1281 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 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.

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.

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.

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.

The invention discloses a grid generation method and device for finite element analysis of an aircraft airfoil quadrilateral structure, equipment and a medium, and relates to the technical field of multi-physics field coupling numerical simulation. Curve distances between discrete grid points and upper airfoil surface trailing edge points of an airfoil are calculated; the method comprises the following steps: initially segmenting an airfoil profile grid to obtain quadrilateral regions, calculating first control angular points at a front edge and a rear edge corresponding to an upper airfoil surface and a lower airfoil surface, second control angular points in an airfoil profile, first grid edges around an airfoil profile, intersecting edges at the front edge and the rear edge, and second grid edges in the airfoil profile, and constructing a computational domain grid; repeated grid point deletion, grid point list updating and grid unit list updating are carried out on the computational domain grid; and according to the target grid point list and the target grid unit list, generating a grid for finite element analysis of the quadrilateral structure of the aircraft airfoil, so that the full quadrangle of the grid units and the consistency of the grid units among different airfoil profiles can be ensured.

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.

The invention provides a modular suction type flap sail device. The modular suction type flap sail device comprises an external rotating mechanism and an internal fixed hollow cylindrical shell. The external rotating mechanism comprises leading edge and flap modules and is used for adjusting the posture according to the wind direction; the internal fixed hollow cylindrical shell comprises a hollow cylinder, a fan and the like and is used for providing a negative pressure field. Through the modular design, each module can be independently manufactured and assembled, manufacturing, transportation, hoisting, debugging, maintenance and upgrading and updating are convenient, the cost is reduced, and the flexibility and the energy efficiency are improved. The front edge and the flap module are designed in sections, so that the height and the length-diameter ratio of the sail body can be flexibly adjusted; the hollow cylinder module is segmented and is specially connected, so that air tightness and air suction uniformity are ensured; the hollow rotating platform enables internal and external mechanisms to move independently, so that energy consumption and load are reduced; the fan installation position is optimized, and the wind energy utilization efficiency is improved. The device is stable, reliable and high in adaptability, and has remarkable economic benefits and engineering application value.

The invention discloses a blade profile thickness distribution construction method and device, a blade and an aero-engine, and blade profile thickness distribution comprises a front edge section with a thickness distribution curve in an elliptic arc shape, a second section with a thickness distribution curve in a convex arc shape, a third section with a thickness distribution curve in a concave arc shape and a tail edge section in an arc shape. The design parameters comprise the length-width ratio of the leading edge, the thickness of the leading edge, the maximum thickness, the position of the maximum thickness / the thickness of the trailing edge and the wedge angle of the trailing edge, the leading edge section is tangent to the second section at the position of the thickness of the leading edge, the second section is tangent to the third section, and the third section is tangent to the trailing edge section. The excellent blade design adjusting capacity is achieved through simple and efficient design parameter input, the trailing edge wedge angle serves as the design parameter, the thickness increasing rate of the third part is controlled, and the optimal aerodynamic performance is achieved. A section of thickness distribution rule which shrinks rapidly is constructed by utilizing the geometrical characteristics of the concave arc, controllable diffusion of airflow at the rear half section of the blade is achieved, and the effect of allowance and low loss is achieved.

An airframe includes a body, a first wing, a second wing and a vane. The body extends longitudinally along a longitudinal centerline from a forward end of the body to an aft end of the body. The first wing and the second wing are disposed to opposing lateral sides of the body. The vane projects vertically out from the body to a tip of the vane. The vane extends longitudinally along the body from a leading edge of the vane to a trailing edge of the vane. A first open rotor propulsion system is vertically above and mounted to the body. A second open rotor propulsion system is vertically above and mounted to the body. The vane is configured as a structural barrier laterally between the first open rotor propulsion system and the second open rotor propulsion system.

The invention discloses a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic thermal constraint, which comprises the following steps of: determining a rudder leading edge radius and a rudder leading edge sweepback angle by taking a current state nominal trajectory as a basis and taking a rudder leading edge thermal response temperature as a judgment standard; secondly, the cone angle of one conical surface of the folded rudder is determined through iterative calculation, so that the chordwise size of the high-temperature-resistant refractory metal material and the thickness of a skin and a heat-proof coating at the cone position are designed; and thirdly, through the design of the interference position of the arched shock wave of the heat-proof cup, the gap between the fixed rudder and the movable rudder avoids the interference area of the arched shock wave, so that the chordwise distance between the folding gap and the rudder shaft is determined. By means of the aerodynamic configuration design method, the radius of the front edge of the rudder can be as low as 3 mm, the taper angle is as low as 4 degrees, it can be guaranteed that ablation deformation does not occur when the maximum heat flow of the front edge of the rudder does not exceed 10 MW / m < 2 >, the resistance characteristic of a high-speed aircraft can be effectively reduced, the characteristic of near-zero ablation is achieved, and the good aerodynamic configuration is maintained.

The present invention discloses a porous material transpiration cooling leading edge structure, which includes multiple layers of porous material layers and a cooling cavity provided on one side of the multiple layers of porous material layers. A coolant supply channel is provided in the cooling cavity, and the coolant supply channel is used to supply coolant to the cooling cavity. The multiple layers of porous material layers include an outer layer, an intermediate layer, and an inner layer, and the material of the inner layer is a shape memory alloy porous material. The present invention can adaptively adjust the porosity of the porous material, automatically adjust the usage amount of the coolant, effectively improve the usage efficiency of the coolant, and further enhance the thermal protection effect in the high-temperature tip region. Moreover, the shape memory alloy material can be reused. The porous material transpiration cooling leading edge structure can be used for structures such as the nose cone of an aircraft, the leading edge of a rudder, and the leading edge of a wing that are subjected to more high-temperature heat fluxes.

The invention relates to a turbine guide vane structure, a turbine and an aero-engine, belongs to the technical field of aerospace, and solves the problems that in the prior art, a turbine guide vane structure with an impact pipe is prone to inclination of a guide pipe during installation, and the cooling effect is not ideal. The device comprises an impact pipe, wherein a plurality of positioning bosses are arranged on the outer wall surface of the impact pipe; an impact hole is formed in the positioning boss; a cooling through hole is formed in the wall face of the front edge of the impact pipe and used for forming impact cooling on the front edge of the blade. According to the invention, the positioning precision and the cooling effect can be improved.

Actuation assembly for a trailing edge flap of an aircraft wing, comprising: a base unit configured to be fixated to a main body of the aircraft wing; an intermediate unit movably connected to the base unit and configured to be hingeably connected to the flap at a leading edge of the flap about a first hinging axis extending along the leading edge of the flap, wherein the movability provides adjustability of the flap between a retracted flap position and an extended flap position; and a second actuator configured to controllably rotate the flap with respect to the intermediate unit about the first hinging axis, the second actuator being configured to be movable along with the intermediate unit with respect to the at least one base unit.

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.

A blended wing body (BWB) aircraft including a blended wing body having a main body and wings with no clear demarcation between the wings and the main body along leading edges of the BWB aircraft, a passenger cabin located within the main body and more than one passenger bays located within the passenger cabin, wherein a perimeter of at least one passenger bay is demarcated by at least a divider and wherein the at least one passenger bay is substantially parallel to at least one leading edge of the BWB aircraft.

A fluid turbine has an annular-airfoil duct and passive, leading-edge flaps on the airfoil that enhance yaw stability and reduce extreme loads in excessive fluid-stream velocities when the turbine is in shutdown mode, or when it is parked or otherwise not producing power. Flaps retract and deploy passively to reduce loads by eliminating the lift over the airfoil, providing manageable loads in high-velocity flow at Up-0. The flaps are intended for use in extreme and unlikely events. Appropriate design and material selection mitigates the detrimental effects of dormancy.

The invention provides a turbine blade wall thickness forward design method and system. The method comprises the steps that 1, the type of a turbine blade, the front edge radius and the tail edge radius of the turbine blade, the internal cooling mode of the turbine blade, the initial wall thickness of the front edge of the turbine blade, the initial wall thickness of the middle chord of the turbine blade and the initial wall thickness of the tail edge of the turbine blade are obtained; 2, pneumatic evaluation is conducted on the turbine blade, whether the type meets the design requirement or not is judged, if yes, the step 3 is executed, and if not, the type of the turbine blade is updated, and the step 1 is executed again; 3, performance evaluation is conducted on the turbine blade, whether evaluation results are qualified or not is judged, if yes, the step 4 is executed, and if not, the initial wall thickness of the turbine blade is adjusted, and the step 1 is executed again; and 4, the initial wall thickness of the front edge, the initial wall thickness of the middle chord and the initial wall thickness of the tail edge of the turbine blade serve as the design wall thicknesses. According to the technical scheme provided by the invention, the design period is effectively shortened.

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 provides a wave rider capable of vertically ascending and descending, and belongs to the technical field of wide-speed-range aircrafts, the wave rider specifically comprises a front edge shock wave generating surface, a central fuselage and sweepback wings located on the two sides of the central fuselage, the two sweepback wings are provided with mounting holes penetrating through the upper surface and the lower surface of the wave rider, and lift fans are mounted in the mounting holes; the lift fans are embedded into the mounting holes and used for providing vertical lift force for the wave rider, the axes of the lift fans are parallel to the vertical axis of the wave rider, and the lift fans on the two sweepback wings are symmetrically distributed on the two sides of the central fuselage. According to the treatment scheme, the vertical lifting capacity is provided for the wave rider, and the high-speed flight aerodynamic performance is considered.

The invention discloses a low-frequency broadband sound absorption metamaterial adaptive wing and a design method thereof, and belongs to the technical field of aircraft design. The problem that in the prior art, a traditional aircraft wing and a design method thereof are difficult to achieve the low-frequency and broadband noise reduction effect is solved. The wing comprises a leading edge inner side slat, a leading edge outer side slat, a main wing, a trailing edge inner side flap and a trailing edge outer side flap, the leading edge inner side slat and the leading edge outer side slat are connected with the front side of the main wing, and the trailing edge inner side flap and the trailing edge outer side flap are connected with the rear side of the main wing; the metamaterial sound absorption strips are embedded into concave cavities of the leading edge inner side slat and the leading edge outer side slat respectively, the metamaterial sound absorption strips are embedded into the leading edge of the main wing, each metamaterial sound absorption strip is composed of a Helmholtz resonator array composed of single-layer Helmholtz resonator units, and each metamaterial sound absorption strip is provided with a micro-perforated surface with continuous curvature. The noise reduction level of the wing is effectively improved, the sound absorption frequency band is wider, and the method can be applied to sound absorption wing design.

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.

According to this invention, an aerial vehicle can be provided that reduces the effect of wind in a given direction striking its frame during flight of the aerial vehicle, thereby improving fuel consumption and stability. The aerial vehicle of this invention is equipped with a flight part including a frame to which a plurality of rotor blades including at least a propeller and a motor are connected, wherein the frame includes a right frame and a left frame extending side by side in the front-rear direction of the aerial vehicle, and at least one of the right frame and the left frame has a substantially wing-shaped portion with a leading edge located outside the aerial vehicle and a trailing edge located inside the aerial vehicle relative to a vertical center line in the frame. The substantially wing-shaped shape is a symmetrical wing shape.

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 a split type adjustable blade structure of a variable geometry turbine and application of the split type adjustable blade structure, and relates to the technical field of variable geometry turbines of aero-engines. An adjustable blade is designed into a fixed petal body and a rotating petal body in a split mode, the fixed petal body is fixed in a turbine blade channel, a blade front edge and a pressure face front blade body are formed, and no radial gap exists between the blade front edge and a casing; the structure stability is ensured; the end region leakage is reduced; the rotating valve body is located on the rear portion of the fixed valve body and connected with an external adjusting mechanism through the rotating boss to achieve angle adjustment, and a blade suction face, a tail edge and a blade body on the rear portion of a pressure face are formed. A molded line transition fit structure is arranged between the fixed petal body and the rotating petal body, and the gap between the fixed petal body and the rotating petal body is kept minimum and mechanical interference does not occur in the initial and adjusting processes. A dynamic and static clearance sealing structure is further designed, and clearance leakage flow is effectively restrained by arranging a sealing groove and sealing teeth. The structure can reduce flow loss and clearance leakage loss and improve turbine efficiency and engine performance.

The invention discloses an incoming flow Mach number control method for a transient transonic wind tunnel airfoil test, and belongs to the technical field of wind tunnel airfoil tests. The problem that in the prior art, a traditional Mach number control method adopting a conventional plenum chamber reference point is difficult to be suitable for carrying out an airfoil test in a transient transonic wind tunnel is solved. According to the method, a standard airfoil model is selected, simulation calculation is carried out, and distribution of Mach numbers in the incoming flow direction of the airfoil model in the free flow state, namely the wind tunnel wall-free state, is obtained; s2, performing simulation calculation in a three-dimensional full-simulation state on the wind tunnel by adopting a standard airfoil model which is the same as the step S2, acquiring axial Mach number distribution of side wall center lines of test sections with different opening-closing ratios of the wind tunnel according to the opening-closing ratios, and calculating the axial Mach number distribution of the side wall center lines of the test sections with different opening-closing ratios; and analyzing to obtain a Mach number which is controlled in the incoming flow direction and is 3 times the chord length of the airfoil profile from the front edge of the airfoil profile, and carrying out a standard airfoil profile model wind tunnel test. According to the invention, the accuracy of simulating the incoming flow Mach number and the opening-closing ratio of the test section is improved, and the method can be applied to the temporary impulse transonic wind tunnel airfoil test.