1131 results about "Trailing edge" patented technology

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.

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.

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.

The utility model relates to the field of aircrafts, aims to solve the problem that a control surface of a known aircraft is heavy or inconvenient to disassemble and assemble, and provides an airfoil structure and an aircraft. The airfoil structure comprises an airfoil body, a control surface, a driving assembly and two rotary connecting mechanisms. And the control plane is arranged on one side of the airfoil body. The two rotary connecting mechanisms are arranged in the length direction of the control surface in a spaced mode. And the control surface is rotatably connected to the rear edge of the airfoil body through the two rotary connecting mechanisms. The driving assembly comprises a steering engine and a transmission mechanism. The steering engine is installed on the airfoil body. The transmission mechanism is in transmission connection between the steering engine and the control plane and used for driving the control plane to rotate relative to the airfoil body. The position, connected with the control plane, of the transmission mechanism is located between the two rotary connecting mechanisms. The utility model has the beneficial effects that the control surface is light in weight and convenient to disassemble and assemble.

The utility model relates to a cross-medium aircraft convenient for adjusting flight attitude, which comprises a water-drop-shaped centrosymmetric aircraft body with an aircraft head and an aircraft body, and the aircraft head of the aircraft body is positioned on the front side of the aircraft body. The aircraft body is further provided with an underwater power mechanism used for providing underwater flight power and an air power mechanism used for providing air flight power, the side portion of the aircraft body is connected with four or six fixed wings, slip-flow rudders controlled to rotate are arranged on the rear edges of the fixed wings, and the rotating axes of the slip-flow rudders are parallel to the wing faces of the fixed wings and perpendicular to the long axis of the aircraft body. Each fixed wing is provided with a slipstream rudder, so that the cross-medium aircraft can be conveniently controlled to pitch, roll, yaw and other actions, and the cross-medium aircraft can be conveniently switched between a hovering flight attitude and a level flight attitude.

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.

The invention provides a wind tunnel experiment-based wing high-attack-angle flow separation combined intelligent control method, which comprises the following steps of: arranging active flow control devices at 10% chord length and trailing edge flap of a wing to inhibit wing high-attack-angle flow separation; the active flow control device at the 10% chord length of the wing adopts direct jet flow, the jet flow direction is perpendicular to the chord line of the wing, three jet flow exciters are uniformly arranged along the spanwise direction, the active flow control device at the trailing edge flap adopts sweeping jet flow, the jet flow direction is parallel to the surface of the wing at the position, and 18 sweeping jet flow actuators are uniformly arranged along the spanwise direction; the jet flow of the two sets of flow control devices is adjusted through corresponding mass flow controllers, jet flow momentum distribution is dynamically optimized in combination with a reinforcement learning algorithm, the lift coefficient is remarkably increased, and energy consumption is reduced. The flow field state is fed back in real time through a pressure measuring system and a force measuring system, a closed-loop control system is constructed, and an execution action is output according to the flow field information around the wing to control an active flow control device arranged on the wing so as to inhibit flow separation on the surface of the wing, so that the wing gets rid of the stall state and the lift force is improved.

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 application provides an integrated flame stabilizer with a tail edge cavity, which comprises a flame stabilizer, an oil injection rod, a cooling gas inlet pipe and a tail edge wing plate arranged at the rear side of the flame stabilizer, and a stable backflow ignition area is formed in the downstream area of the flame stabilizer; a plurality of cooling channels are arranged on the rear side plate of the flame stabilizer between the two tail edge wing plates, and the airflow direction at the outlet of the cooling channels is inclined to the flow direction, so that a cooling vortex is formed in the tail edge cavity, the flame stabilizer and the backflow ignition area are spaced from each other, the flame stabilizer and the downstream backflow ignition area are isolated from each other to reduce the thermal load of the components such as the flame stabilizer and the fuel injection pipe, the ablation and coking problems are avoided, meanwhile, the flame stabilizing function of the flame stabilizer is not damaged, and the influence on the backflow ignition area behind the stabilizer is small.

The invention belongs to the technical field of glide control, and relates to a direct lift active disturbance rejection control method suitable for flying wing layout. According to the method, a trailing edge flap is introduced to serve as a main lift control surface, meanwhile, a constant attack angle is kept through a lifting aileron so as to balance part of pitching moment generated by the trailing edge flap, and therefore the attitude of the aircraft is stabilized, and the constant speed is maintained. In the actual flight process, external disturbance is inevitable, so that an original glide control strategy is improved and optimized in combination with an active disturbance rejection control method. Compared with a traditional glide control method, the flying wing layout direct lift active-disturbance-rejection control method provided by the invention has remarkable anti-interference capability and excellent glide control performance, and shows a wide application prospect.

A design method for a blade with orientation structures, a blade and a performance test method for a blade are provided, which relate to the technical field of blade design, and are universal to blades whose parameters are completely or partially known. In the design method, orientation awls are provided at a trailing edge of a blade from a variable-pitch propeller; orientation awls are provided at a trailing edge of a blade from a fixed-pitch propeller or a fan blade from a jet engine, and orientation grooves are provided on a back of the blade. Meanwhile, during providing orientation awls for either blade, the blade is stretched in a chord direction to ensure that an area of a pressure surface of a modified blade is equal to an area of a pressure surface of its original.

The invention provides a friction force identification method for a trailing edge flap driving mechanism with a bearing, and the method comprises the steps: applying a series of improved triangular waves with different speeds to a driver, carrying out the closed-loop control, carrying out the zero-phase digital filtering of a collected flap angle signal, and suppressing the interference of random noise to the signal, adjusting a buffeting phenomenon of a derivation step length inhibition speed to obtain a flap angular speed, then intercepting a constant-speed ascending section and a constant-speed descending section, then assigning a series of angular speeds, carrying out friction model identification on a relation between friction force and the speed, and then identifying a relation between parameters obtained by identification and a flap angle; finally, the friction force about the flap angle and the flap angular velocity in the flap driving mechanism is obtained.

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.

The invention relates to a stator blade and an open rotor engine comprising the stator blade. The stator blade is used for the rear row of the open rotor engine. The stator blade extends from the blade root part to the blade tip part, so that the noise of the open type rotor fan can be effectively reduced. A tip portion of the stator blade has a tip surface, a leading edge of the stator blade at the tip surface extending a distance from a root portion of the blade is farther than a trailing edge of the stator blade at the tip surface extending a distance from the root portion of the blade such that the tip surface is an inclined surface with a higher leading edge and a lower trailing edge when viewed from a side view.

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.

Cambering device for a profiled sail including: —an elastically deformable semi-rigid structure, including: —two length sections, extending one in front of the other, forming respectively a first length section secured to a first surface of the profiled sail, and a second length section secured to a second surface of the profiled sail, —a U-shaped connecting section, connecting together the two length sections terminating at the trailing edge by a tensioning of one of the two free ends, —a spacing system configured for working in compression and in tension, for maintaining a separation between the two length sections, —the action of the actuator, connecting together the two free ends of the semi-rigid structure.

The invention provides an AI-based engine blade defect identification and detection method and system, and the method comprises the steps: S1, scanning an engine blade through an eddy current detection probe according to the sequence of the front edge, the blade body, the rear edge and the root of the engine blade, obtaining an eddy current response signal, and transmitting the eddy current response signal to a signal processing system; s2, extracting defect features from the defect eddy current signal of the defective engine blade; extracting a defect-free feature from the defect-free eddy current signal of the defect-free engine blade; s3, establishing and training an engine blade detection model through the extracted defect features and defect-free features; and S4, whether the engine blade to be detected has defects or not is detected through the engine blade model. According to the invention, based on artificial intelligence eddy current detection, the accuracy and efficiency of detection are improved, the probability of missing detection and false detection is reduced, and the technical problem of missing detection or false detection caused by low detection efficiency in the prior art is solved.

The invention relates to a steering engine and piezoelectric patch hybrid driven variable camber wing, belongs to the technical field of aerospace, and solves the problems of poor wing surface deformation continuity and poor transmission effect in the prior art. The plurality of flexible trailing edge deflection modules are arranged in the straight wing main body; the skin structure is arranged on the outer surfaces of the straight wing main body and the flexible trailing edge deflection module; the module connecting structure is connected with the plurality of flexible trailing edge deflection modules; wherein the straight wing main body comprises a front edge D-shaped structure wing box, a plurality of wing ribs arranged on the rear side of the front edge D-shaped structure wing box, and a bearing rear wall arranged on the rear sides of the wing ribs; each flexible trailing edge deflection module comprises a driving rib assembly fixed to the bearing rear wall, driven rib assemblies arranged on the two sides of the driving rib assembly, carbon rods sequentially connected with the driving rib assembly and the driven rib assemblies, an inhaul cable driving mechanism and a piezoelectric driving mechanism, and the driving rib assembly and the driven rib assemblies are sequentially arranged in the longitudinal direction of the variable-camber wing.

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 invention provides a blade, a noise reduction system, a fan rotor, an open rotor engine and a noise reduction method. The blade comprises a plurality of sawteeth, the sawteeth are distributed on the tail edge of the blade in the spanwise direction of the blade, and the sawteeth are rotatably connected with the tail edge around the axis extending in the spanwise direction of the blade. The actuator is connected with the sawteeth and used for adjusting the rotation angle of the sawteeth; by the adoption of the sawteeth, the phenomenon that vortexes periodically fall off can be effectively restrained when airflow passes through the blades, so that noise generated by the falling vortexes is reduced, the angle of the sawteeth relative to the tail edge is adjusted, the direction of the airflow at an outlet can be changed, the airflow is smoother, the continuity of the vortexes is damaged, and therefore high-frequency noise is reduced.

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.

The utility model relates to the technical field of storage and transportation equipment of empennages of unmanned aerial vehicles, in particular to a loading structure of an empennage of an unmanned aerial vehicle, which comprises two frame bodies arranged at an interval, and the interval between the two frame bodies forms an accommodating space of an empennage body; each frame body comprises an upper cross beam and a lower cross beam which are vertically arranged at an interval, and an oblique beam is jointly welded at the front ends of the upper cross beam and the lower cross beam; a plurality of supporting plates are welded between the lower cross beams of the two frame bodies; a plurality of empennage fixing plates are mounted between the oblique beams of the two frame bodies through screws; and a trailing edge mounting plate and a trailing edge middle supporting plate are mounted between the upper cross beams of the two frame bodies through screws. The loading structure provided by the utility model is simple in structure, can load the empennage body and the tail edge of the wing at the same time, and effectively improves the loading efficiency. The loading structure is used for loading the empennage body and the tail edge of the wing, so that the whole can be conveniently put into a container for transportation.

The present application belongs to the technical field of flame stabilizer design in aircraft engines, and specifically relates to a double-bend premixed preevaporated flame stabilizer, which is a double-bend flame stabilizer that uses external airflow for cooling and premixed preevaporation with fuel. The premixed preevaporation of fuel and external airflow can be used to improve the oil-gas ratio at the trailing edge of the stabilizer, thereby improving the ignition and flame stabilization capabilities at the left boundary of the aircraft engine, and can form an air film cooling on the tail plate through two rows of cooling holes to ensure structural reliability at the right boundary of the aircraft engine and when the intake temperature is high.

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.

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 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 belongs to the technical field of unmanned aerial vehicles, and particularly relates to a tailstock type overwater vertical take-off and landing fixed-wing unmanned aerial vehicle which can float and park on the water surface, can vertically take off to leave the water and is switched into a fixed-wing mode for cruising and level flight. The unmanned aerial vehicle comprises a four-rotor support buoy and a wing body fusion body fixed to the upper portion of the four-rotor support buoy, small wing type buoys are arranged on wing tips of the two sides of the wing body fusion body respectively, and lifting ailerons capable of deflecting up and down are arranged on the rear edges of wings on the two sides of the wing body fusion body respectively. And four motor rotor propeller propelling devices are uniformly arranged at the tip part of the four-rotor bracket buoy.

A wing designed to pass through a fluid medium is provided, the wing having a plurality of alternating protrusions and valleys along a trailing edge thereof. The described wing uses the plurality of protrusions and valleys to counteract noise that is a function of wind speed and chord length. In an alternative variant, instead of or in addition to cancelling noise, the plurality of protrusions and valleys are adapted to generate or correct noise such that the noise is generated in a desired frequency band. In this variation, the plurality of protrusions and valleys are configured such that the generated noise is in the desired frequency band based at least on a frequency band that repels various types of animals given a specified local wind speed. In another variation, the plurality of protrusions and valleys are adapted to improve the aerodynamic performance of the wings, whether in the context of separate use or with multiple wings working together within the system.