of the Fowler flaps, and spoilerons located over the leading edge of the Fowler flaps when the flaps are in the fully extended position.

Variable pitch fan blade (8a) for an unfaired turbomachine fan, the blade (8a) comprising an aerodynamic blade (9) and a base (10), the blade (9) comprising a recess opening simultaneously onto the upper and lower surfaces (14) and a leading edge (15) of the blade (9), the blade (8a) further comprising a measuring device (26) housed in the recess and clipped onto the blade (9), the measuring device (26) comprising a streamlined body (28) whose outer surface (29) is complementary to the outer surface (30) of the blade (9) to ensure aerodynamic continuity, the body (28) carrying at least one measuring instrument (31) configured to measure at least one parameter of an airflow upstream of the blade (9). Figure for the abstract: 2

The present application relates to the technical field of aero-engine aerodynamic optimization and computer-aided design, and discloses a numerical optimization design method and structure for aero-engine low-pressure turbine blade leading edge modeling based on response surface modeling. The method realizes precise matching of Klebanoff stripe and large-scale vortex generation timing by adjusting flow or rotor speed to change wake injection angle, and promotes stripe impact and accelerates large-scale vortex breakup. Meanwhile, by constructing a blade leading edge modeling parameterization model and establishing a response surface proxy model for the combination of modeling design variable parameters and boundary layer velocity mean square deviation, the optimal design variable parameter combination is determined in the constraint domain with the maximum velocity mean square deviation in the boundary layer as the target. The present application effectively reduces the boundary layer momentum thickness and total pressure loss under the premise of ensuring constant blade load, improves the design efficiency and engineering applicability under low Reynolds number conditions, and does not need to introduce additional flow control structures, thus being simple in structure and easy to implement.

The present application relates to a fan blade (10) for an industrial axial fan. The fan blade (10) comprises an airfoil (12) and a blade root (14) arranged proximal to the airfoil (12) and configured for attaching the fan blade (10) to a hub (62) of an axial flow fan. The airfoil (12) includes a tip (22) at a distal end (23) and a base (24) at a proximal end (25), a leading edge (26) and an oppositely disposed trailing edge (28), whereby both edges (26, 28) extend from the tip (22) to the base (24). The airfoil (12) has a variable thickness defined as a distance between a top surface (30) and a bottom surface (32). Furthermore, the airfoil (12) is an injection molded unitary structure integrally formed in one piece from a composite material, where the composite material comprises carbon fibers and a thermoplastic selected from the group consisting of polyamide, polybutylene terephthalate and polypropylene.

A composite laminated layup panel structure suitable for wing combined connection, comprising a main box section skin (100) and a plurality of stiffeners (200) which are integrally cured and formed. The main box section skin is sequentially provided with recesses (300) at a front edge and a rear edge, forming a front recessed edge (110) for overlapping with a leading edge skin (400) and a rear recessed edge (120) for overlapping with a trailing edge skin. A front spar stiffener (600) configured to be connected to a front spar web (500) is fixedly arranged at the bottom of a transition region between the front recessed edge (110) and a main box section skin main body. A rear spar stiffener (700) configured to be connected to a rear spar web is fixedly provided at the bottom of a transition region between the rear recessed edge (120) and the main box section skin main body. The panel structure simplifies connection forms of panels and related structures, reducing the number of parts, improving assembly efficiency, and also facilitating disassembly and replacement maintenance during aircraft use and maintenance.

The utility model discloses a gliding unmanned aerial vehicle with folding wings. A canard-type middle-single-wing aerodynamic layout is adopted; the full-moving front wing can be folded downwards by 90 degrees; the two main wings are an outer main wing and an inner main wing, the outer main wing can be folded downwards by 180 degrees at the joint of the outer main wing and the inner main wing, and the lower surfaces of the wings are opposite; the inner main wing can be folded upwards by 90 degrees at the joint of the inner main wing and the main wing reinforcing rib; main wing reinforcing ribs on two sides of the fuselage are fixed with main wing rotating shafts and can rotate around the main wing rotating shafts, so that the main wing reinforcing ribs are switched between a horizontal state and a vertical state; when the two sections of main wings are folded and placed in parallel to the fuselage, the front edges of the folded main wings are below; the vertical empennage can be retracted into the fuselage. Compared with a conventional layout, the canard layout has better maneuverability, the canard wings and the main wings generate lift force, and the folded wings have a large aspect ratio after being unfolded, so that the large lift-drag ratio is favorably provided, and gliding flight of the unmanned aerial vehicle is facilitated.

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 turbine blade (30) comprising: a flow profile section (34) extending in a blade height direction and having a pressure surface (46, 46A) and a suction surface (48, 48A), each of which extends between a leading edge (42, 42A) and a trailing edge (44, 44A); a cover section (52) arranged at an outer end face of the flow profile section (34); a rounding section (40) formed by a curved surface (40a) and connected to an end section of the cover section (52) on one side of the flow profile section (34); at least one first cooling hole (60, 60a-60i) extending along the blade height direction within the flow profile section (34); and at least one cooling cavity (70) arranged at least partially within the cover section (52) and connected to with at least one first cooling hole (60, 60a-60i),and a second cooling hole (62) connected to the at least one cooling cavity (70) and open to a surface of the cover section (52), wherein the flow profile section (34) has a reference flow profile (34A) in which a maximum blade thickness at a reference position (PA) in the blade height direction is minimal, wherein the at least one cooling cavity (70) comprises a cavity (72) extending such that it overlaps the fillet section (40) in the blade height direction, wherein in a cross-section enclosing the cavity (72) perpendicular to the blade height direction, the cavity (72) extends inside and outside a first region where a contour of the reference flow profile (34A) is projected onto the cross-section in the blade height direction, and wherein the cavity (72) has: a bottom surface (76),which extends along a direction perpendicular to the blade height direction; and an inner wall surface (78) which extends along the blade height direction into an extension area (Ra1, Rb1) into which the rounding section (40) extends, and which is connected to the bottom surface (76) in an area, viewed from the blade height direction, outside the first area at the cross-section.

The utility model belongs to the technical field of aerospace, and particularly relates to a multifunctional test carrier for a wing leading edge structure, which comprises a test section, a test section tip transition section and a test section root transition section, a metal carrier, a metal carrier tip and a metal carrier root; the end face of one side of the test section root is in butt joint with the end face of the test section root transition section, the end face of one side of the test section tip is in butt joint with the end face of the test section tip transition section, and the test section, the test section root transition section and the test section tip transition section form a wing leading edge part. The end face of one side of the metal carrier root is in butt joint with the end face of the metal carrier root, the end face of one side of the metal carrier tip is in butt joint with the end face of the metal carrier tip, and the metal carrier, the metal carrier tip and the metal carrier root form a wing rear section part; the rear end surface of the wing front edge part is butted with the front end surface of the wing rear section part to form a microwave darkroom test carrier; the front end face of the metal carrier is butted with the rear end face of the test section to form an icing wind tunnel test carrier.

This invention belongs to the field of milling technology for thin-walled blade surfaces of aero-engines, specifically relating to a compensation method for shortening the chord length during blade surface milling. The compensation method includes the following steps: Step 1, obtaining the leading edge thickness deviation Δh after precision milling of the nth (n = 1, 2, 3…) section of the blade surface. qn Step 2: Obtain the leading edge chord length deviation ΔA after precision milling of the nth (n = 1, 2, 3...) section of the profile. qn Step 3: Calculate the actual compensation value Δ of the leading edge chord length of each section. qn (n = 1, 2, 3…); Step 4: Compensate for the leading edge chord length of each section; Step 5: Construct new trailing edge profiles for each section; Step 6: Regenerate the machining model; Step 7: Machining according to the new profile model and inspecting again. By using the compensation method of this invention, the milling accuracy of the leading and trailing edges of the blade is greatly improved, and the chord length deviation distribution is more reasonable, which is beneficial to improving the performance of the product.

An apparatus is provided for an aircraft propulsion system. This apparatus includes a variable guide vane. The variable area vane includes an airfoil, and the variable guide vane is configured to pivot about an axis. The airfoil extends spanwise along the axis to a tip. The airfoil extends longitudinally along a camber line between a leading edge and a trailing edge. The airfoil extends laterally between a first side and a second side. The airfoil includes a vane acoustic structure. This vane acoustic structure includes a first skin and a plurality of chambers within the airfoil. The first skin at least partially forms the first side. A first of the chambers is fluidly coupled with one or more perforations in the first skin.

The invention provides a trailing edge variable camber airfoil composed of a bistable composite material structure and a design and control method of the trailing edge variable camber airfoil, and relates to the technical field of aircraft design and aerodynamics. Wherein the trailing edge variable camber airfoil sequentially comprises a leading edge section, an airfoil middle section and a trailing edge section in the chordwise direction, a front beam is arranged between the leading edge section and the airfoil middle section, the trailing edge section is fixed to an airfoil rear beam, and the trailing edge section is a core function section and is formed by arranging multiple layers of glass / carbon hybrid bistable composite material structures from top to bottom in the wingspan direction; each glass / carbon hybrid bistable composite material structure has a first stable state and a second stable state which correspond to a first flight state and a second flight state of the airfoil profile respectively, one end of each glass / carbon hybrid bistable composite material structure is fixedly connected to the airfoil profile rear beam, and the other end of each glass / carbon hybrid bistable composite material structure is fixed into a whole through glue joint. According to the invention, the steady-state characteristic of the bistable composite material is utilized, so that non-driven conversion between two stable bending states is realized.

The present invention describes a rotor blade (2) having a leading edge (23) and a trailing edge (24), wherein the rotor blade (2) is designed and configured to be exposed to a fluid flowing substantially from the leading edge (23) to the trailing edge (24), the rotor blade (2) including at least two sensors (31) designed to detect the flow characteristics of the fluid and provide corresponding sensor signals, and the rotor blade (2) further including at least two actuators (32) for generating an anti-noise signal (36) based on the sensor signals, the sensors (31) being arranged along the spanwise direction at a surface of the rotor blade (2), and the actuators (32) being arranged along the chord direction at a surface of the rotor blade (2), and the actuators (32) being arranged and prepared such that flow-induced edge noise (35) of the rotor blade (2) generated by the fluid is at least partially eliminated by the anti-noise signal (36). The present invention further describes a method for forming such a rotor blade, and a related wind turbine.

Disclosed are a wing for an aircraft, a leading edge high lift assembly and an aircraft, the wing comprising a main wing and a leading edge high lift assembly, the leading edge high lift assembly comprising a slat and a connecting assembly movably connecting the slat to the main wing, and a connecting assembly for connecting the slat to the main wing such that the slat is movable between a retracted position, a first extended position, and a second extended position, where the connecting assembly comprises an elongate slat track, a first end and / or a middle portion of which is mounted to the slat, the slat track being mounted to the main wing by means of a roller support. The aim of providing a wing with a beneficial ratio of aerodynamic performance to noise resulting from the extension of the slat is achieved, in which the slat track and the roller support are configured such that, in a retracted position and a first extended position, the slat rests against the main wing and, in a second extended position, the slat rests against the main wing. A gap is formed between the trailing edge of the slat and the leading edge of the main wing.

A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

Aspects described herein relate to flow turning across an aircraft wing. An aircraft may include propulsors configured to propel an airflow across the upper surface of the aircraft wing. Due to the Coandă effect, the airflow may adhere to the upper surface of the aircraft wing. A movable flap may be attached to a trailing edge of the aircraft wing. The flap may be configured to rotate between retracted and rotated positions. The flap may be configured to rotate up to and beyond 90 degrees. During rotation, another flap may pivot to a deployed position to provide a continuous surface between a trailing edge of the aircraft wing and a leading edge of the moveable flap for the airflow to adhere to, which facilitates turning the flow of air across the aircraft wing.