329results about "Supersonic aircraft" patented technology

The corrugated plate type integrated thermal protection system based on the porous dot matrix composite phase change material comprises an outer plate and an inner plate, the outer plate and the inner plate are arranged in a spaced mode, and a web, a thermal insulation material and the composite phase change material are arranged in a spaced cavity formed by the outer plate and the inner plate; the multiple web plates are arranged in the interval cavities in a penetrating mode and used for forming a supporting structure for connecting the outer plate and the inner plate, and the multiple web plates are inclined to one another; the spacing cavity comprises a first area and a second area, the first area is close to the outer plate and used for containing a heat insulation material, and the second area is close to the inner plate and used for containing a composite phase change material; wherein the composite phase change material is arranged in a porous dot matrix mode, and the porous dot matrix is adopted as a heat conduction reinforcing body. The high heat storage efficiency of the porous dot matrix composite phase change material is utilized to quickly absorb and store heat of the web, so that the heat transferred into the aircraft is reduced, and the system has higher heat insulation performance.

The invention discloses a supersonic aircraft head thermal protection structure, and belongs to the technical field of hot end component cooling, the supersonic aircraft head thermal protection structure comprises a cooling system, a preheating system, a CBC power generation system and a secondary cooler, the cooling system comprises a coolant liquid storage tank, an electromagnetic adjustable pump and a cooling channel arranged in the wall surface of the aircraft head, a cooling medium flowing in a direction opposite to that of external high-temperature gas is arranged in the cooling channel; the preheating system comprises a fuel tank, a fuel adjustable pump and a fuel channel; the secondary cooler is used for transferring heat in the cooling channel into the fuel channel; the CBC power generation system comprises a heater, a primary cooler and a circulating working medium, the heater is used for transferring heat in the cooling channel to the circulating working medium, and the primary cooler is used for transferring heat in the circulating working medium into the fuel channel. Efficient thermal protection of the head of the supersonic aircraft can be achieved, the risk of structural failure caused by high temperature is reduced, and the safety of the aircraft is improved.

The invention relates to a heat protection structure of a dredging coupling semi-open type dissipation structure, which belongs to the technical field of heat protection of high-speed aircrafts and comprises a dredging structure shell, a dredging structure capillary core body, a dredging structure working medium, a semi-open type dissipation structure, a semi-open type dissipation working medium, a semi-open type dissipation structure exhaust pipe and a semi-open type dissipation structure control switch. Aiming at a local high-heating part of a hypersonic aircraft, on the basis of semi-active dredging type uniform-temperature heat protection, phase change heat absorption and discharge heat dissipation of an active dissipation working medium are combined, and heat absorption and dissipation ways are increased, so that the total heat amount is effectively reduced while the internal heat distribution of the structure is optimized, and the heat dissipation efficiency is improved. Internal heat control and management of the structure are realized through coupling of dredging semi-active heat protection and active dissipation, the problem of non-ablative heat protection in an extreme environment is solved, and a feasible heat protection technical approach is provided for performance improvement of high-speed aircrafts in the future.

The invention discloses a variable sweepback aircraft variant structure with inner and outer telescopic wings. The variable sweepback aircraft variant structure comprises outer wings, a fuselage, a lead screw system, push rods, inner wings, push rod connecting pieces, graphite copper sleeves and upper and lower guide rails, wing accommodating grooves are formed in the two sides of the fuselage respectively and used for accommodating the outer wings and the inner wings; the inner wing is located in the outer wing and connected with the outer wing through a rotating pair, and rotating motion can be achieved. The inner and outer wings are provided with chutes with special tracks, one end of a push rod is connected with the chutes, the other end of the push rod is fixed with a sliding table of a ball screw through a connecting piece, and a motor drives the screw to rotate to realize symmetrical movement of the push rod, control of actions of the inner and outer wings and wing deformation. Through the design of the sliding grooves of the inner and outer wings, the push rods are connected with the inner and outer wings through the sliding grooves, a certain supporting effect is achieved, the length of the sliding grooves is increased, the change range of the sweepback angle can be increased, the constant-speed movement of the lead screws realizes constant-speed extension of the outer wings and constant-speed rotation of the inner wings through movement of the push rods along the sliding grooves, and the overall structural design is simple and ingenious and flexible and convenient to use.

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.

The invention provides a supersonic-speed jet adopting supersonic-speed double wings and variant design, and belongs to the field of novel supersonic-speed civil aircrafts, a large-area truncated delta wing with edgings on the inner side is adopted, an engine nacelle is arranged on the upper portion of the outer end of the large-area truncated delta wing, and the supersonic-speed double wings are installed on the outer side of the nacelle; the front edges and the rear edges of the upper wing and the lower wing of the supersonic double-wing adopt intelligent materials to realize a variant function; the empennage can rotate around a mounting shaft in the longitudinal direction of the airplane to change the inclination angle; the front edge and the rear edge of the supersonic double-wing are deformed, and the inclination angle of the empennage is changed to jointly form a supersonic state and a subsonic state. The problem that the conventional double-wing non-design-point performance is poor is solved, high aerodynamic performance can be achieved within the flight envelope range, and it is guaranteed that an aircraft achieves ideal supersonic speed environment-friendly performance such as acoustic explosion and noise and aerodynamic efficiency.

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.

Method and system for reducing a temperature differential across a flight vehicle. The system includes an exterior of the flight vehicle. The exterior of the flight vehicle includes a first exterior panel being at a first initial temperature and a second exterior panel being at a second initial temperature to define an initial temperature differential. A fluid loop, containing a working fluid, is coupled to the first exterior panel and the second exterior panel such that heat is transferred from the first exterior panel to the working fluid to decrease the first initial temperature of the first exterior panel to a first altered temperature, and to heat the working fluid. The heated working fluid is circulated from the first exterior panel and used to heat the second exterior panel to a second altered temperature greater than the second initial temperature.

The invention relates to a self-adaptive skin structure suitable for a cross-airspace speed domain morphing aircraft, and belongs to the technical field of intelligent skin. The skin structure is of a multi-layer form and sequentially comprises a heat insulation protection layer, a flexible skin base body, a sensing layer and a bearing and driving layer from the outer side to the inner side. Wherein the outermost side adopts a fiber sandwich structure formed by coating aerogel with an alumina fiber reinforced alumina composite material as a heat insulation protective layer; a high-temperature-resistant carbon fiber elastic matrix composite material is used as a flexible skin matrix; a bearing layer and a driving layer on the inner side are integrally designed, and a honeycomb cell element structure driven by an integrated high-temperature shape memory alloy wire is adopted; an optical fiber sensor is embedded between the flexible skin substrate and the bearing and driving layer, the optical fiber sensor realizes direct strain measurement, and the thickness change of the skin is calculated through a strain measurement value; the method can realize self-adaptive active flexible deformation of the skin when the cross-airspace speed domain aircraft is in a hypersonic flight stage, and has important application value in the field of skin design of the cross-airspace speed domain aircraft.

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 belongs to the technical field of aviation, and particularly relates to a supersonic civil aircraft layout. According to the supersonic civil aircraft layout, the engine nacelle is installed above the rear fuselage between the wing and the horizontal tail, the lower portion of the air inlet is shielded by the rear edge of the wing, the lower portion of the air outlet is shielded by the front edge of the horizontal tail, shock waves generated by the engine are prevented from spreading to the ground through the shielding effect of the wing and the horizontal tail, and the ground sound explosion loudness of the aircraft is effectively reduced; the engine and the hanger can perform small-angle pitching deflection around the rotating shaft in the hanger, so that the mounting angle of the engine is adjusted, and the propulsive efficiency of the engine under the non-design-point working condition is improved; the horizontal tail tip can be folded upwards to provide additional vertical tail capacity during supersonic cruising, loss of vertical tail efficiency during supersonic cruising is made up, and expansion waves generated by an engine exhaust nozzle and a horizontal tail are prevented from spreading to the ground. And the wingtip is put down during low-speed take-off and landing, so that a larger horizontal tail area is ensured, and the trimming requirement of a take-off and landing structure is met.

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 relates to the technical field of aircraft design, in particular to a sectional type hybrid power supersonic commercial passenger plane which comprises a supersonic main aircraft and a subsonic babysitter machine. In the take-off and landing stages, the subsonic babysitter carries the supersonic main aircraft to form a combined body, an engine of the supersonic main aircraft supplies power to a ducted fan of the subsonic babysitter and charges an energy storage device of the ducted fan, and the main aircraft and the subsonic babysitter cooperate to generate thrust to jointly fly; in the cruising stage, the supersonic main aircraft is separated from the subsonic babysitter, the supersonic main aircraft flies independently, and the subsonic babysitter returns. Through the innovative sectional type structural design, combination and configuration of the propelling system and power distribution and switching integration in the flight stage, efficient, safe and environment-friendly flight of the supersonic commercial airliner is achieved.

The invention provides a supersonic jet adopting supersonic double wings and belongs to the field of novel supersonic civil aircrafts, an S-shaped bent large slenderness ratio variable cross-section fuselage is adopted as a fuselage, large-area truncated delta wings with edgings on the inner sides are adopted as inner sections of wings, engine nacelles are arranged on the upper portions of the outer ends of the wings, the small-area supersonic double wings are installed on the outer side of the engine nacelle, the tail wing adopts the design of a single vertical tail and an inverted V tail, the acoustic explosion intensity can be reduced, meanwhile, the functions of the wing inner section, the supersonic double wings and an engine nacelle supporting structure are achieved, the structural efficiency is effectively improved, and the structural weight is reduced. By the adoption of the supersonic double-wing design, it can be guaranteed that the aircraft achieves ideal environment-friendly performance and aerodynamic efficiency, the front edge and the rear edge of the supersonic double wings are movable, the problem that the performance of non-design points of conventional double wings is poor is solved, and the aircraft can have high aerodynamic performance within the flight envelope range.

Apparatuses and systems are provided herein for unducted propulsion systems. The system includes an aft housing for low drag for high subsonic sustained flight. A plurality of blades are affixed to the aft housing, wherein the housing defines a flowpath curve extending from the axial extent of the aft blade root to the aft end of the aft housing. The flowpath curve is described by an axial direction parallel to an axis of rotation and a radius from the axis of rotation. The flowpath curve includes first point having a first radius where the radius reaches a maximum aft of the aft blade root and a second point forward of the first point having a second radius where the radius stops decreasing. The ratio of the first radius to the second radius is greater than or equal to 1.081.

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.