258results 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.

The invention discloses a thermal photovoltaic Brayton cycle power generation system for an aircraft and a control method. The aircraft comprises a combustion chamber and a cooling channel, the system comprises a fuel assembly, the fuel assembly comprises a fuel storage tank and a first diverter valve in pipeline connection with the fuel storage tank, an outlet pipeline of the first diverter valve is connected with a cooling pipeline, and an outlet pipeline of the cooling pipeline is connected to the combustion chamber; the thermophotovoltaic power generation assembly is arranged outside the cooling channel; the cooling pipeline is arranged in the cooling channel, and fuel in the cooling pipeline is used for cooling the cooling channel and the photovoltaic power generation assembly; and the Brayton cycle power generation assembly comprises a cooler, and the cooling pipeline is connected with a first inlet pipeline of the cooler. On the basis of Brayton cycle power generation, a thermal photovoltaic power generation system is coupled, so that the generating capacity can be increased while cooling of a cooling channel is ensured, and the thermoelectric conversion efficiency of the system is improved.

An airfoil for an aircraft having an airfoil geometry. The airfoil geometry includes an upper surface and a lower surface, the upper and lower surfaces are in accordance with Cartesian coordinate values of X and Y. The Cartesian coordinate values of X and Y are non-dimensional values from 0 to 1 convertible to dimensional values by multiplying the Cartesian coordinate values of X and Y by a chord length of the airfoil. The upper surface and the lower surface form an enclosed airfoil shape. The aircraft is a supersonic aircraft, and the airfoil has a thickness-to-chord ratio of about 10%.

The invention provides a wing-body fusion wide-speed-range aircraft layout design method based on consideration of an afterbody effect. The method comprises the following steps: firstly, generating a cone-guided waverider precursor of a wide-speed-range aircraft for supersonic flight; then, based on the basic airfoil profile, performing pneumatic optimization on the basic wing profile and the strake wing profile of the wide-speed-domain aircraft to obtain a multi-section wide-speed-domain waverider airfoil profile; then calculating the length of a waverider afterbody and designing a symmetry plane of the wide-speed-range aircraft; and finally, based on an osculating cone method and transonic speed strake wing constraints, obtaining three-dimensional layout parameters by using the waverider precursor parameters and the waverider afterbody length, and assembling the basic wing optimized airfoil profile, the strake wing optimized airfoil profile and the symmetry plane to generate the layout of the wide-speed-domain aircraft. According to the method, the cone-guided waverider generation method, the wide-speed-range airfoil optimization method and the osculating cone double-sweepback waverider generation method are coupled, and the aerodynamic design of the afterbody is considered, so that the aerodynamic performance of the wide-speed-range aircraft under the subsonic / transonic / supersonic speed is improved.

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 discloses a lattice composite structure thermal protection air rudder for a high-speed aircraft and a manufacturing method, and the manufacturing method comprises the steps: 1, according to the thermal environment conditions and load conditions of the aircraft and the response characteristic requirements of the air rudder, a rudder shaft and a control surface of a metal framework are determined, the rudder shaft adopts a solid structure, and the control surface adopts a lattice structure; 2, according to the thermal environment condition of the aircraft, in combination with the temperature resistance of the metal framework, a composite material heat-proof structure is determined, and the control surface is coated with the composite material heat-proof structure, and the lattice structure is completely filled with the composite material heat-proof structure; 3, a metal framework is manufactured through a selective laser melting additive manufacturing process; and 4, the composite material heat-proof structure is manufactured through a resin transfer molding forming technology or an integrated mold pressing technology. The problem that a traditional air rudder is limited in the aspects of weight, structural efficiency, heat conduction management and the like is solved.

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 provides a hypersonic velocity leading edge combined cooling structure, system and method capable of coping with space-time non-uniform heat flow, and is mainly used for solving the problems that in the existing active thermal protection technology, head cooling insufficiency and downstream cooling excess are likely to occur in space, the cooling effect is difficult to dynamically adjust in time, cooling insufficiency or excess in some time periods is caused, and the cooling effect is poor. And the whole porous medium structure is inconvenient to check and maintain. The invention discloses a hypersonic velocity leading edge combined cooling structure capable of coping with space-time non-uniform heat flow. The hypersonic velocity leading edge combined cooling structure comprises a diverging part, a structure part, a rear cover plate and a shunting core, through discrete arrangement of the diverging parts and setting of different porosities for the diverging parts at different positions, the porosities of the N diverging parts are sequentially reduced from the head part to the tail part, and the larger the porosities are, the more the coolant distribution amount is, so that the technical problems of insufficient head cooling and excessive downstream cooling easily occurring in space are solved.

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.

An aircraft wing including a wing structure supporting a wing cover having an outer aerodynamic surface, and a pop-up spoiler unit including an actuation mechanism and a spoiler coupled to the actuation mechanism. The spoiler is moveable by the actuation mechanism to deploy through an aperture in the wing cover. The pop-up spoiler unit is mounted to the wing structure to be substantially isolated from wing bending deflections. The wing structure has a neutral plane which is substantially neither stretched nor compressed during bending of the wing structure. The pop-up spoiler unit is mounted to the wing structure at only two pin joints each having a respective pin axis about which the pop-up spoiler unit is substantially free to rotate, and each pin axis substantially lies in the neutral plane.

A graded multilayered composite comprises a metal matrix material having a first side and a second side opposite the first side. A first layer of microspheres is dispersed on the first side of the metal matrix material. A second layer of microspheres is dispersed on the second side of the metal matrix material.

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.

The invention relates to an improved deflection flap for a vehicle with a fuselage featuring an airfoil profile. The deflection flap at the bow is two-part. It includes a device for raising and lowering it.

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 invention discloses a Ca3Co4O9-based metal ceramic material with a nano / micro honeycomb sandwich structure and a preparation method of the Ca3Co4O9-based metal ceramic material, and belongs to the technical field of high-temperature ceramic-based thermoelectric conversion materials.The preparation method comprises the steps that a heated and molten alloy is infiltrated into a surface-modified porous Ca3Co4O9 ceramic matrix through a normal-pressure infiltration or double-face-vacuum melting infiltration technology, and the Ca3Co4O9-based metal ceramic material with the nano / micro honeycomb sandwich structure is obtained; and the Ca3Co4O9 metal ceramic material is successfully prepared. According to the method, after the surface of ceramic is modified, a Ca3Co4O9 ceramic matrix is vertically placed on an alloy under the normal pressure or vacuum condition, after the alloy is heated and melted, an alloy liquid phase spontaneously fills pores in the ceramic or permeates into the ceramic due to the action of capillary force, and after the alloy is solidified, a two-phase combined honeycomb sandwich structure is formed. The conductivity and seebeck coefficient of the material can be effectively improved, so that the ZT value of the material is improved.

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.

The invention relates to an improved hinged deflection panel for a vehicle comprising a fuselage having a wing profile. The hinged deflection panel at the front is a two-part piece. Also provided is a device for lifting and lowering.

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.