125 results about "Wingspan" patented technology

The invention discloses a deformable wing with the characteristics of large unfolding-folding ratio and high bearing capacity. The deformable wing comprises a bearing structure, a driving mechanism, a wing surface dimensional structure and an accessory structure, the bearing structure comprises a plurality of groups of telescopic mechanisms, wing ribs and corner boxes; the accessory structure comprises a mounting plate, and the mounting plate is connected with the machine body; each group of telescopic mechanisms comprises a plurality of stages of telescopic rods which are arranged along the chord length direction and are fixed on the mounting plate; the wing ribs are arranged in the wingspan direction and are connected with the multi-stage telescopic rods through the angle boxes to form a wing framework; the driving mechanism penetrates through the hollow parts of the wing ribs, and two ends of the driving mechanism are respectively connected with the mounting plate and the wing rib on the outermost side of the wing; the wing skeleton is externally coated with the wing surface dimensional structure to form a wing shape.

The invention relates to the technical field of unmanned aerial vehicles, in particular to an inflatable wing type variable wingspan unmanned aerial vehicle. The unmanned aerial vehicle comprises a vehicle body, variable inflatable wings are arranged on the two sides of the vehicle body, the variable inflatable wings are connected with an inflation and deflation mechanism arranged in the vehicle body, each variable inflatable wing comprises a hollow fixed inflatable wing body and an extended inflatable wing body, and an inflatable part of each extended inflatable wing body is arranged outside a hollow cavity of the corresponding hollow fixed wing body; the uninflated part of the extended inflatable wing body is arranged in the hollow cavity, and the extended inflatable wing body is connected with the telescopic part of the extended driving mechanism. According to the inflatable wing type variable wingspan unmanned aerial vehicle, the variable inflatable wings are adopted, the wingspans of the whole wings are adjusted by lengthening the inflatable wing bodies, the unmanned aerial vehicle adapts to different flight states, and the unmanned aerial vehicle is of a relative folding structure, so that multi-stage adjustment is achieved.

An aircraft wing tip that is movable by an actuator between a flight configuration, and a ground configuration which has a shorter wingspan for use in ground-based operations. Lugs that connect to arms of the actuator are integrally formed as part of the primary structure of either the fixed wing or outboard wing section, such as a folding tip section.

Disclosed is a wing assembly for an aircraft. The wing assembly comprises a fixed wing, a wing tip device, a wiring harness and a harness guide. The wing tip device is moveably mounted at a joint at an end of the fixed wing. The wing tip device is moveable about the joint between: (i) a flight configuration for use during flight, and (ii) a ground configuration for use during ground-based operations, in which ground configuration the wing tip device is moved relative to the fixed wing such that the span of the wing is reduced. The wiring harness extends between the fixed wing and the wing tip device. The wiring harness comprises a plurality of conductors and is arranged to transmit electrical power and / or data to the wing tip device. The harness guide is located between the fixed wing and the wing tip device. The harness guide comprises a helical channel configured to receive the wiring harness and guide the wiring harness between the fixed wing and the wing tip device.

A method for reducing aerodynamic loads on an aircraft located on the ground includes moving a wing tip relative to a fixed wing to an on-ground configuration in which the span of the wing is reduced. The wing tip includes an airflow channel, which is in an open configuration in which airflow through the channel between upper and lower surfaces of the wing tip is enabled, when the wing tip is in the on-ground configuration. This has been found to reduce aerodynamic loads on the wing tip and / or parts of the aircraft connected thereto during this phase of the aircraft's use.

A vertical take-off and landing aircraft includes a fuselage, at least one propulsion system, at least four lift propellers and at least two fixed wing planes. The fixed wing planes which are located behind the most forward lift propellers are located above the lift propellers. The lift propellers are distributed on either side of at least one of the fixed wing planes and on either side of the fuselage such that the lift propellers are longitudinally separated by at least the magnitude of the chord of the fixed wing plane located therebetween, and are laterally separated by at least the width of the fuselage located therebetween. At least one of the fixed wing planes has a variable span and the aircraft includes a control system for varying this span in flight.

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 multi-mode self-rotating tailstock type vertical take-off and landing aircraft. Multi-mode motion comprises a multi-axis mode, a self-rotating mode and a fixed wing mode. The unmanned aerial vehicle comprises a fuselage, a variable collective pitch structure and four symmetrical wings on the two sides; the wing type of each wing is a symmetrical wing type, the tail end of each wing is a full-moving wing tip, and each wing is provided with a driver; the driver is a propeller driven by a motor; each wing, the driver and each full-motion wingtip can rotate by + / -90 degrees around a rotating shaft in the spanwise direction of the wing, and switching among multiple modes and attitude control are achieved. The aircraft disclosed by the invention has more flexible maneuvering capability, has vertical take-off and landing performance of a multi-axis mode, efficient hovering performance of a spinning mode and high-speed cruising performance of a fixed wing mode, can meet different flight requirements, and has better environment adaptability.

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.

A wing structure for a supersonic aircraft including a pair of supersonic wave-tripping channels formed on each of two laterally extending wings of the supersonic aircraft, wherein each of the pair of supersonic wave-tripping channels extend in a span-wise direction of the wings respectively, wherein an upper supersonic wave-tripping channel of the pair of supersonic wave-tripping channels is disposed on an upper surface of each of the wings and a lower supersonic wave-tripping channel of the pair of supersonic wave-tripping channels is disposed on a lower surface of each of the wings, wherein the upper supersonic wave-tripping channel and the lower supersonic wave-tripping channel are set back from a leading edge of the wings respectively.

The application discloses a variable wing based on a metamorphic mechanism and an aircraft. The variable wing comprises a plurality of ribs, a metamorphic module and a driving mechanism. The plurality of ribs are sequentially connected by the metamorphic module in the transverse direction to form a wing body. The metamorphic module comprises two groups of multi-link metamorphic mechanisms. Each multi-link metamorphic mechanism comprises a first link and four second links. The four second links are arranged in two groups respectively on opposite sides of the first link. One end of each second link is hinged to the first link, and the other end of each second link is hinged to a spar of a rib. The first link, the two spars and the four second links form a double parallelogram structure sharing the first link. The driving mechanism is configured to act on the two groups of multi-link metamorphic mechanisms respectively to drive the wing body to switch between a variable-span configuration, a singular configuration and a variable-sweep configuration. The application can realize both wing span change and sweep angle change, and has the advantages of compact structure and good stability.

This application provides a method, apparatus, device, and medium for determining aerodynamic characteristic parameters of an aircraft, relating to the field of data processing technology. The method acquires multiple airfoil parameter sets, as well as the airfoil shape parameters and control parameters of the aircraft under test. Based on the wingspan of the aircraft under test and the wingspan range of each airfoil parameter set, it determines the target airfoil parameter set corresponding to the aircraft under test. Based on the airfoil shape parameters of the aircraft under test and the airfoil shape parameters of the target airfoil parameter subset, it determines a reference subset from the multiple target airfoil parameter subsets. Based on the correspondence between the control parameters of the aircraft under test and the control parameters and aerodynamic characteristic parameters in the reference subset, it determines the aerodynamic characteristic parameters of the aircraft under test. By finding a similar reference subset from multiple known airfoil parameter sets and obtaining the aerodynamic characteristic parameters of the aircraft under test from the reference subset, the efficiency of determining the aerodynamic characteristic parameters of an aircraft can be improved.

A vertical take-off and landing aircraft (1) comprises a fuselage (2), at least one propulsion system (5), at least four lift propellers (10), and at least two fixed wing planes (20, 30, 40), the fixed wing planes (30, 40) located rearward of the forwardmost lift propellers (10) being located above the lift propellers (10). The lift propellers (10) are distributed on either side of the fixed wing planes (30) and on either side of the fuselage (2) such that the lift propellers (10) are separated in the longitudinal direction by at least the size of the chord (Cl, C2) of the fixed wing planes located between the lift propellers (10) and in the lateral direction by at least the width (L1, L2) of the fuselage (2) located between the lift propellers (10). At least one of the fixed wing planes (30) has a variable wingspan and the aircraft (1) comprises a control system for varying the wingspan in flight.

The application discloses a kind of torsion wing structure and aircraft, it is related to the field of morphing wing technology, including linear drive device, wing rib and pressure-torsion coupling cell, wing rib and pressure-torsion coupling cell are multiple, multiple wing ribs are arranged along the wing span direction interval, and pressure-torsion coupling cell is equipped between two adjacent wing ribs;Pressure-torsion coupling cell includes intermediate connecting component and the load-bearing end block fixedly connected with the two ends of intermediate connecting component, load-bearing end block is fixedly connected with adjacent wing rib, and all wing ribs and pressure-torsion coupling cell form wing main body;The first end wing rib of wing main body is used to be fixedly connected with fuselage;Two load-bearing end blocks of pressure-torsion coupling cell are connected with linear drive device;Linear drive device can make the two load-bearing end blocks of each pressure-torsion coupling cell approach each other or away from each other, to make intermediate connecting component torsion deformation and drive wing rib around wing span torsion by load-bearing end block.The application has the effect of "lightweight-structure simplification-aerodynamic efficiency-high life-low maintenance cost".

The invention relates to the technical field of computer simulation, and discloses a rapid modeling simulation method and system for a deformed wing structure, and the method comprises the steps: recognizing wing section key points of wing section data, fitting an upper surface curve and a lower surface curve of the deformed wing structure, and determining a parameterized wing section of the deformed wing structure; analyzing key change parameters, corresponding to the wingspan direction, of the deformed wing structure so as to construct a parameter change curve in the wingspan direction, and fitting a wing geometric model of the deformed wing structure; determining a deformation mode of the deformed wing structure, calculating an airfoil variable quantity and a torsion angle variable quantity of the deformed wing structure in a deformation process, and fitting an initial deformed wing model of the deformed wing structure; carrying out simulation calculation on an aerodynamic coefficient of the initial deformation wing model, and determining aerodynamic performance of the initial deformation wing model; and optimizing the initial deformation wing model to obtain a target deformation wing model. According to the method, the modeling simulation accuracy and design efficiency of the deformed wing structure can be improved.

The invention discloses a cross-medium variable tilt rotor aircraft and a control method thereof, and relates to the technical field of cross-medium aircrafts. Comprising a fuselage, front rotor units, a tail tilting rotor assembly, main wing units and wing unfolding and folding mechanisms, wherein the front rotor units are arranged on high-position mounting platforms on the two sides of the fuselage and are in bilateral symmetry; the tail tilting rotor assembly is arranged at the tail end of the fuselage and can rotate around a transverse tilting shaft; the main wing units are in bilateral symmetry and can be folded; the buoyancy adjusting unit and the control system are arranged in the fuselage. Through collaborative design of fuselage bionic integration, the composite rotors, the variable main wings, active buoyancy adjustment and closed-loop control, the traditional single propulsion bottleneck is broken through, cross-medium adaptation, consideration of take-off and landing and cruising efficiency and smooth transition are achieved, multi-scene adaptation is achieved, and breakthrough support is provided for engineering application.

The utility model discloses water conservancy and hydropower hoisting equipment suitable for materials of different specifications, and relates to the technical field of hoisting equipment, the water conservancy and hydropower hoisting equipment comprises a hoisting tool body, a sling body and a wire harness, hoisting assemblies are arranged at the two side end parts of the hoisting tool body, each hoisting assembly comprises a wingspan bearing plate, a hook seat and a pressure-bearing stand column, adjusting assemblies used for being connected with a wingspan bearing plate are arranged at the two side ends of the lifting appliance body, each adjusting assembly comprises a bracket, a reinforcing stand column and an electromagnetic block, a bolt at the top end of a hook base is screwed, the hook base slides on the outer side of the wingspan bearing plate, the position of a lifting point is adjusted, and the wingspan bearing plate can be lifted. Meanwhile, the wingspan bearing plate slides at the side end part of the lifting appliance body and extends towards the two sides, the hook seat is aligned with the outer side of the equipment according to the volume of the equipment after the hook seat is moved and adjusted, and the hook and the gravity center of an object can be on the same vertical line by adjusting the position of the hook seat, so that the object is kept balanced in the lifting and hoisting processes.

A temperature monitoring unit may be removably installed inside an aircraft wing structure for monitoring temperature conditions along the span of the wing. The wing structure has a temperature-sensitive device (162) for monitoring a temperature con- dition, which is attached to a support frame (173). The support frame and attached temperature-sensitive device may be installed as a unit within the wing structure. The support frame may be configured for sliding engagement inside the wing structure, for example, with a set of tracks.

This invention discloses a variable cross-section tail wing suitable for FSAE racing cars, comprising a variable cross-section wing assembly, an upper sparsity wing, a first lower sparsity wing, a second lower sparsity wing, and endplates. The variable cross-section wing assembly consists of a main wing, a first flap, and a second flap; all three wings are variable cross-section wings. The variable cross-section wing is a small, continuous section raised at mid-span. The main wing, the first flap, and the second flap are sequentially mounted between the two endplates, roughly diagonally distributed, with the second flap featuring a grunt flap. The upper sparsity wing is mounted on the corner of the endplate, while the first and second lower sparsity wings are located near the bottom of the endplate. There are two endplates, symmetrically mounted on both sides, each featuring louvers, an outer edge flange, a leading edge notch, a trailing edge notch, and guide strips. The variable cross-section wing can reduce the turbulence caused by the headrest and minimize airflow slippage in the spanwise direction, resulting in a more uniform pressure distribution on the wing surface. The louver structure also reduces drag.

The invention discloses a vertical take-off and landing aircraft, and belongs to the field of aircrafts. The vertical take-off and landing aircraft comprises a fuselage, a first fixed wing, a second fixed wing and a plurality of rotor wing mechanisms. And the first fixed wing and the second fixed wing are distributed at intervals in the first direction. The number of the first fixed wings and the number of the second fixed wings are both two. The two first fixed wings are distributed on the two sides of the fuselage in the second direction and connected with the fuselage. The two second fixed wings are distributed on the two sides of the fuselage in the second direction and connected with the fuselage. And each first fixed wing and each second fixed wing are connected with at least one rotor wing mechanism. The first fixed wing and the second fixed wing are distributed in a staggered mode in the third direction. According to the vertical take-off and landing aircraft, through the first fixed wing and the second fixed wing which are arranged in series, on the premise that the same load capacity is achieved, the wingspan corresponding to the fixed wing on the single side of the aircraft body can be effectively shortened, and the space adaptability and the structural rigidity of the aircraft are improved.

A method of actuating a wing tip device (11, figure 1B) on an aircraft (1), from a flight configuration to a ground configuration in which the span is reduced. The method comprises the steps of 100 receiving data indicative of a condition, such as the location 23 or speed (124, figure 5) of the aircraft; and determining an actuation profile 102 in dependence on the condition. The actuation profile may include determining a time period, and / or speed profile of the wing tip actuation over that time period. In this manner the actuation of the wing tip device can be tailored to minimize loads and wear on the actuator assembly.

An unmanned aircraft according to the present disclosure is provided with an airframe (2) and a main wing (4). The main wing (4) has: a pair of left and right wing body members (12) that are connected to the airframe (2) at the wing root side, and extend in the wing width direction toward the wing tip side; and at least one movable wing member (14) that is provided at the trailing edge of the wing body members (12), and can be displaced relative to the wing body members (12). The at least one movable wing member (14) has a wing root-side portion disposed at the wing root side, and a wing tip-side portion disposed at the wing tip side. The wing tip-side portion has a smaller lift coefficient than the wing root-side portion.

The application provides a kind of aircraft with telescopic folding wing, it relates to aircraft technical field.The aircraft includes: the cavity is formed in the main wing mechanism inside chamber, and the opening is arranged at the wing tip, the aileron mechanism and the main wing gear group are arranged in the chamber, so that the wing body of aircraft becomes extensible design;Aileron outer shell is connected with the worm transmission of main wing gear group through worm nut;Power mechanism is arranged on the fuselage, and power mechanism is used to control main wing mechanism to unfold, retract;When power mechanism moves main wing outer shell rotation, through the transmission of main wing gear group, so that aileron outer shell is stretched out or shrunk from the wing tip opening of main wing outer shell.In the case where the length, width of fuselage and the space occupied after wing folding are unchanged, the wingspan is increased and it is convenient to recycle and reuse, effectively improves the range of aircraft;And, when power mechanism provides driving force, the main wing outer shell of both sides of fuselage always keeps synchronous rotation, guarantees the stability of fuselage.

A system is provided for an aircraft. This aircraft system includes an aircraft and an air system. The airframe includes a body and a vane connected to the body. The vane projects spanwise away from the body to a vane tip. The vane extends longitudinally between a leading edge and a trailing edge. The vane extends laterally between a first vane side and a second vane side. The air system includes a circuit inlet, a circuit outlet and an air circuit. The circuit inlet is arranged with the body and laterally offset from the vane. The circuit outlet is arranged with the vane. The air circuit extends within the airframe from the circuit inlet to the circuit outlet.

The invention discloses a catapult blade variable wingspan rotor determining method and device, a medium and a product, and relates to the field of unmanned rotorcrafts. The method comprises the steps that an aerodynamic configuration three-dimensional model of a catapult blade variable wingspan rotor is constructed; according to the aerodynamic configuration three-dimensional model of the catapult blade variable wingspan rotor, establishing a rotor aerodynamic lift numerical model of a variable wingspan process from a conventional mode to a full wingspan mode; according to the rotor aerodynamic lift numerical model, determining the rotor aerodynamic lift change in the variable wingspan process from the conventional mode to the full wingspan mode; according to the change of the aerodynamic lift force of the rotor wing, determining the action relationship of the size and the rotating speed of the variable-wingspan rotor wing of the catapult blade on the aerodynamic lift force by combining fluid mechanics, fluid-solid coupling and boundary layer flow theories; and optimizing the size of the variable-wingspan rotor wing of the ejection paddle according to the action relationship. The aerodynamic lift of the aircraft can be rapidly and remarkably improved.

The application discloses a vertical take-off and landing fixed-wing aircraft based on thrust vector nozzles, which comprises four engines, four thrust vector nozzles, a wing, a fuselage, a T-shaped tail wing and a support; four sets of power devices adopt a four-point isosceles trapezoidal symmetrical layout, and front and rear power devices are symmetrically arranged about the gravity center of the whole machine. The transverse spacing of the front power device is 2.5-3 times that of the rear power device, the spanwise projection is located at 40%-60% of the wing span length, and the height is matched with the wing leading edge to realize jet flow lift increase; the height and horizontal position of the rear power device are optimized to ensure stable air intake. The thrust vector nozzle takes the horizontal axis of the fuselage as the 0° reference, the vector angle adjustment range is not less than -95°-+15°, and the hovering, mode transition and steady switching of the flat flight are realized through differential adjustment. The application cancels the tilting mechanism, reduces the structural dead weight, reduces the aerodynamic interference, and improves the cruising speed and flight safety.

Herein is disclosed a blade assembly (1) for providing wind propulsion for a ship (2), as well as such ship (2) having one or more blade assemblies (2), the blade assembly (1) comprising a top end part and a lower end part for being coupled to a part of the ship (2), and a first and a second aerofoil assembly (3, 4), each of the and the second aerofoil assembly (3, 4) comprising at least two aerofoils (5a, 5a', 5b, 5b', 5c, 5c', 105a, 105a', 105b, 105b', 105c, 105c'), wherein the first aerofoil assembly (3) is arranged at one side of a central plane (CP) of the blade assembly (1) and the second aerofoil assembly (4) is arranged at the opposite side of the central plane (CP) than the first aerofoil assembly (3), wherein the at least two aerofoils (5a, 5a', 5b, 5b', 5c, 5c', 105a, 105a', 105b, 105b', 105c, 105c') of each of the first and the second aerofoil assembly (3, 4) are oriented with a leading edge (LE) of the aerofoil (5a, 5a', 5b, 5b', 5c, 5c', 105a, 105a', 105b, 105b', 105c, 105c') towards an inlet gap (IG) of the blade assembly (1) between the first aerofoil assembly (3) and the second aerofoil assembly (4), wherein the at least two aerofoils (5a, 5a', 5b, 5b', 5c, 5c', 105a, 105a', 105b, 105b', 105c, 105c') of each of the first and the second aerofoil assembly (3, 4) are arranged in an overlapping sequence so that one or more flow channels (FC) are formed between neighbouring aerofoils (5a, 5a', 5b, 5b', 5c, 5c', 105a, 105a', 105b, 105b', 105c, 105c'), and wherein for each of the first and the second aerofoil assembly (3, 4) an assembly camber line (ACAL) connecting the leading edge (LE) of the front aerofoil (5a, 5a', 105a, 105a') nearest the inlet gap (IG) with the trailing edges (TE) of each of the aerofoils (5a, 5a', 5b, 5b', 5c, 5c', 105a, 105a', 105b, 105b', 105c, 105c') in the aerofoil assembly (3, 4) from the front aerofoil (5a, 5a', 105a, 105a') towards the back aerofoil (5b, 5b', 105b, 105b') nearest an outlet gap (OG) of the blade assembly (1) deviates from an assembly chord line (ACHL) connecting the leading edge (LE) of the front aerofoil (5a, 5a', 105a, 105a') with the trailing edge (TE) of the back aerofoil (5b, 5b', 105b, 105b') in a direction with a component away from the central plane (CP), the blade assembly (1) further comprising a flow blocking arrangement (10, 11, 13, 14) which can be selectively shifted between a first position, where flow of air from the inlet gap (IG) and out through the flow channel(s) (FC) of the first aerofoil assembly (3) is limited, and a second position, where flow of air from the inlet gap (IG) and out through the flow channel(s) (FC) of the second aerofoil assembly (4) is limited.

This invention belongs to the field of aircraft technology, specifically disclosing a multi-dimensional continuously variable wing structure, including a fixed plate, an inner wing section, an outer wing section, a sweep actuation mechanism, and a telescopic actuation mechanism. The inner wing section is driven by the sweep actuation mechanism to rotate around the wing root in the spanwise plane to change the wing sweep angle. The outer wing section is slidably connected to the inner wing section and is driven by the telescopic actuation mechanism to extend or retract along the spanwise direction to change the wing span. The leading and trailing edges of the rear side of the inner wing section are respectively connected to a flexible trailing edge and a rigid trailing edge. The flexible trailing edge is used to change the trailing edge camber; the rigid trailing edge is used to change the trailing edge deflection angle. The sweep actuation mechanism includes a turntable, a push rod motor, and a rocker arm. The push rod motor drives the rocker arm to rotate the inner wing section around the turntable to achieve variable sweep. This invention can achieve multi-dimensional wing variability through a simple combination of mechanisms, improving the aerodynamic performance of the aircraft and enabling the aircraft to adapt to different flight conditions and meet various flight mission requirements.

The application discloses an underwater flapping wing structure and a bionic flapping wing underwater robot, and relates to the technical fields of underwater vehicles and bionic robots. The underwater flapping wing structure comprises a connecting member and a first skin, and the first skin is arranged on the connecting member. The connecting member is used for stretching and contracting in the span direction when deformed by water pressure. The first skin can stretch and contract with the deformation of the connecting member, and the chord width size of the first skin remains unchanged. The bionic flapping wing underwater robot comprises a base, a driving mechanism, a tail fin oscillation mechanism and the underwater flapping wing structure. The tail fin oscillation mechanism and the driving mechanism are arranged on the base, and the output end of the driving mechanism is connected with the connecting member. The first skin can realize adaptive stretching and contraction in the span direction, and the chord width size of the first skin remains unchanged, so that the robot can always maintain the most efficient standard hydrodynamic wing profile boundary in the process of stretching in the span direction and greatly changing the water sweeping area of the wing surface, the flexible variable wing span with low resistance flow is realized, and the propulsion efficiency is improved.