100results about How to "Good aerodynamic shape" patented technology

The bay (12) of a turbofan engine (10) comprises a front structural element (30), whose external surface is continuous and extends over at least 50% of the geometrical chord of the bay. Said element (30) is installed on maintaining and guiding members (44), such as slides, which prevent a significant deformation in flight and allow a sliding to the front of the element (30) for maintenance purposes. A laminar air flow around the front half of the bay (12) is consequently ensured.

A toroidal shaped or ring fuel tank located within the loft line of a blended wing body aircraft is disclosed. The ring tank may be used in an aircraft to store liquid hydrogen fuel with a reduced tank weight. The ring tank may be continuous with no tank end domes typically found on cylindrical pressure tanks, reducing tank weight for a given fuel volume. The ring tank configuration avoids increasing the aerodynamic shape of the aircraft and does not encroach on usable passenger or payload areas of the aircraft. In one example the ring tank may be configured in a nose down position such that the forward portion of the ring tank is outside the pressurized cabin area.

A blade for use in a wind turbine comprises a pressure side and suction side meeting at a trailing edge and leading edge, the pressure side having a pressure side surface and the suction side having a suction side surface, the pressure side surface and suction side surface for providing lift to the turbine blade upon the flow of air from the leading edge to the trailing edge and over the pressure side and suction side surfaces, the pressure side and suction side extending from a root portion to a tip portion of the turbine blade. In some situations, the root portion is non-aerodynamic. The aerodynamics of such a blade is improved with the aid of pneumatic blowing through one or more blown passages for providing pressurized air (or other fluid) to a suction side and / or pressure side of the blade.

The invention discloses a novel inorganic thermal protection structure and a manufacturing method thereof. The novel inorganic thermal protection structure comprises an aluminum alloy basal body, a heat insulation layer and a heat-protection layer, wherein the heat insulation layer is a space reticular structure with a plurality of staggered hole gaps, the staggered hole gaps comprise a hollow sleeve pipe braided by a plurality of layers of quartz fibre wound on the aluminum alloy basal body, and the heat-protection layer is a model body which is formed by pressing a silicon nitride powder body and is matched with the heat insulation layer. The internal diameter of the quartz fibre sleeve pipe is 0.5-1.0 mm. The quartz fibre sleeve pipe is wound by a wet method, and the distance of the sleeve pipe is 1-4 mm; the quartz fibre sleeve pipe comprises the following parameters from the inside to the outside: the internal ring direction winding layer thickness H 1 is 1-4 mm, the spiral winding layer thickness H2 is 20-35 mm, and the external ring direction winding layer thickness H3 is 5-20 mm; the winding parameter of the spiral winding layer thickness is 75-85 degrees, and the winding direction between layers is formed by opposite superposition. The thickness of the silicon nitride heat-protection layer is 2-4 mm.

The invention discloses a variable unmanned aerial vehicle. The variable unmanned aerial vehicle comprises a vehicle body (1), multi-section wings (2) and an empennage (3). The multi-section wings (2) are arranged on the vehicle body (1), the adjacent wing sections of the multi-section wings (2) can relatively stretch out and draw back, and therefore the total elongation of the multi-section wings (2) is changed. The empennage (3) is arranged on the vehicle body (1), and can be far away from or close to the vehicle body (1) relative to the flying direction of the unmanned aerial vehicle. All the wing sections of the variable unmanned aerial vehicle can relatively stretch out and draw back, and therefore the elongation of the wings is changed. When the wing sections relatively stretch out and draw back to change the pneumatic appearance of the variable unmanned aerial vehicle, the gravity center position and the pneumatic characteristic changed by stretching out and drawing back of the wings are adjusted by enabling the empennage to be far away from or close to the vehicle body (1) relative to the flying direction of the unmanned aerial vehicle, and therefore the pneumatic appearance of the variable unmanned aerial vehicle is kept at the optimal state whether the variable unmanned aerial vehicle is in any state.

The invention discloses a split streamline steel-concrete composite box girder comprising two or three horizontal steel-concrete composite box girder units. An air nozzle is disposed outside each steel-concrete composite box girder unit. A channel 4m-20m in horizontal width is formed between each two steel-concrete composite box girder units. Connecting beams are disposed at intervals of 4m-20m longitudinally along the channels. A top plat of each steel-concrete composite box girder unit is a concrete plate, and a bridge deck pavement for automobiles is paved on the top plate; a bottom plate and an oblique web in each girder unit are reinforced steel plates; rails for trains or light rail vehicles are arranged on the bottom plate; each oblique web is provided with round, elliptical or polygonal light holes. The split streamline steel-concrete composite box girder has the advantages that wind proofing stability of bridges can be improved effectively, vortex-induced vibration is suppressed, durability of the bridge deck pavement is improved, and the girder is applicable to highway bridges or highway and light-railway bridges and is simple in structure and well economical.

The invention relates to an inflatable wing with a controllable sway trailing edge and is characterized in that: the inflatable wing adopts the structural design of split skin lacing wires and comprises a main wing, an auxiliary wing, a connection skin and a piezoelectric fibrous composite unit, wherein the main wing and the auxiliary wing are both of inflatable air bag structure, and form a common wing-shaped curve; the auxiliary wing is connected with the main wing as a body through wire joint; the inner parts of air bags of the main wing and the auxiliary wing are provided with a pluralityof lacing wires that divide the wing into bays which are mutually communicated; two ends of the connection skin are respectively bonded with the main wing and the auxiliary wing the two of which are connected together; and the piezoelectric fibrous composite unit is attached to the connection skin and the bottom edge of the piezoelectric fibrous composite unit is level with that of the connectionskin. The inflatable wing effectively solves the flight control problem of inflatable wings.

The invention discloses a novel efficient tilt rotor unmanned aerial vehicle. Unmanned aerial vehicle layout has higher aerodynamic efficiency and better level-flight longitudinal direction and coursecontrol performance, and high-speed and high-efficiency flight can be achieved on the premise of vertical take-off and landing. The unmanned aerial vehicle comprises tilt rotors, vertical flight mechanisms, fixed wings, a fixed vertical fin, a stabilator, a body structure and a landing gear. The unmanned aerial vehicle needs to experience a vertical flight state, a transition flight state and a level flight state from take-off to level flight. The tilt rotors and the vertical flight mechanisms can rotate through an actuation mechanism, so that the unmanned aerial vehicle can vertically take off and horizontally fly at high speed, the tilt rotors can change a distance, take account of rotating efficiency in the vertical flight state and the level flight state and drive the unmanned aerialvehicle to fly at high speed, the vertical flight mechanisms and the tilt rotors collectively provide lift force and control force of vertical take-off and landing in the vertical flight state, the vertical flight mechanisms are stored on two sides of a body in the level flight state, and a good aerodynamic shape of the unmanned aerial vehicle is kept.

A variable camber wing based on composite corrugated plate flexible skins is composed of a plurality of wing sections, rotating shafts, hinge structures, composite corrugated plates and elastomer layers. The connection relationship is as follows: the wing sections are connected end to end through the rotating shafts, the composite corrugated plates cover between the adjacent wing sections, the composite corrugated plates are respectively fixed on platforms of the front wing section and the rear wing section, each hinge structure is arranged in the middle of each composite corrugated plate andis connected with the corresponding wing section, and the elastomer layers are adhered to the outer surfaces of the composite corrugated plates. The characteristic that the composite material corrugated plate flexible skins have large deformability in the corrugated direction is utilized to design the variable camber wing structure, out-of-plane constraint is applied to the composite corrugated plate flexible skins through the hinge structures, the wing can achieve the variable camber function, out-of-plane deformation outside the allowable range of the composite corrugated plate flexible skins can be avoided, and a good aerodynamic configuration can be maintained.

The present invention relates to airplane structure design, and is especially improved lip for intake duct of airplane. The improved lip for intake duct is one ring part comprising connected ring lip outer wall and ring lip inner wall, features that between the ring lip outer wall and the ring lip inner wall, there are reinforcing ribs distributed in the circumference direction and in the interval of 200-400 mm and that the lip is formed integrally with plate material through a machining process. The improved lip has high dimensional precision, good pneumatic appearance, greatly raised strength, great load bearing capacity and long service life.

The invention provides an unmanned aerial vehicle retardation releasing device. The unmanned aerial vehicle retardation releasing device comprises a head cover, a middle cabin, a tail skirt, a retarder parachute, a controller and a cutting rope. The middle cabin is a hollow cylinder, one end of the middle cabin is connected with the bottom face of the hemispherical head cover, the other end of themiddle cabin is connected with the upper bottom of the circular-table-shaped tail skirt, and an unmanned aerial vehicle is fixed to the interior of a cavity formed by the head cover and the middle cabin; and the retarder parachute and the controller are installed in the tail skirt, and the cutting rope is bonded to the inner surfaces of the head cover and the middle cabin and the joint of the middle cabin and the tail skirt along the axis of the middle cabin in an axially symmetrical mode. According to the unmanned aerial vehicle retardation releasing device, the unmanned aerial vehicle can be rapidly and effectively released from a parent aircraft, and successful take-off of the unmanned aerial vehicle can be ensured.

The present invention provides a steering gear installation fastener applied to a miniature unmanned aerial vehicle and an installation manner. A steering gear installation fastener is designed for installing a steering gear, which can install the steering gear to lie inside a stable face within limited space, is especially fit for the unmanned aerial vehicle which has high requirement towards a pneumatic appearance and realize the fast disassembly and assembly of the steering gear. When the steering gear is installed, firstly the steering gear is pressed and fastened to be blocked at the fastener, and then the steering gear is installed to lie inside a stable face by the connection of the fastener and a skin. A pressing connection fastening method is applied to the medium, small and miniature unmanned aerial vehicles to replace the prior bolt connection fastening, which can realize the convenient and fast disassembly and assembly of the steering gear, simplifies the installation and manufacture technology and reduces the weight and the cost.

The invention discloses a large-span foldable reusable rocket landing buffer structure, which comprises a main leg (1), two subsidiary legs (2) and a landing plate (3), wherein an upper joint at one end of the main leg (1) and an upper joint at one end of each subsidiary leg (2) are connected with a mounting seat of a rocket body through knuckle bearings arranged on the main leg (1) and the subsidiary legs (2) respectively, and the knuckle bearings are arranged on the upper joints and the lower joints of the main leg (1) and the subsidiary legs (2); a rotating shaft is arranged on one side ofthe landing plate (3); the lower joint at the other end of the main leg (1) and the lower joint at the other end of each subsidiary leg (2) are connected to the same rotating shaft on the landing plate (3) through knuckle bearings arranged on the main leg (1) and the subsidiary legs (2) respectively; and a rubber buffer pad is arranged on the other side of the landing plate (3) and can make buffercontact with the ground when the rocket lands slowly.

The invention provides a composite material blade for a wind tunnel. The composite material blade for the wind tunnel comprises a covering skin, a front edge reinforcement area, a rear edge reinforcement area, foam core materials, a main bearing beam and a metal blade handle, wherein the metal blade handle comprises a blade portion in square frame shape and a shaft portion in long rod shape, the blade portion is stretched into a central area in the covering skin in blade shape, the front edge reinforcement area and the rear edge reinforcement area are respectively arranged at the front end and the rear end of the interior of the covering skin, the covering skin and the blade portion are filled with the foam core materials, the square main bearing beam is embedded between the covering skin and the blade portion, the thickness of the main bearing beam is varied and decreases progressively in ladder mode from one side of the shaft portion, and the width of the main bearing beam is varied and decreases progressively in ladder mode from the side of the shaft portion. The composite material blade for the wind tunnel is made of carbon fiber composite material, and thereby is light and strong. Furthermore, the foam core materials have an effective supporting function for a wing-shaped profile surface of the composite material blade for the wind tunnel, and the main bearing beam bears a majority of loads on the composite material blade for the wind tunnel.

The invention provides a semi-open airship, which comprises a fixed rigid framework (1), a buoyancy-lifting gas bag (2), a converging skin (3), a converging device (6) and an air bag (7). By converging and diverging the converging skin (3), the effective buoyancy volume is adjusted, the situation that the air throughput of a fixed-displacement airship in ascending and descending processes is large is avoided, the burden of a gas management system is reduced and the mass and inertia change of the entire airship is reduced; through the converging skin (3) and the fixed rigid framework (1), the airship is enabled to have better aerodynamic configuration, the overall rigidity of the structure of the airship is guaranteed and the effects of structural adjustment of the airship and flying without depending on an external carrier are realized; the deformation of the converging skin (3) and a buoyancy-lifting gas bag skin (22) is realized mainly through folding and unfolding, tensile deformation theoretically does not exist, the requirement on fatigue resistance of skin and bag materials is reduced and the extension and retraction motion of the variable airship air bag in the prior art is structurally replaced.

The invention discloses a solid carrier rocket sub-stage capable of being recycled and reused. The solid carrier rocket sub-stage comprises a sub-stage body, an inflatable heat protection structure and a solid rocket engine; the inflatable heat protection structure comprises a front-end airbag which covers the windward side of a front cabin section under an expanding state and is in a shape of a rotating body, and the front-end airbag comprises a flexible heat-insulating layer and a heat-proof coating which are sequentially arranged from inside to outside; and the solid rocket engine comprisesa combustion chamber shell with a cavity allowing a grain to be placed, and a combustion chamber heat protection structure is arranged on the inner wall of the combustion chamber shell. By adopting the inflatable heat protection structure with heat protection, deceleration and buffering combined, the front-end airbag can maintain the good aerodynamic shape, the sub-stage is not damaged due to pneumatic heating and large landing impact force in the reentry process, and can be recycled, through re-charging, replacing of disposable parts and the like of the rocket engine, the sub-stage and the engine thereof can be reused, and thus the launching cost of a solid carrier rocket is lowered.

The invention discloses a moving target atmospheric disturbance characteristic simulation method. The method includes obtaining target geometry parameters, establishing target three-dimensional model,dividing grids of a target and flow field solving area; setting flow field calculation conditions, selecting a flow field calculation model, obtaining target atmospheric disturbance flow field parameters, calculating flow field refractive index distribution, calculating a refractive index gradient according to a Barron operator, calculating a disturbance atmospheric deflection transmission path through Runge-Kutta light tracing, and obtaining a deflection angle of light passing through a disturbance field. According to the method, simulation analysis is carried out on light deflection generated when ground object light passes through an atmospheric disturbance area formed by a high-speed target for the first time. The method can be used for atmospheric disturbance simulation analysis of any target flying in the atmospheric layer, can analyze the deflection condition of the ground object light of any spectral band passing through the target atmospheric disturbance area, and can also analyze the deflection condition formed by the ground object light of different spectral bands passing through the atmospheric disturbance area of the target at different incident angles. A new means isprovided for target atmospheric disturbance simulation analysis, and the method can be applied to atmospheric disturbance field characteristic research.

The invention relates to a mould for testing the aerodynamic configuration of a repaired wind turbine blade, a mould manufacture method and a repair condition testing method through the mould. The mould manufacturing method mainly comprises the steps that a position, corresponding to a damaged part of the blade, of an undamaged blade or an original complete blade mould is found out and used as a mould manufacturing area, the size of the mould manufacturing area is determined according to the size of the blade repair area, and the testing mould is manufactured based on the mould manufacturing area. Due to the fact that the testing mould meets the requirement for the standard appearance, existing before damage, of the damaged part and is small in size and light in weight, the testing mould can be conveniently carried to the blade repair area to test the aerodynamic configuration of the repaired part of the repaired blade, repair personnel are guided to make an accurate judgment and adjust repair work, the aerodynamic configuration of the repaired blade can meet the design requirement, the requirement for the aerodynamic performance of the blade is met to the maximum extent, the work efficiency of direct blade repair at a high altitude is improved, and repair cost is reduced.

The invention discloses a rear fuselage protection system, which relates to the technical field of airplane tail protection. The rear fuselage protection system comprises a mechanical part and a control part, wherein the mechanical part comprises an actuator, a rocking arm, a linkage mechanism, a sliding cover and a rail; a piston rod of the actuator is connected with the rocking arm, and the other end, which is opposite to the piston rod, of the actuator is fixed to a fuselage; the rocking arm is connected to the fuselage by virtue of a rotating shaft and can rotate around the rotating shaft, and a roller is arranged on the other end, which is opposite to the rotating shaft, of the rocking arm; one end of the linkage mechanism is fixed to the rocking arm, and the other end of the linkage mechanism is fixed to the sliding cover; the control part is used for controlling the rocking arm to retract and extend. The rear fuselage protection system has the beneficial effects that in normal flight of an airplane, the mechanical part is retracted in the fuselage, when the airplane takes off or lands, if a signal of the control part is received by the mechanical part, the rocking arm extends to protect the rear fuselage, and when the airplane takes off, the rocking arm is retracted, so that the airplane is in a good aerodynamic shape, and the aerodynamic performance is improved.

The invention provides a grid rudder and an aircraft. The grid rudder comprises a plurality of partition plates which are connected with each other and form a grid structure; wherein each partition plate is oppositely provided with a windward side and a leeward side along the two ends of the intersection line of the partition plates, at least one part of the windward side is provided with a notch,and the absolute value of the tangent slope of the outer surface where the notch is located is continuously reduced along the direction from the windward side to the leeward side, a flow guide curvedsurface is formed at the windward side, square arrangement at the windward side is avoided, and under hypersonic speed environment, on one hand, differential pressure resistance and friction resistance are reduced, power efficiency is improved, the total heat generated by heating of the windward side is reduced, and the structural strength of materials of the grid structure is guaranteed; on theother hand, the condition of local high heat generated by protruding parts such as edges and corners under the condition of airflow friction is improved, the aerodynamic heat of the grid structure isreduced, the thermal environment distribution of the grid structure is improved, and the temperature distribution uniformity is improved, so that the structural thermal stress is reduced, and the strength performance is optimized.

The invention discloses an electric folding wing system with a high aspect ratio, which comprises a pulley support, a servo motor, a synchronous belt, pulley axle systems, universal couplings, clutchaxle systems, elastic wing axle systems, axial compression springs, elastic wing rotating shaft mounting seats, elastic wings, locking pin assemblies and a cabin body. At the initial state, the wingsare in the folded state. After an unfolding command is issued, the unfolding function of four groups of elastic wings is realized under the driving of the servo motor, and the unfolding state of the elastic wings is locked by the locking pin assembly. The electric folding wing system with the high aspect ratio overcomes shortcomings of the prior art, parameterization of the design process is realized, the system can be transplanted quickly according to different weapon system requirements. The system has the advantages of simple mechanism, reliable function, small envelope size in folded state, good aerodynamic shape in unfolded state, less space occupation by the mechanism inside the cabin, smoothness and no impact in working process, and can be used as a large aspect ratio folding wing assembly of an airborne, ground-based and water-surface barrel-launched missile.

The invention relates to a method for molding aeronautic sheet metal parts, belongs to the technical field of aeronautic manufacturing and provides a method for segmented stretch forming molding of 2000-series aluminum alloy skin. The method for segmented stretch forming molding of the 2000-series aluminum alloy skin comprises the following steps that firstly, according to the radian characteristics of the skin stretch forming die profile, the skin stretch forming die profile is divided into multiple areas of a first area, a second area and the like; secondly, two ends of skin are clamped by jaws of a stretch forming machine; and thirdly, stretch forming molding is conducted on the first area, the second area and the like of the skin in sequence according to working procedures so that profiles, located in the first area, the second area and the like, of the skin and a skin stretch forming die are fitted completely, and namely, stretch forming molding of the skin is completed. The method for segmented stretch forming molding of the 2000-series aluminum alloy skin has the following advantages that firstly, the stretch forming molding uniformity is improved, and defects such as slip lines and cracks are prevented from being generated in partial areas of the skin; and secondly, the shape radian conformity after skin stretch forming is improved so that project drawing requirements are met by the shape radian conformity, and then the skin quality is improved so that more outstanding aircraft pneumatic shape can be achieved.

The invention relates to development of high-speed wind tunnel test models, in particular to a whisker enhanced photosensitive resin aero-elastic high-speed wind tunnel test model. The technical scheme is that the whisker enhanced photosensitive resin aero-elastic high-speed wind tunnel test model comprises a framework and a skin externally wrapped on the framework, wherein the skin is made of photosensitive resin filled with whiskers, and the whisker content occupies 6% of mass of the photosensitive resin. The framework comprises a beam frame and a girder, the skin is externally wrapped on beam frame and the girder which are made of photosensitive resin filled with whiskers, and the whisker content occupies 6% of mass of the photosensitive resin. The whisker enhanced photosensitive resin aero-elastic high-speed wind tunnel test model is characterized in that the framework comprises the beam frame and the girder which are made of high strength metal materials, the whiskers are filled into the photosensitive resin, and the skin or the beam frame and the girder are produced in a photo-curing forming process. The test model is similar to an aircraft in shape, rigidity distribution and mass distribution and is identical to the aircraft in the vibration mode, and the test model is simple, fast and convenient to produce and low in cost.

The invention relates to an airborne wingtip smoke puller used for flight performances. The front end of the chimney made of light material is connected with a conical fluid fairing. Inside the chimney, there are three triple-ignition smoke bomb ballasts arranged in the shape of "pin" or "one" in parallel. In the hollow structure of the chimney The peripheral surface is fixedly connected with bolts to connect the suspension beam of the wing tip and two front shoulder shafts 5 and rear shoulder shafts 7 that are fixedly connected on the suspension beams. The frustum of a circle with the surface in contact with the airfoil. The invention adopts the "product" character array cabin light-weight smoke tube and the conical fluid fairing connected to it, which is composed of three times of ignition and smoke in parallel. It has a good aerodynamic shape, the smallest outline size, small cross-sectional area, small volume and weight. Light advantage. It can meet the comprehensive factors of overall aerodynamics and strength limit. The contact between the top surface of the beam body and the airfoil is avoided by the two round tables assembled on the front and rear shoulder shafts on the upper part of the beam body. It provides an easy way for the show machine to restore to a fighter.

The invention discloses a fusion structure of a missile wing connection lug channel and a missile body. The fusion structure is formed by direct numerical control integration processing of the missile body and the lug channel. The missile wing connection lug channel and the missile body are integrally designed and fused, an appropriate aluminum sectional material is selected, the missile body and lug channel are directly finished by numerical control processing, an assembly link is not required, a connection tool is omitted, working efficiency is improved, and parallel operation is facilitated.

The invention relates to the field of construction structure, in particular to a curtain wall supporting structure optimizing aerodynamic configuration of a high-rise building. The curtain wall supporting structure is arranged on the periphery of the building body, and the building body comprises a top structure reinforcing layer and a bottom structure reinforcing layer. The curtain wall supporting structure is characterized by comprising ring beams, radial support rods, hanging rods and anti-torsion parts, wherein the ring beams are vertically arranged in multiple layers, each layer of ring beams surrounds the periphery of the building body, two ends of the radial support rods are connected with the building body and the ring beams respectively, several radial support rods are arranged on each layer of ring beams, the hanging rods are connected with the ring beams, the tops of the hanging rods are connected with the top structure reinforcing layer, and the anti-torsion parts are arranged between the building body and the ring beams. The curtain wall supporting structure is firm and reliable, regular simplicity of the building body and excellent aerodynamic configuration and appearance of the curtain wall can coexist, and manufacturing cost of engineering is reduced. Meanwhile, heat insulation and preservation can be realized by setting the internal curtain wall, and building operating efficiency is improved.

The invention relates to a segmented airborne radome which comprises a front cowling, an antenna working cover, a back cowling, a front separating frame and a back separating frame. The front separating frame and the back separating frame supply rigid supporting to the front cowling, the antenna working cover and the back cowling. The segmented airborne radome satisfies mechanical performance requirements such as rigidity and load of an airborne radome. The flange of the front cowling and the flange of the back cowling realize direct combination with an airplane body structure, thereby realizing simple structure and preventing structure weight increase caused by combination of other structures with the airplane body. The back cowling has a mode that a streamline type composite material rectifying sheath is combined with a metal-reinforced section bar. Not only is a good pneumatic shape ensured, but also enough strength and rigidity of the back cowling are ensured. The front separatingframe and the back separating frame utilize a mode that the metal section bar is combined with a composite material web plate, thereby realizing sufficient bearing capability and effectively reducingstructure weight.

The invention relates to an aviation sheet metal part and a forming method thereof, belongs to the technical field of aeronautical manufacturing, and discloses a large aircraft skin blank structure and a skin pull stretching method. According to the large aircraft skin blank structure, the end of a skin blank is provided with a structure protruding outwards in the stretching direction. A skin is of the above skin blank structure. The skin stretching forming method comprises the following steps that firstly, a jaw of a stretching machine is lifted up by a certain angle and is bent, and the twoends of the skin blank are clamped in the jaw of the stretching machine; secondly, the jaw of the stretching machine oppositely moves in the horizontal direction, and the skin blank is warped; thirdly, the height of the jaw of the stretching machine is reduced, and the skin blank is in contact with the face of a skin stretching mold; and fourthly, the jaw of the stretching machine does stretchingmotion in the direction away from the skin stretching mold in the horizontal direction, and the skin blank and the face of the skin stretching mold are completed stretched and attached. The large aircraft skin blank structure and the skin pull stretching method have the following advantages that the skin quality is improved, and therefore the more excellent aircraft pneumatic appearance is achieved.

The present invention is a large-span, foldable and reusable rocket landing buffer structure, comprising: a main leg (1), two auxiliary legs (2) and a landing plate (3); an upper joint at one end of the main leg (1) The upper joint at one end of each auxiliary leg (2) is connected to the mounting seat of the rocket body through the joint bearing provided on each side, and joint bearings are placed on the upper joint and the lower joint of the main leg (1) and the auxiliary leg (2) One side of the landing pad (3) is provided with a rotating shaft; the lower joint at the other end of the main leg (1) and the lower joint at the other end of each auxiliary leg (2) are connected to the landing pad (3) ) on the same reel. The other side of the landing plate (3) is provided with a rubber buffer pad, which can be in contact with the ground buffer when the rocket lands slowly.