181results about "Supersonic aircraft" patented technology

The invention provides an ultrahigh-temperature ceramic precursor, a preparation method thereof, a modified high-performance inorganic fiber and application, and relates to the technical field of high-temperature-resistant materials. The preparation of the ultra-high-temperature ceramic precursor comprises the following steps: placing a transition metal compound mixed solution in a low-temperature environment, standing, adding a coordination agent and an exchange agent, reacting, filtering in a low-temperature nitrogen atmosphere, adding a mixed solvent into the filtrate, heating and refluxing, and carrying out a polymerization reaction to obtain the ultra-high-temperature ceramic precursor. The superhigh-temperature ceramic precursor provided by the invention forms a high-temperature protective coating on the surface of a high-performance inorganic fiber and other materials, and meanwhile, a transition metal ion solution is placed under a low-temperature condition through a metal ion stabilization method, so that the internal structure of transition metal ions is more stable, and uniform dispersion at a molecular level is achieved; and the uniformity of the elements can be kept in the curing and cracking processes, and the phenomenon of coating stripping caused by mismatching of thermal shrinkage coefficients of a coating and a substrate material in an extreme environment is avoided.

Combined aerospace system is a multi-stage technology to deliver payloads to orbit and to midair. The system comprises a carrier aircraft, a tow aircraft, a towed spaceplane or a rocket plane or a glider. The carrier aircraft, coupled with the spaceplane and / or the rocket plane or several gliders, takeoff from a runway and ascends to a staging altitude and speed under the carrier aircraft power. After staging, the tow aircraft will pick up the spaceplane or the rocket plane or the glider from the carrier aircraft in flight. The towed spaceplane or the rocket plane or the glider continues ascending and accelerating by the tow aircraft. The spaceplane or the rocket plane will disconnect from the tow aircraft and continue to ascend and accelerate in an atmosphere by using the ramjet engine. The spaceplane or the rocket plane equipped with a propulsion unit is used to place the spacecraft into the earth orbit. The glider will disconnect from the tow aircraft and lands in designated place. Gliders can be used for instantly delivering of correspondence, parcels, and civil goods from a supplier to wide variety of individual consumers or corporate customers.

A thermal gradient management assembly comprising a functional component. The functional component includes an exterior facing segment, where the exterior facing segment is configured to experience aerodynamic heating, and an interior segment with the exterior facing segment, where the interior segment is configured for isolation from the aerodynamic heating. The functional component includes a thermal gradient extending between the exterior facing segment and the interior segment. The thermal gradient management assembly also comprises a gradient mitigating heating system includes a power source and a gradient heating element in communication with the power source, wherein the gradient heating element is coupled with the interior segment. The gradient heating element is configured to diminish the thermal gradient between the exterior facing and the interior segments.

A thermal protection structure of a high-temperature and supersonic aircraft belongs to the technical field of thermal protection of aircrafts under high heat flux and comprises a water absorption layer, a front edge area and two side areas, and the front edge area comprises a first evaporative cooling porous medium layer, a cross-scale porous medium layer and a high-temperature-resistant wall surface which are sequentially laid on the outer side of the front edge of the water absorption layer; the two side areas comprise a second evaporative cooling porous medium layer and a third evaporative cooling porous medium layer which are laid on the outer sides of the two sides of the water absorption layer respectively, and a coolant conveying pipeline communicated with the front edge and the two side areas is arranged on the inner side of the water absorption layer; fluid is sucked to the high-temperature-resistant wall face through the capillary acting force of micron-scale and nano-scale pores of the first evaporative cooling porous medium layer and the cross-scale porous medium layer, the millimeter-scale pores of the cross-scale porous medium layer provide steam escape channels, and the high-temperature-resistant wall face blocks direct disturbance of high-temperature main flow to the fluid in the porous medium. Steam outlets on two sides of a front edge stagnation point discharge steam; the device has the advantages of no need of an additional liquid supply power device, high temperature resistance and high cooling efficiency.

The invention provides a wave rider capable of vertically ascending and descending, and belongs to the technical field of wide-speed-range aircrafts, the wave rider specifically comprises a front edge shock wave generating surface, a central fuselage and sweepback wings located on the two sides of the central fuselage, the two sweepback wings are provided with mounting holes penetrating through the upper surface and the lower surface of the wave rider, and lift fans are mounted in the mounting holes; the lift fans are embedded into the mounting holes and used for providing vertical lift force for the wave rider, the axes of the lift fans are parallel to the vertical axis of the wave rider, and the lift fans on the two sweepback wings are symmetrically distributed on the two sides of the central fuselage. According to the treatment scheme, the vertical lifting capacity is provided for the wave rider, and the high-speed flight aerodynamic performance is considered.

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.

The invention relates to the technical field of layered products, and particularly discloses a heat-insulating and wave-absorbing skin material as well as a preparation method and application thereof. The heat-insulating and wave-absorbing skin material is of an integrated layered structure based on an aluminum oxide-mullite system, comprising a functional gradient layer with double gradients of fiber density and aperture, an impedance matching layer with a gradient dielectric constant, a resonance wave absorbing layer with a periodic dot matrix formed by silicon carbide and aluminum oxide core-shell microspheres, and a heat insulation layer with a nano-micron double-scale pore structure. According to the preparation method, a green body is constructed mainly through tape casting and silk-screen printing, and then co-sintering integrated forming is carried out. The material has the excellent characteristics of high temperature resistance, toughness, efficient heat insulation, broadband wave absorption and the like, and is suitable for extreme environments such as hypersonic aircraft skin and the like.

The invention relates to the technical field of aerospace, and particularly provides a liquid metal closed-loop thermal management system and method for a hypersonic aircraft. The system comprises a liquid metal circulation loop which comprises a liquid metal pump used for driving a liquid metal working medium to circularly flow; the heat absorption device is integrated in the liquid metal circulation loop and used for receiving operation heat of the hypersonic flight vehicle and transmitting the operation heat to the liquid metal working medium flowing through the hypersonic flight vehicle. The thermo-acoustic power generation device at least comprises a hot end heat exchanger connected to the heat absorption downstream of the liquid metal circulation loop in series and used for converting part of heat energy of the liquid metal working medium heated after heat absorption into electric energy so as to drive a liquid metal pump. The liquid metal circulation loop further comprises a first diverter valve and a second diverter valve, the first diverter valve is connected into the heat absorption device, and the second diverter valve is connected into the thermo-acoustic power generation device. The dynamic cooperative controller is used for controlling the dynamic cooperative controller based on the flight state parameters of the hypersonic aircraft, the engine thrust requirement and the temperature of the key thermal protection part. The total flow of the liquid metal working medium and flow distribution in the heat absorption device and the thermo-acoustic power generation device are dynamically and cooperatively adjusted by controlling the liquid metal pump, the first diverter valve and the second diverter valve. According to the invention, the compactness of the system can be obviously improved, the heat sink consumption is greatly reduced, and a reliable and efficient heat management solution is provided for long-endurance hypersonic flight.

The invention discloses a testing device of an active and passive combined thermal protection system, and relates to the technical field of thermal protection of hypersonic aircrafts. The micro-channel cold plate comprises a micro-channel cold plate assembly, the top of the micro-channel cold plate assembly is fixedly connected with a passive thermal insulation layer, and a plurality of thermocouples are installed between the passive thermal insulation layer and the micro-channel cold plate assembly; a gear pump is installed at an inlet of the micro-channel cold plate assembly through a fuel oil pipeline, a fuel oil tank is installed at the inlet end of the gear pump, and an adjusting valve is installed at the outlet end of the gear pump. By adjusting the flow of the micro-channel cold plate in real time, the change of an external pneumatic thermal environment is well matched, so that efficient thermal protection of a large-area fuselage of a high-speed aircraft under long-time and high-heat-flow conditions is realized, and the consumption of a coolant is effectively reduced; and when the control module adjusts the flow, constant-pressure supply of the cooling working medium is achieved by changing the opening degree of the pressure adjusting valve, the working stability is improved, and the thermal failure problem of the active and passive thermal protection system is effectively avoided.

The invention relates to the technical field of thermal protection of high-speed aircrafts, in particular to a pump-driven two-phase fluid loop for high-temperature heat dissipation in a fixed plane area and a design method. The pump-driven two-phase fluid loop for high-temperature heat dissipation in the fixed plane area comprises a pipeline system, a condenser, a liquid storage container and a power pump. The pipeline system comprises a main loop and a plurality of heat dissipation branches; the heat dissipation branches are arranged in the corresponding gaps; and the two ends of the main loop are communicated with the plurality of heat dissipation branches through multi-stage branches. Under the action of the power pump, the cooling working medium in the liquid storage container is pumped into each heat dissipation branch of the multi-stage branches; in the heat dissipation branch, the cooling working medium completes heat exchange, at least part of the liquid-phase cooling working medium is converted into a gas phase in the process, then the gas-phase and liquid-phase mixed cooling working medium returns to the main loop and enters the condenser for heat exchange and cooling, finally, the cooling working medium is completely converted into the liquid-phase cooling working medium, and the cooling working medium can be evenly distributed in a target area. The cooling effect is improved.

The invention relates to a hypersonic aircraft leading edge active and passive composite thermal protection structure and a preparation method thereof. The thermal protection structure comprises a high-temperature metal heat pipe, a parting flow channel layer plate and a porous sweating layer plate which are arranged in a stacked mode from inside to outside, the parting flow channel layer plate is provided with a branch-shaped parting network flow channel, and the porous sweating layer plate is provided with concave grooves which are distributed in a staggered mode in the shape of the Chinese character'yao '. And the bottom of the sweating groove is communicated with the dendritic parting network flow channel through the sweating holes. Passive heat dissipation of the high-temperature metal heat pipe and active cooling of the parting flow channel-porous sweating layer are used for cooperative work, and advantage complementation of passive heat protection and active heat protection is achieved; the parting runner-porous sweating layer design can realize uniform distribution of a coolant on the porous sweating layer plate, and a large number of micro-pore channels formed by the porous sweating layer micro pin fins-interconnected embedded grooves significantly promote sweating cooling, so that efficient active and passive composite thermal protection is realized.

A shielded, multi-layer heatshield material for hypersonic flight applications includes an ablative / insulative structure having a thickness and providing both thermal insulation and a first ablation resistance, and a separate shield layer bound to an outer surface of the ablative / insulative structure, the shield layer being thinner than the ablative / insulative structure while providing a second higher ablation resistance greater than the first ablation resistance. In one arrangement the ablative / insulative structure includes a carbon / carbon composite layer attached to a syntactic carbon foam layer, using carbon fiber stitched loops for reinforcement, and the shield layer is a carbide outer layer. In another arrangement the ablative / insulative structure includes a silica composite layer attached to a silica insulative foam layer, using stitched loops for reinforcement, and the shield layer is a ceramic outer layer.

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.

The invention relates to an engineering-oriented wave rider airfoil efficient design method. The method comprises a traditional wave rider theoretical aerodynamic shape cutting method and a wave rider airfoil shape and position layout design method for a wave rider wing, a vertical tail, a pelvic fin and other airfoils. Aiming at the problems that the traditional wave rider theoretical appearance is only suitable for a single design point and the stability and the controllability are insufficient during cross-speed-domain flight, a system solution is provided for cross-speed-domain stable flight of the hypersonic flight vehicle. According to the method, the design technology breakthrough is realized by combining the parameterized cutting technology of the traditional wave rider theoretical appearance and the multi-airfoil collaborative optimization design. Through a global parameter joint debugging algorithm, after airfoil design is completed, the aerodynamic center position variation of the waverider under different Mach numbers is reduced, trimming is easy, and engineering adaptability design of a theoretical shape is achieved.

To provide multilayered material systems tuned to locally meet stringent thermomechanical loading requirements on an extreme environment hypersonic airframe structure.SOLUTION: A graded multilayered composite (1100) comprises a metal matrix material (1110) having a first side (1112) and a second side (1114) opposite the first side (1112). A first layer (1120) of microspheres is dispersed on the first side (1112) of the metal matrix material (1110). A second layer (1122) of microspheres is dispersed on the second side (1114) of the metal matrix material (1110).SELECTED DRAWING: Figure 1

The present invention discloses a deformation mechanism for the front middle section of a supersonic aircraft cabin. The driving mechanism includes a curved beam, which is the main load-bearing structure of the driving mechanism and is fixed to the front and rear ends of the front middle section of the aircraft cabin; four fan-shaped bottom plates of the same size are equally fixedly connected to the curved beam; a linear guide rail is connected to and fully fits the bottom of the curved beam, and a linear motor group is provided on the linear guide rail, which moves along the linear guide rail; the linear motor group is connected to a slider group through a push rod; a guide rail is provided on the fan-shaped bottom plate for the slider group to move up and down; a synchronous connecting rod passes through the holes on the slider group and the perforated plate to realize the synchronous movement of the slider group; the driving mechanism of the present invention is a synchronous motion device of three groups of slider mechanisms driven by a single motor module, which realizes the problem that the front middle section of a single aircraft cabin has a variable lift-to-drag ratio to meet different flight missions.

The invention provides a large-scale supersonic civil aircraft wing body fusion low-sonic-boom aerodynamic layout, and relates to the field of supersonic civil aircraft design, the large-scale supersonic civil aircraft wing body fusion low-sonic-boom aerodynamic layout comprises a nose, a middle fuselage, a wing body transition section, wings and a V-shaped empennage; the machine head is in a duckbilled flat shape, the upper contour line is a downward-inclined curve, the lower contour line is an upward-warping curve, the side contour line is a sharp-arch-shaped curve, the section of the head is in the shape of a combination of a triangle and an arc, the upper portion of the section of the head is a triangle with the top in arc transition, and the lower portion of the section of the head is an arc. The camber of the center line of the middle fuselage is small, and only the front section bends down slightly; the middle fuselage is connected with the roots of the wings through wing body transition sections; the wings are arranged on the two sides of the middle fuselage and adopt an arrow-shaped wing layout with a small aspect ratio, a variable dihedral angle and a variable torsion angle and a camber wing type; the V-shaped empennages are arranged on two sides of the tail end of the According to the method, the problem that a gap still exists in the fusion low-sonic-boom aerodynamic layout of the wing body of the large supersonic speed civilian aircraft is solved, and reference is provided for research on the fusion layout of the supersonic speed wing body.

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.

A supersonic aircraft is disclosed herein. The supersonic aircraft comprises one or more fuselages, one or more wings, at least one wing-mounted engine, at least one inlet mounted on the upper surface of or above the wing, an upper wing surface propulsion system fairing, and a lower wing surface propulsion system fairing or lower wing surface modification to account for the presence of the engine.

The three-inlet integrated design method of double-waverider forebody is considered, which is related to the integrated design of hypersonic vehicle forebody / inlet. According to the design conditions, the double-waverider forebody and one middle internal contraction inlet and two symmetrical external internal contraction inlets are given. The middle internal contraction inlet is designed by tracing the flow lines of three-dimensional internal contraction axisymmetric reference flow field, and the external internal contraction inlets are also designed by tracing the flow lines of three-dimensional internal contraction axisymmetric reference flow field. The internal contraction reference flow field of the internal waverider forebody and the external contraction flow field of the external waverider forebody are aerodynamically coupled, so that the forebody and the lower surface of the middle internal contraction inlet maintain waverider characteristics, maintain the overall aerodynamic performance, and arrange two symmetrical external internal contraction inlets to improve the captured flow rate, further compress the incoming flow, and improve the aerodynamic performance and stability of the body; both efficiency and performance requirements are considered.

The application belongs to the field of hypersonic active flow control, and discloses a drag reduction and heat protection system and method for a hypersonic aircraft. The method combines energy deposition and magnetic fluid flow control. Energy deposition has excellent drag reduction characteristics, which can bring good drag reduction effect to the aircraft. Meanwhile, energy deposition can be used as auxiliary ionization technology to improve the magnetic control effect of magnetic fluid flow control by improving the electrical conductivity of the fluid. Magnetic fluid flow control has outstanding heat protection performance, which can bring excellent heat protection effect to the aircraft after dealing with the heat flow growth caused by energy deposition. Meanwhile, the Lorentz force of the magnetic field fully suppresses the resistance impact and shock caused by pulse energy deposition. The method can achieve better hypersonic drag reduction and heat protection effect, effectively reduce the resistance of the hypersonic aircraft and improve the heat protection ability of the aircraft, and has wide application prospect.

The application discloses a variable sweep in-out telescopic wing aircraft variable structure, which comprises an outer wing, a fuselage, a screw system, a push rod, an inner wing, a push rod connecting piece, a graphite copper sleeve and upper and lower guide rails; wing storage grooves are formed on the two sides of the fuselage respectively and used for accommodating the outer wing and the inner wing; the inner wing is located inside the outer wing and connected with the outer wing through a rotating pair to realize rotating movement; the inner and outer wings are designed with special trajectory sliding grooves, one end of the push rod is connected with the sliding grooves, the other end is fixed with a sliding table of a ball screw through a connecting piece, a motor drives the screw to rotate, realizes the symmetrical movement of the push rod, controls the movement of the inner and outer wings and realizes the deformation of the wings. Through the design of the sliding grooves of the inner and outer wings, the push rod is connected with the inner and outer wings through the sliding grooves and plays a certain supporting role; the length of the sliding grooves is increased, the change range of the sweep angle is increased, the uniform movement of the screw is realized through the movement of the push rod along the sliding grooves, the outer wing is uniformly extended and the inner wing is uniformly rotated, the overall structure is simple and ingenious and is flexible and convenient to use.

This invention relates to the field of layered product technology, specifically disclosing a heat-insulating and microwave-absorbing skin material, its preparation method, and its application. The heat-insulating and microwave-absorbing skin material is an integrated layered structure based on an alumina-mullite system, comprising a functionally gradient layer with dual gradients in fiber density and pore size, an impedance matching layer with a gradient in dielectric constant, a resonant microwave-absorbing layer composed of a periodic lattice of silicon carbide@alumina core-shell microspheres, and a heat-insulating layer with a nano- to micron dual-scale pore structure. The preparation method mainly involves constructing a green body through casting and screen printing, followed by co-sintering to form an integrated structure. The material of this invention possesses excellent properties such as high temperature resistance, toughness, high-efficiency heat insulation, and broadband microwave absorption, making it suitable for extreme environments such as hypersonic vehicle skins.

The application discloses a kind of inner and outer flow integrated high-speed full-flow surface waverider aircraft design method, it is related to aerodynamic shape design field, comprising: S1, based on the constraint condition of incoming flow condition and axial symmetry curved surface cone, utilize characteristic line method to obtain two mutually matched first axial symmetry curved surface cone flow field, second axial symmetry curved surface cone flow field;S2, in two axial symmetry curved surface cone flow field, respectively construct upper waverider surface, lower waverider surface, precursor inlet of waverider aerodynamic configuration using streamer tracking method;S3, upper waverider surface, lower waverider surface and precursor inlet are assembled according to flow field characteristics, and the appearance structure of inner and outer flow integrated high-speed full-flow surface waverider aircraft is obtained.The application realizes the design of high-speed air-breathing waverider configuration from the level of theoretical method, all components are designed using flow surface and the flow field between each component is matched, which provides a very effective and efficient method and technical support for designing waverider configuration meeting engineering practical requirements.

The invention discloses an aramid fiber structure oil tank of a supersonic unmanned aerial vehicle and a forming process method of the aramid fiber structure oil tank, and belongs to the technical field of advanced fiber reinforced resin matrix composite manufacturing. In order to solve the problems that an existing metal oil tank is heavy, an existing carbon fiber oil tank is poor in impact resistance, and in the forming process, mold precision is low, curing defects are many, splicing sealing is weak, and detection is single, aramid fiber prepregs are symmetrically laid by + 45 degrees / 0 degrees / 90 degrees / -45 degrees / -45 degrees / 90 degrees / 0 degrees / + 45 degrees; the invention discloses a Q235 female die block die. Synchronously heating, pressurizing and curing the autoclave; splicing is carried out by using Eongda 420 / XM22 glue, and epoxy resin is coated inside; and the detection is combined with an 80Kpa air tightness test and Hyperworks stress simulation. The fuel tank is light in weight, high in strength, resistant to impact and high in airtightness, adapts to supersonic extreme environments, improves the voyage and safety of an unmanned aerial vehicle and has remarkable engineering application value.

The invention is notably directed to an aircraft, which includes an aircraft structure, a thermal exchange system, a thermal insulation material, and fasteners. The thermal exchange system includes a fluid-based cooling circuit and heat exchanger tiles. The latter cover, at least partly, the aircraft structure. They may thus form the skin of the aircraft. The fasteners secure the tiles to the aircraft structure. Each tile of the heat exchanger tiles embeds one or more fluidic channels, which are connected to the fluid-based cooling circuit, whereby the thermal exchange system is adapted to actively cool the tiles, in operation. Remarkably, the thermal insulation material extends between the aircraft structure and the heat exchanger tiles. The thermal insulation material is clamped on either side by the heat exchanger tiles and by the aircraft structure, thanks to the fasteners.

The invention discloses a preparation method of large-size high-strength Y2O3-MgO nano-composite ceramic, and relates to the field of transparent ceramic materials, the preparation method comprises the following steps: loading Y2O3-MgO nano-composite powder with the particle size of 10-60 nm into a cold isostatic pressing mold, then loading into a sealing bag, and vacuumizing; the sealed sample piece is put into a cold isostatic press for cold isostatic pressing forming, the forming pressure is 140-200 MPa, and a low-defect-density green body is obtained; and carrying out vacuum and hot isostatic pressing sintering on the formed Y2O3-MgO nano complex-phase green body. According to the preparation method, the problems that the large-size Y2O3-MgO nano-composite ceramic is low in yield, relatively large in grain size, relatively poor in mechanical property and the like can be solved through a stepped cold isostatic pressing molding technology and a vacuum + hot continuous sintering technology, and the preparation method is a preparation scheme of the Y2O3-MgO nano-composite ceramic with practical application value.

This invention discloses a composite thermal protection aerodynamic rudder for aircraft and its manufacturing method, aiming to solve the problems of high manufacturing cost, poor thermal insulation, and poor overall mechanical performance in existing structures. The composite thermal protection aerodynamic rudder of this invention includes a metal frame, lightweight filler, and a composite material thermal protection structure arranged sequentially from the inside out. The metal frame includes a rudder shaft and a rudder surface, which are integrally formed. The rudder surface has an internal frame rib structure forming an internal cavity, which is filled with lightweight filler. The composite material thermal protection structure covers the outside of the rudder surface and the lightweight filler, forming the aerodynamic shape of the aerodynamic rudder. Through the frame rib structure design of the metal frame, the amount of material used is significantly reduced while ensuring load-bearing capacity, achieving lightweighting, which is beneficial to improving the thrust-to-weight ratio and maneuverability of the aircraft. Optimizing the rib layout according to the load path achieves efficient material utilization.

The invention relates to a wide-speed-range airfoil profile suitable for aerodynamic layout of a high-pressure capture wing, which is characterized in that the maximum thickness position of the airfoil profile is located at the position 72% times of the chord length of the airfoil profile away from the front edge of the airfoil profile, and the maximum thickness of the airfoil profile is 4% times of the chord length of the airfoil profile; the maximum camber position of the airfoil profile is located at the position of the airfoil profile chord length 74% times away from the front edge of the airfoil profile, and the maximum camber of the airfoil profile is 1.92% times of the airfoil profile chord length. The aerodynamic characteristics of subsonic velocity, transonic velocity and hypersonic velocity can be well considered.

A method of constructing a skin assembly comprises forming a first skin panel and a second skin panel; removing material from the first skin panel and the second skin panel, respectively, to form one or more first fingers and one more second fingers. The method also comprises joining the first skin panel and the second skin panel such that the first skin panel and the second skin panel define a plurality of staggered expansion gaps therebetween.