1389 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 additive manufactured (AM) metal coupon for a turbomachine component is provided. The component includes an airfoil body having a pressure side, a suction side, a trailing edge, and a coupon opening. The AM metal coupon has an AM metal member with at least one slanted coupon side wall at an angle of less than 90° with an exterior surface of the AM metal member. The AM metal member also has a first porous region having a first porosity and a second porous region having a second porosity different than the first porosity. A braze material couples the AM metal coupon in the coupon opening in the airfoil body and infiltrates into at least one of the first porous region based at least on a characteristic of the first porosity and the second porous region based at least on a characteristic of the second porosity.

An additive manufactured (AM) metal coupon (200) for a turbomachine component (202) is provided. The component (202) includes an airfoil body (206) having a pressure side (152, 178), a suction side (154, 180), a trailing edge (158, 184), and a coupon opening (204). The AM metal coupon (200) has an AM metal member (290) with at least one slanted coupon side wall (292, 294) at an angle of less than 90° with an exterior surface (296) of the AM metal member (290). The AM metal member (290) also has a first porous region (300A) having a first porosity and a second porous region (300B) having a second porosity different than the first porosity. A braze material (310) couples the AM metal coupon (200) in the coupon opening (204) in the airfoil body (206) and infiltrates into at least one of the first porous region (300A) based at least on a characteristic of the first porosity and the second porous region (300B) based at least on a characteristic of the second porosity.

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.

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 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 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.

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 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 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 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 utility model provides a wind resistance balancing wing plate, which belongs to the technical field of balancing wing plates, and comprises a wing plate main body, the wing plate main body is divided into a front edge, a wing-shaped middle part and a rear edge along the spanwise direction, the wing plate main body is provided with a sawtooth structure, the sawtooth structure is continuously arranged triangular sawteeth, and the triangular sawteeth are arranged in the wing plate main body. The sawtooth structures are integrated on the edge of the wing plate body to form an integrated structure, the vertex of each sawtooth is located on the edge line of the wing plate body, gears of the sawtooth structures are arranged in a staggered mode, and the sawtooth structures are continuously distributed along at least one side edge of the wing plate body. The sawtooth structures are continuously distributed along the edge of at least one side of the wing plate body, preferably distributed on the edges of the two sides, so that the airflow optimization effect is maximized, turbulent flow of a traditional wind resistance balance wing plate is reduced, the smoothness of airflow is improved, and therefore wind resistance is reduced, and stability and efficiency are enhanced.

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.

An airfoil includes an airfoil section that defines a trailing edge region that includes a trailing edge. The airfoil section is formed of a ceramic matrix composite that includes core fiber plies and skin fiber plies. The core fiber plies define a radial tube that has a radiused end in the trailing edge region. The skin fiber plies wrap around the core fiber plies and a filler element in the trailing edge region is aft of the internal cavity and is sandwiched between the skin fiber plies on the pressure side and the skin fiber plies on the suction side. There is at least one cooling passage that includes a first, inlet orifice section that opens to the internal cavity at a location forward of the radiused end and that extends through the core fiber plies, and a second, outlet orifice section that extends through the trailing edge.

The invention relates to the technical field of aircraft control, in particular to an airfoil gust retarding feedforward control method based on a circular rector control technology, which comprises the following steps: modifying the shape of the rear edge of an NACA0012 airfoil, and arranging jet holes in the rear edge to form a circular rector control airfoil; performing a fluid simulation experiment on the circular rector control airfoil based on the gust speed, and constructing a model of the airfoil lift coefficient disturbed by the gust speed; performing a fluid simulation experiment on the circular rector control airfoil based on the jet flow momentum coefficient, and constructing a model of which the airfoil lift coefficient is controlled by the jet flow momentum coefficient; based on the model, obtaining a gust slowing-down feedforward control model of which the input quantity is the vertical gust speed and the output quantity is the jet flow momentum coefficient; controlling jet flow of a trailing edge by adopting the gust slowing-down feedforward control model to complete gust slowing-down control of the circular rector control airfoil; the method can improve the ability of the aircraft to resist the gust disturbance environment.

The present invention provides a noise reduction blade and a fan thereof, which relates to the technical field of fan devices, including a wind wheel blade, the wind wheel blade including a blade pressure surface, a blade suction surface, a blade leading edge and a blade trailing edge; the blade leading edge adopts a corrugated noise reduction design, and the blade is designed with corrugated noise reduction in both the circumferential and axial directions from the blade tip to the blade root. The present invention designs corrugations in the circumferential and axial directions on the blade leading edge, so that the blade leading edge has airflow inlet angles of different angles from the blade tip to the blade root, so that when the airflow enters the wind wheel blade, it interacts with the blade leading edge to generate pressure pulsations in different directions, which can effectively reduce the fan inlet noise; the blade is designed with corrugations in the circumferential and axial directions from the blade tip to the blade root, so that an airflow guide flow channel from deep to shallow from the blade leading edge to the blade trailing edge can be formed on the blade pressure surface and the blade suction surface, which can increase the blade strength, and can also play a role in guiding the airflow so that the airflow flows more evenly on the blade surface, thereby achieving the effect of improving the fan efficiency and reducing noise.

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 discloses a flight-oriented fan wing trailing edge adjusting device, and relates to the technical field of aircrafts, the flight-oriented fan wing trailing edge adjusting device comprises a trailing edge angle adjusting mechanism, a rectification angle adjusting mechanism, a circular end plate and a rectification plate, the trailing edge angle adjusting mechanism comprises a steering engine, a first connecting rod mechanism, a trailing edge plate and a rotating shaft; the steering engine is fixedly installed on the circular end plate and controls the rear edge plate to move through the first connecting rod mechanism, the rotating shaft is arranged on one side of the rear edge plate, and the rear edge plate is hinged to the rectifying plate through the rotating shaft. By freely adjusting the trailing edge angle and the rectification angle, the aerodynamic force of the fan-wing aircraft is changed, the lift force and thrust distribution is improved, the aerodynamic effect is enhanced, the energy consumption is reduced, and the flight efficiency is improved. The fan wing trailing edge adjusting device is applied to a serial fan wing, forward coupling gain between two wings is achieved by changing the size of a trailing edge angle and a rectification angle, different flight conditions are met, various flight requirements are flexibly met, and the load performance and the endurance are optimized.

A process for manufacturing a blade made of composite material for a turbomachine. The blade includes an airfoil having a pressure side and a suction side which extend from a leading edge to a trailing edge of the airfoil. The blade further includes a metal sheath that extends along the leading edge of the airfoil. The process includes the steps of: placing a preform, produced by three-dimensionally weaving fibers, in a mold, the sheath being positioned on an edge of the preform intended to form the leading edge of the airfoil; and injecting polymerizable resin into the mold to impregnate the preform so as to form the airfoil after solidifying. At least one double-sided adhesive film may be inserted between the sheath and the edge of the preform prior to injection of the resin.