643 results about "Trailing edge" patented technology

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 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.

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.

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.

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.

An aircraft may comprise propulsors. The propulsors may be integrated into an array of propulsors. An array of propulsors may be integrated in an airfoil of the aircraft. The aircraft may include a flow turning control system. The flow turning control system may include an end wall and one or more movable flaps. The flow turning control system may enable selective flow turning of the flow of propelled by the array of propulsors. The selective flow turning may include actuating the one or more movable flaps. The selective flow turning may achieve flow turning of over a trailing edge of a movable flap.

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 self-propelled watercraft vehicle has a floating body (1), underwater holding structure that is connected to a propulsion device (8) and consists of a mast (6) and a beam (7) that is transverse to the mast, a hydrofoil (9), a stabilizing foil (10), a central control unit (5), speed and roll sensors. Hydrofoil panels (11) have an asymmetric streamlined profile, stabilizing foil panels (12) have a streamlined profile. On the trailing edge (14) of both hydrofoil panels (11) there are control elements (15) that enable a hydrodynamic lifting force. The central control unit (5) provides automatic roll elimination depending on the speed and is designed to interact with a remote control.

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.

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.

An aircraft may comprise propulsors. The propulsors may be integrated into an array of propulsors. An array of propulsors may be integrated in an airfoil of the aircraft. The aircraft may include a flow turning control system. The flow turning control system may include an end wall and one or more movable flaps. The flow turning control system may enable selective flow turning of the flow of propelled by the array of propulsors. The selective flow turning may include actuating the one or more movable flaps. The selective flow turning may achieve flow turning of over a trailing edge of a movable flap.

Disclosed is an aerodynamic system for an underside of a vehicle, the system comprising: an aerodynamic chassis, the aerodynamic chassis comprising a main compartment bounded by a plurality of walls and a floor extending between the walls; and an aerodynamic device, the aerodynamic device comprising: a mounting structure, the mounting structure being fixedly mounted to the floor of the aerodynamic chassis; and an aerodynamic body, the aerodynamic body comprising a leading edge, a trailing edge and at least two outboard edges, the aerodynamic body being fixable to the floor of the aerodynamic chassis by the mounting structure, the mounting structure being located wholly inboard of the outboard edges of the aerodynamic body.

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 invention discloses a built-in main control actuator connecting structure of an aircraft wing box, and relates to the technical field of aircraft structure design, an actuator is located in the aircraft wing box structure, an actuating rod of the actuator penetrates through an actuator rod avoiding hole of the aircraft wing box structure, and the actuating rod stretches out and draws back to cooperate with the actuator to rotate up and down in the wing box to achieve deflection of a control surface. The invention is suitable for small airfoils and small and medium-sized airfoils with low airfoil height, solves the problems that the small airfoils of the small airfoils are small, the trailing edge space is limited, and a control surface actuating system cannot be mounted, and avoids the problem that the aerodynamic efficiency is influenced as an actuating structure is arranged outside the airfoils; the structure of the actuator connector is designed and arranged on the front main bearing beam of the wing box, the load transfer continuity of the wall plate structure is not broken, meanwhile, the actuating load can be directly diffused to the skin through the support connector, and the load transfer efficiency is higher. By separating a daily inspection and maintenance passage from a disassembly and assembly passage, the opening size of the airfoil wall plate is reduced, the load transfer capacity of the lower airfoil skin is guaranteed, and structural weight reduction is facilitated.

The invention discloses a control surface structure with a tangential air blowing seam in the technical field of control surface structures, and particularly comprises a control surface supporting beam, a control surface front edge and a control surface rear edge are arranged on the control surface supporting beam, an air blowing cavity is formed in the control surface front edge, the air blowing seam is formed in the joint of the control surface front edge and the control surface rear edge, and the air blowing seam communicates with the air blowing cavity; the lower surface of the air blowing seam is flush with the control surface rear edge, and the air blowing direction of the air blowing seam faces the tangential direction of the control surface rear edge. The air blowing direction of the air blowing gap faces the tangential direction of the control surface rear edge, so that flow separation inhibition of air on the upper surface of the control surface rear edge is better completed, the lift coefficient of the control surface structure is effectively improved, and improvement of the overall lift coefficient of an aircraft is affected.

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 wide-speed-range tailless variable-structure aircraft comprises an aircraft body, the aircraft body comprises an aircraft body-fixed wing fusion body and variable-structure wing rudders, the variable-structure wing rudders are symmetrically arranged on the two sides of the aircraft body-fixed wing fusion body, and a wing body fusion body layout is integrally formed; the fuselage-fixed wing fusion body is connected with the variable-structure wing rudder through a driving mechanism, and the variable-structure wing rudder can deflect in the axial direction of the fuselage; the tail of the fuselage-fixed wing fusion body is provided with a trailing edge rudder, the trailing edge rudder of the fuselage-fixed wing fusion body achieves pitching channel control through same-direction deflection, and under the condition that the channel coupling coefficient is smaller than 1.5, rolling channel control is achieved through differential deflection. The problem that the yaw channel operation stability characteristic is insufficient under the tailless layout condition is solved, and the wide-speed-range operation performance of the aircraft is effectively improved; compared with other variable-structure aircrafts in which additional variables are introduced, the combination of the variable-structure wing rudder and the trailing edge rudder is simple and efficient, and the burden on a control system is small.

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.

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.

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.

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.

The application discloses a rudder surface structure with a tangential blowing slit, and specifically comprises a rudder surface support beam, a rudder surface front edge and a rudder surface rear edge are arranged on the rudder surface support beam, a blowing cavity is arranged in the rudder surface front edge, a blowing slit is arranged at the joint of the rudder surface front edge and the rudder surface rear edge, and the blowing slit is communicated with the blowing cavity; the lower surface of the blowing slit is flush with the rudder surface rear edge, and the blowing direction of the blowing slit is tangential to the rudder surface rear edge. The blowing direction of the blowing slit is tangential to the rudder surface rear edge, so that the flow separation of the gas on the upper surface of the rudder surface rear edge is better inhibited, and the lift coefficient of the rudder surface structure is effectively improved, and the overall lift coefficient of the aircraft is improved.

The invention discloses a wing, an aircraft, a wing design method, a storage medium and electronic equipment. The wing comprises a wing body and a control piece, and the control piece comprises a first end, a second end, a third end and a fourth end, the second end is arranged on one side of the first end, and a first surface is formed between the second end and the first end; wherein a preset angle is formed between the first surface and the upper surface of the wing body, and a preset height is formed between the first end and the second end. The wing can eliminate the head lowering pitching moment while having a larger maximum lift coefficient, and has the advantages of increasing the lift of the wing, reducing the area requirement of the wing, reducing the balancing loss of the pitching moment, being convenient to realize the balancing of the pitching moment and the like.

A propeller fan blade for aircraft propulsion. A propeller fan blade (110) with a saw-toothed trailing edge (BF), wherein a first and second skeleton line connect the leading edge (BA) to a tooth crest and an adjacent groove in the trailing edge (BF), respectively, and the absolute value of a variation (ΔβS) of a tilt angle (βS) of the first skeleton line is less than a variation (ΔβC) of a tilt angle (βC) of the other skeleton line over corresponding terminal segments adjacent to the trailing edge (BF). Each of said terminal segments of the first and second skeleton lines extends over 30%, or even 20%, of the total length of the respective skeleton line between the leading edge (BA) and the trailing edge (BF). Figure for the abbreviation: Fig. 2.

The invention belongs to the field of aerodynamics, and particularly relates to a quick treatment method for a trailing edge step of an airfoil profile. The method comprises the following steps: standardizing airfoil profile data, calculating a thickness distribution curve and a camber distribution curve of the airfoil profile through airfoil profile coordinates, constructing a new continuous disturbance function which is a continuous derivable function, and the disturbance quantity of the function at the rear edge is just the required step thickness. And the smooth and continuous disturbance function is superposed on the thickness distribution, so that local rapid adjustment of the thickness of the step on the rear edge of the airfoil is realized, finally, the coordinates of the airfoil are inversely calculated through the camber distribution and the thickness distribution, and the airfoil profile meeting the requirement is obtained. Due to the fact that the camber distribution and the maximum thickness of the airfoil profile modified through the method are not changed, the requirement for continuity of the appearance is met, and it can be guaranteed that the overall aerodynamic characteristics of the airfoil profile before and after adjustment are basically consistent.

This invention belongs to the field of propulsion and control technology for vertical takeoff and landing (VTOL) aircraft, and discloses a cycloidal propeller control mechanism and method for driving trailing edge flap deflection using a dual four-bar linkage mechanism. Based on the traditional single four-bar linkage mechanism used to control the periodic pitch motion of the cycloidal blades, this mechanism adds a four-bar linkage mechanism with a similar structure and principle, specifically designed to independently drive the trailing edge flaps for periodic deflection. By separately setting the pitch eccentricity and pitch eccentricity angle of the pitch control four-bar linkage, and the deflection eccentricity and deflection eccentricity angle of the flap deflection four-bar linkage, the pitch motion of the blades and the deflection motion of the flaps can be independently controlled. Through numerical optimization design, a preset phase difference can be established and maintained between the deflection motion of the flaps and the pitch motion of the blades, and the maximum deflection angle of the flaps is constrained within the target amplitude range, thereby achieving effective improvement and active control of the cycloidal propeller's aerodynamic performance.

The application discloses a kind of leaf type thickness distribution construction method, device, blade and aero-engine, leaf type thickness distribution includes that the front edge section of thickness distribution curve is oval arc, the second section of thickness distribution curve is concave circular arc, the third section of thickness distribution curve is convex circular arc and the trailing edge section of thickness distribution curve is circular arc, design parameter includes front edge length-width ratio e, front edge wedge angle, front edge thickness, maximum thickness, maximum thickness position and trailing edge thickness, the front edge section is tangent with the second section at front edge thickness position, the second section is tangent with the third section, and the third section is tangent with the trailing edge section.This method utilizes the geometric characteristics of concave circular arc to construct a segment of negative curvature thickness distribution law, realizes the deceleration of airflow in the first half of blade, controls wavefront Mach number, completes the controllable rotor blade thickness rapid construction of blade thickness growth rate on the basis of ensuring wedge angle, realizes the best aerodynamic performance under the condition of having enough structural strength stability.