22 results about "Swept wing" patented technology

The invention discloses a wide-speed-range variant high-speed aircraft aerodynamic layout, and belongs to the technical field of aircrafts, the wide-speed-range variant high-speed aircraft aerodynamic layout comprises a fuselage, the front end of the fuselage is provided with a drag reduction rod, the rear end is provided with a thrust vectoring nozzle, the left and right sides are symmetrically provided with triangular sweepback wings, and the top is provided with a supporting structure; a high-pressure capturing wing is arranged on the top of the supporting structure. The shock wave separation effect of the drag reduction rod is combined with the high-pressure capture and reflection shock wave induction mechanism of the high-pressure capture wings, so that the aircraft can achieve effective drag reduction and lift augmentation effects in different flight states, and the lift-drag ratio and adaptability of the high-speed blunt cone aircraft are remarkably improved.

An aircraft 310 has a pair of swept wings 314A meeting at a central region 312. Each wing has an inner wing section 326A containing an integrated payload fuselage (fig.1,20); a thinner thickness oute

This application provides a rapid wing deployment and locking mechanism for swept-wing aircraft. The mechanism includes: a base, a slider, a guide mechanism, a linkage mechanism, a rotating mechanism, and a locking mechanism. The guide mechanism, linkage mechanism, rotating mechanism, and locking mechanism are all mounted on the base. The slider is slidably connected to the guide mechanism and connected to the linkage mechanism. The end of the linkage mechanism away from the slider is connected to the rotating mechanism. The wing is fixedly connected to the rotating mechanism. The locking mechanism is located on the side of the rotating mechanism and is used to lock the rotating mechanism. When the slider is impacted by a pyrotechnic device, the slider moves along the guide mechanism, causing the linkage mechanism to swing. The linkage mechanism drives the rotating mechanism to rotate, which in turn causes the wing to deploy. After the wing is fully deployed, the locking mechanism locks into place with the rotating mechanism. This application eliminates the need for wing reassembly, reduces the need for large packaging space, and allows for rapid wing deployment from under the fuselage. Locking the deployed wing ensures stable flight.

The application discloses a rear-swept wing distributed non-equal-diameter rotorcraft and a control method thereof, which comprises a fuselage, a wing, a tilt-rotor group, a landing gear, a horizontal tail and a vertical tail; the tilt-rotor group comprises four tilt-rotors which are connected to the leading edge of the wing through a support rod; the rotor disc area near the wing tip is larger than the rotor disc area near the wing root, so as to balance the flight efficiency at low speed and the flight speed at high speed, and the rear sweep of the wing is beneficial to improving the stability of the tilt-rotor aircraft in the fixed-wing mode flight, and the environmental adaptability of the aircraft is improved. The application balances the efficiency of the tilt-rotor aircraft in the hovering and cruising states, the aircraft structure is relatively simple, is beneficial to the attitude control of the aircraft and the flight control software design.

The application relates to a variable sweep wing stealth aircraft afterbody wedge shape design method, and belongs to the field of aircraft general design and stealth technology. The application is simple and practical in shape; comprehensive design factors are considered, and comprehensive requirements such as variable tail wing rotation, component coupling, stealth principle and aerodynamic resistance are covered; design constraints can be accurately converted into geometric design; variable tail wing rotation, the stealth principle, aerodynamic resistance and component coupling can be effectively converted into afterbody wedge shape point position and line tangent vector control design to be realized; the fusion design of the afterbody wedge and the surrounding curved surface is better realized; under the premise of guaranteeing the variable tail wing rotation space requirement and the binary nozzle requirement, the aerodynamic and stealth shape integration design of the afterbody wedge shape is realized. The application is mainly applied to the scheme design of the afterbody wedge shape of the variable tail wing. The method is novel, ingenious and compact in structure, has strong applicability, has a wide application range, and has a wide application prospect in the design of military aircraft.

The invention belongs to the technical field of aircrafts, and aims to solve the problems of resistance divergence effect and difficult engine air intake of a traditional aircraft. The air-breathing high-subsonic-velocity aircraft comprises a nose, a fuselage, a tail, wings and an empennage, the nose is connected with the tail through the fuselage, the wings and the empennage are connected with the fuselage, the wings are supercritical airfoil type sweepback wings, an embedded air inlet channel is formed in the lower surface of the tail and communicated with an engine in the tail, and the engine is connected with the tail through the embedded air inlet channel. The embedded air inlet channel can introduce air on the lower surface of the tail into the engine so as to meet the air inlet requirement of the engine. According to the aircraft, the wings are arranged to be the supercritical wing type sweepback wings, through the unique geometric shape and sweepback arrangement, the resistance divergence phenomenon is effectively delayed, the effective speed of airflow perpendicular to the front edge is reduced, accordingly, shock wave generation is delayed, an embedded type air inlet channel is formed in the lower surface of the tail, and the impact resistance of the aircraft is improved. The air inlet requirement of an engine of an aircraft is met, and meanwhile the flight resistance and the radar cross section are reduced.

An aircraft (10) comprising: a starboard swept wing (14A) and a port swept wing (14B), each swept wing having an inner wing section (26A) comprising an integrated payload fuselage (20) and an outer wing section (30A) for extending the wingspan and providing lift; the inner wing section comprising: a rear spar (40) disposed close to a trailing edge (22A) of each inner wing section to carry loads along each inner wing section; a wing surface (30), in which the wing surface is substantially shaped as an aerofoil, and in which the wing surface comprises a rear fairing (62) disposed behind the rear spar, and near to the trailing edge, and in which the wing surface comprises an upper curved portion (50) which has a substantially arcuate cross section, a front curved portion (52) which has a substantially arcuate cross section, and a lower wing surface, the upper curved portion meeting the rear spar at a rear of the upper curved portion, and the upper curved portion meeting the front curved portion at a front of the upper curved portion, and the front curved portion meeting the lower wing surface at a front of the lowerwing surface, a cabin structure (32) in each inner wing, in which the cabin structure comprises one or more floor beams (36) and a rear wall defined by the rear spar; in which the lowerwing surface comprises a front lower curved portion (117), in which the front lower curved portion is substantially arcuate in profile and extends from a front edge of the floor beam(s) of the cabin structure towards the rear wall; in which the lower wing surface of each inner wing extends below a first line A, the first line being a line on the rear wall of the cabin structure where a projection of the arcuate front lower curved portion of the wing surface meets the rear wall, and in which a rear lower portion of the lower wing surface meeting the front lower portion and continuing the lowerwing surface to a second line, which is spaced from and below the first line.

The utility model discloses a fixed-wing unmanned aerial vehicle based on flexible perovskite solar cells, which adopts a wing body fusion design and comprises an equipment compartment in the middle of a vehicle body and wings on two sides, and a pair of ailerons with rolling and pitching control functions is arranged on the rear edge of the outer section of each wing; a pair of vertical stabilizers are arranged at the tips of the wings so as to enhance the course stability during flight; two flexible perovskite solar cells which accord with the shapes of the wings and are tightly attached to the wings are mounted on the surfaces of the fuselage and the wings, and power is provided for the unmanned aerial vehicle. According to the utility model, the sweepback wing flying wing type layout is adopted, so that the effective lift area of the wings and the laying area of the solar cells are increased while the structural weight is reduced; the battery is laid on the surface of the wing, damage to the battery in the laying process can be reduced, and the installation difficulty is lowered; meanwhile, the flexible perovskite solar cell has the advantages of low cost, good folding and bending performance and the like, so that the flight efficiency of the unmanned aerial vehicle is improved, the effective load is increased, and the endurance is prolonged.

An aircraft (10) comprising: a port swept wing (14B) and a starboard swept wing (14A), each swept wing having an inner wing section (26A) comprising an integrated payload fuselage (20) and an outer wing section (30A) for extending the wingspan, in which each outer wing section has a thickness less than a thickness of the corresponding inner wing section; and in which the sweep angle of each outer wing section is lower than the sweep angle of the corresponding inner wing section; a central region (12) where the port swept wing meets the second swept wing; a port transition region (23A) where the port inner wing section meets the port outer wing section; a starboard transition region (23B) where the starboard inner wing section meets the starboard outer wing section; at least one rear spar on each wing, in which each rear spar is disposed close to a trailing edge (22A) of the wing to carry loads along the respective wing; one or more elevons (34A) or ailerons on each outer wing section; and one or more roll spoilers (32A) on each inner wing section.

The application provides a method for assigning aerodynamic heat boundary conditions of variable sweep wing structure heat transfer simulation, which comprises defining a wing surface position with a first sweep angle in a whole trajectory as a reference position, rotating heat environment data to the reference position, and dividing a structure heat transfer grid at the reference position; then, creating a "point method" equation for the outer edge of a wing box, establishing a structure heat transfer grid surface grid point space position criterion based on the equation, and dividing the grid points into a set of aerodynamic heating points and a set of non-aerodynamic heating points; taking the set of aerodynamic heating points as an aerodynamic heat data interpolation target area, performing heat environment interpolation, and assigning 0Kw / m 2 to the cold wall heat flux of the set of non-aerodynamic heating points, restoring the constant value of enthalpy, and forming an adiabatic boundary condition. The method solves the problems that aerodynamic heat environment data cannot be directly interpolated to the structure heat transfer calculation grid surface due to the change of the spatial position of the variable sweep wing structure, and the structure heat transfer grid spatial position changes, resulting in that heat transfer analysis cannot be continuously performed along the whole trajectory.

An aircraft 10 has a pair of swept wings 14A,14B meeting at a central region 12. Each wing has an inner wing section 32A,32B containing an integrated payload fuselage; an upper surface and a lower su

This invention discloses a variable-sweep wing and an aircraft. The variable-sweep wing includes an external linkage system based on a multi-link mechanism, an internal linkage system based on a crank-slider mechanism, and a power assembly. The external linkage system is hinged to the fuselage of the aircraft to form the wing body. The internal linkage system is also hinged to the fuselage to maintain the attitude of the wing body formed by the external linkage system. The internal linkage system is connected to the external linkage system via at least one intermediate link, forming a single-degree-of-freedom planar multi-link structure with multiple hinges. The power assembly can drive changes in the attitude of each link in the external and internal linkage systems, thereby altering the sweep angle of the wing body. This invention has the advantages of a simple driving method and more uniform overall force distribution.

Wing assembly with slats and aircraft. A wing assembly comprising: a swept wing body, a leading edge of the wing body extending outwardly and rearwardly from a wing root to a wing tip; a first slat selectively movably connected to the wing body; and a second slat selectively movably connected to the wing body, the second slat disposed outwardly of the first slat, a flexible seal member disposed and connected between the first slat and the second slat; at least a portion of the first slat, at least a portion of the second slat, and at least a portion of the flexible seal member defining therebetween a slat gap, at least a majority of the slat gap being substantially parallel to a predetermined local airflow direction. Also disclosed is an aircraft comprising a fuselage; and two oppositely disposed wing assemblies connected to the fuselage.

The application provides a low-subsonic to high-subsonic wide-speed-range deformable aerodynamic layout of a cruise missile, comprising double-axis variable sweep wing mechanisms arranged at the bottom of a fuselage, capable of arbitrarily changing and locking a sweep angle in a range of 0-90 degrees, each double-axis variable sweep wing mechanism comprising an inner wing segment and an outer wing segment, the inner wing segment being pivotally connected with the fuselage at a wing root, and the inner wing segment being adjusted at a front sweep angle in a range of 0-90 degrees; the outer wing segment being pivotally connected with the inner wing segment at a wing tip, and the outer wing segment being adjusted at a rear sweep angle in a range of 0-90 degrees; and the nose being blunt. In the embodiment, the double-axis variable sweep wing mechanism has opposite influence trends of the inner wing segment and the outer wing segment on the center of pressure of the cruise missile during stretching, folding and flying, so as to control the change range of the center of pressure of the cruise missile in a very small interval, and make the wing sweep process have good maneuvering and stability characteristics and flight stability. In addition, the double-axis variable sweep wing mechanism is parallel to the fuselage when being folded, so that the cruise missile is convenient to store and transport, and the storage and deployment capability is improved.

The sub-span and super-span speed domain aerodynamic layout comprises double-shaft variable-sweep-wing mechanisms arranged at the bottom of a fuselage, the sweep angle can be freely changed and locked within the range of 0-90 degrees, each double-shaft variable-sweep-wing mechanism comprises an inner wing section and an outer wing section, the wing root of each inner wing section is in pivot connection with the fuselage, and the wing root of each outer wing section is in pivot connection with the fuselage. The sweep-forward angle of the inner wing section is adjusted within any angle of 0-90 degrees; the wing root of the outer wing section is pivotally connected with the wing tip of the inner wing section, and the sweepback angle of the outer wing section is adjusted at any angle of 0-90 degrees; the machine body is in a flat shuttle shape, and the machine head is in a pointed-end shape. According to the implementation mode, in the stretching, folding and flying processes, the influence trends of the double-shaft variable-swept-wing mechanism on the pressure center focus of the patrolling bomb are opposite, so that the pressure center focus change range of cross-speed domain deformation flight is controlled within an extremely small interval, and the pneumatic control stability is remarkably improved. In addition, when being stored, the double-shaft variable-swept-wing mechanism can be completely folded and parallel to the fuselage and is adaptive to narrow spaces such as a launch canister or a transport case, so that the storage and deployment capabilities are improved.

The application discloses a method and device for longitudinal aerodynamic force rapid analysis and water-exit attitude stability control of a variable-sweep wing trans-medium aircraft and electronic equipment, and comprises the following steps: aerodynamic analysis is carried out on the wings, fuselage and elevators of the aircraft separately; the aerodynamic force analysis of the wings is carried out by using open-source aerodynamic analysis software XFLR5 software, and the results of the aerodynamic force analysis of the wings in the multi-physical field simulation and computational fluid dynamics software STAR-CCM+ software are compared and verified; a mathematical model of the aircraft is established, and the control problem of the aircraft is mathematically abstracted; an extended state observer (ESO) is designed to observe the disturbance and the nonlinearity of the model, and a backstepping controller is designed to control the attitude of the variable-sweep wing trans-medium aircraft, so that the aircraft is rapidly stabilized. The application can significantly improve the aerodynamic parameter acquisition efficiency, accelerate the theoretical verification process, and effectively solve the stability control of the variable-sweep wing trans-medium aircraft in the water-exit stage.

The embodiment of the present application provides a main beam (3) of a curved variable sweep wing. The main beam (3) of the curved variable sweep wing is generated by combining a flexible leaf spring and a flexible hinge. According to the embodiment of the present application, on the basis of the leaf spring, a flexible cross hinge geometry feature is added, so that the main beam (3) has good large deformation stiffness decoupling performance keeping ability, and is suitable for the curved variable sweep wing.

The invention discloses a variable sweepback wing unfolding mechanism based on self-locking type high-pressure nitrogen springs. The variable sweepback wing unfolding mechanism comprises a projectile body, sweepback wings, the nitrogen springs and control electromagnetic valves. A control electromagnetic valve is arranged at the other end of the middle part of the projectile body; the nitrogen spring is arranged between the sweepback wing and the control electromagnetic valve, one end is hinged to the sweepback wing, and the other end is hinged to the electromagnetic valve; one end of the sweepback wing is connected with the projectile body, and the other end extends outwards; and the electromagnetic valve is mounted on the outer side of the bomb body. According to the invention, the backswept wing can be unfolded and locked at any target angle, and the limitation that a traditional unfolding mechanism can only be fixed at a single position is avoided, so that the aerodynamic configuration on-line adjustment capability of the projectile body backswept wing is more flexible and stable.

The vertical take-off and landing composite wing unmanned aerial vehicle comprises a vehicle body, front wings are installed on the two sides of the front portion of the vehicle body respectively, the front wings are sweepback wings, a vertical tail wing is installed on the rear top of the vehicle body, and rear wings are installed on the top of the vertical tail wing. Wingtips at two ends of the rear wing are respectively connected with wingtips of the two front wings through two end plates, vertical take-off and landing propellers are mounted at the bottoms of the front wings and the rear wings, and a level flight propeller is mounted at the tail of the fuselage. According to the technical scheme, resistance can be reduced, stability is improved, and then the loading capacity and cruising ability of the unmanned aerial vehicle are improved.

An aircraft (10, fig 1) comprising:a starboard and port swept wing (14A, 14B) , comprising an integrated payload fuselage (20);inner wing sections (26A): a rear spar 40 disposed close to a trailing ed

The invention provides a double-shaft variable-sweep upper single-wing layout of a speed-domain-crossing patrolling missile, which comprises a double-shaft variable-sweep wing mechanism arranged on the back of a fuselage and can form an M-shaped integral structure; each double-shaft variable-swept-wing mechanism comprises an inner wing section and an outer wing section, and the wing root of the inner wing section is pivotally connected with the fuselage, so that the sweep-forward angle of the inner wing section is adjusted; the wing root of the outer wing section is pivotally connected with the wing tip of the inner wing section, so that the sweepback angle of the outer wing section is adjusted. In the embodiment, the angles of the inner wing section and the outer wing section can be independently changed, the left side and the right side have asymmetric deformation capacity, and the overall aerodynamic configuration is changed to adapt to different flight speeds from low subsonic velocity to supersonic velocity. The upper single-wing layout is adopted, and space can be released for placement of an air inlet channel and task loads. The high-aspect-ratio wings bend and deform upwards to deviate from the fuselage under the aerodynamic load, and the problem of structural interference of the wings and the fuselage caused by aerobombs is solved. Most airfoils and the whole variable-sweep rotating mechanism are located in the leeward shadow area of the fuselage, so that the stealth performance is improved, and the aerodynamic heating problem of the deformed airfoils during hypersonic flight is relieved.

This invention provides a plasma device and method for de-icing swept wings, including a plasma power supply and a plasma actuator. The plasma actuator is a surface dielectric barrier discharge plasma actuator, comprising an insulating dielectric layer, exposed electrodes, and buried electrodes. This invention uses a high-voltage AC surface dielectric barrier discharge plasma actuator for de-icing, which has advantages such as small size, light weight, low power consumption, fast response, no mechanical parts, and high heating efficiency. Furthermore, by utilizing the aerodynamic effect of the plasma actuator, energy can be injected into the airfoil boundary layer, delaying surface flow separation and improving the aerodynamic performance of the wing.