131results about "Heat reducing structures" patented technology

A heat exchanger system for an aircraft comprises nacelle having an inlet and an outlet, a nacelle heat exchanger within the nacelle, and a fan system within the nacelle. Air flows into the nacelle through the inlet and out of the nacelle through the outlet. The nacelle heat exchanger is configured to transfer heat away from a coolant using the air in an airflow to the nacelle heat exchanger. The fan system is configured to increase the airflow to the nacelle heat exchanger.

The invention belongs to the field of electric vertical take-off and landing aircrafts, and discloses a novel vertical take-off and landing remote high-speed manned electric tilt-rotor aircraft structure which comprises an aircraft body, wings, a V-shaped empennage, six supporting arms, six tilt-rotor units, an energy system, an avionics system and an undercarriage. The six tilting rotor wing units are symmetrically installed on the wings and the V-shaped empennage through the six supporting arms. The aircraft is mainly characterized in that the design of six tilting rotor units is adopted, so that the aircraft has the advantages of vertical take-off and landing, high-speed forward flight, no structural'dead weight 'and useless power consumption; the design that the supporting arms extend out of the wings and the empennage to exceed the radius length of the rotors is adopted, and the downward projection of a rotor disc is not overlapped with the wings and the empennage in the vertical flight state, so that downward washing flow of the rotors cannot generate downward loads on the wings and the empennage, and power consumption for overcoming the downward loads is avoided; the design of the wings with high aspect ratio and winglets is adopted, so that the forward flying lift-drag ratio and the cruising efficiency are improved, and the aircraft has high-speed forward flying and remote flying capabilities; and the six tilting rotor units adopt the same tilting mechanism, so that the complexity of design and manufacturing is reduced.

The thermal protection structure comprises a first porous medium layer arranged on the arc-shaped wall face of the front edge of the aircraft and a plurality of second porous medium layers arranged on the upper wall face and the lower wall face of the front edge of the aircraft at intervals in the flow direction. Solid walls are arranged between the first porous medium layers and the second porous medium layers and between every two adjacent second porous medium layers; the porosity of each second porous medium layer is firstly increased and then reduced along the flow direction, and the porosity of the first porous medium layer is smaller than that of the second porous medium layer at the most upstream. The method is applied to the field of thermal protection of aircrafts, the problem of local overheating existing in a traditional sweating cooling technology is effectively solved through the synergistic effect of the intermittent sweating layout and the gradient porosity material, and the situation that the cooling efficiency is reduced due to overheating of a cooling agent is avoided.

An aircraft includes a wing and a leading-edge aerodynamic surface coupled to the wing. An actuator is disposed within the wing and coupled to the leading-edge aerodynamic surface via a linkage system. Two or more hinge systems are coupled to the leading-edge aerodynamic surface and configured to rotate about a fixed axis internal to the wing to move the leading-edge aerodynamic surface between a retracted position and a drooped position.

The invention discloses a flight-oriented fan wing trailing edge adjusting device, and relates to the technical field of aircrafts, the flight-oriented fan wing trailing edge adjusting device comprises a trailing edge angle adjusting mechanism, a rectification angle adjusting mechanism, a circular end plate and a rectification plate, the trailing edge angle adjusting mechanism comprises a steering engine, a first connecting rod mechanism, a trailing edge plate and a rotating shaft; the steering engine is fixedly installed on the circular end plate and controls the rear edge plate to move through the first connecting rod mechanism, the rotating shaft is arranged on one side of the rear edge plate, and the rear edge plate is hinged to the rectifying plate through the rotating shaft. By freely adjusting the trailing edge angle and the rectification angle, the aerodynamic force of the fan-wing aircraft is changed, the lift force and thrust distribution is improved, the aerodynamic effect is enhanced, the energy consumption is reduced, and the flight efficiency is improved. The fan wing trailing edge adjusting device is applied to a serial fan wing, forward coupling gain between two wings is achieved by changing the size of a trailing edge angle and a rectification angle, different flight conditions are met, various flight requirements are flexibly met, and the load performance and the endurance are optimized.

The invention discloses a wing end standing vortex lift augmentation device which comprises a device upper surface, a device lower surface, a leading edge surface and a wing profile, a wing upper wing surface extends outwards to form the device upper surface, a wing lower wing surface extends outwards to form the device lower surface, and the extension length of the device upper surface is larger than that of the device lower surface. A distance exists between the upper surface and the lower surface of the device, and wingtips do not coincide with each other; the device upper surface, the device lower surface, the front edge surface and the airfoil profile jointly form a semi-closed cavity, when incoming flow exists in the semi-closed cavity, a wing end standing vortex rotating in the direction away from the central axis of the fuselage is formed in the semi-closed cavity, and the rotating direction of the wing end standing vortex is opposite to that of a wingtip vortex. According to the three-dimensional wing, airflow on the lower wing surface of the wing is prevented from flowing to the upper wing surface of the wing through the wing end standing vortex in the semi-closed cavity, so that the trend that the airflow on the lower wing surface of the wing is reversed upwards can be inhibited under the condition of not depending on physical obstruction, the lift force of the three-dimensional wing is effectively increased, and the induced resistance is reduced.

The invention discloses a variable sweepback wing based on flexible corrugated skin and a design method, and relates to the field of aircraft structure design, the variable sweepback wing comprises a wing mounting seat matched with a fuselage, and further comprises a rigid edge strip matched with the outer edge of the variable sweepback wing; the rigid edgings are arranged up and down in space, and the wing mounting seats and the rigid edgings are connected into an integrated structure so as to construct a flexible corrugated skin I and a flexible corrugated skin II of the sweepback wing; the supporting assemblies are arranged between the flexible corrugated skin I and the flexible corrugated skin II; the driving mechanism is arranged on the wing mounting seat and is used for pulling the flexible corrugated skin to deform; and the linear displacement sensor is used for measuring the variable quantity of the sweepback angle of the wing. According to the variable-sweepback wing based on the flexible corrugated skin and the design method, the aerodynamic configuration of an aircraft can be adjusted along with the flight path, the cross-speed domain aerodynamic configuration contradiction is effectively solved, the higher available lift-drag ratio is provided for the aircraft, and the comprehensive aerodynamic performance of the aircraft is further improved.

The invention discloses a low-frequency broadband sound absorption metamaterial adaptive wing and a design method thereof, and belongs to the technical field of aircraft design. The problem that in the prior art, a traditional aircraft wing and a design method thereof are difficult to achieve the low-frequency and broadband noise reduction effect is solved. The wing comprises a leading edge inner side slat, a leading edge outer side slat, a main wing, a trailing edge inner side flap and a trailing edge outer side flap, the leading edge inner side slat and the leading edge outer side slat are connected with the front side of the main wing, and the trailing edge inner side flap and the trailing edge outer side flap are connected with the rear side of the main wing; the metamaterial sound absorption strips are embedded into concave cavities of the leading edge inner side slat and the leading edge outer side slat respectively, the metamaterial sound absorption strips are embedded into the leading edge of the main wing, each metamaterial sound absorption strip is composed of a Helmholtz resonator array composed of single-layer Helmholtz resonator units, and each metamaterial sound absorption strip is provided with a micro-perforated surface with continuous curvature. The noise reduction level of the wing is effectively improved, the sound absorption frequency band is wider, and the method can be applied to sound absorption wing design.

The invention relates to the field of aircrafts, in particular to a portable single-person low-altitude vector jet aircraft with a folding wing structure, the portable single-person low-altitude vector jet aircraft comprises a rack part, a power part and a load part, the power part is arranged below the rack part, and the load part is arranged below the power part; according to the portable single-person low-altitude vector jet aircraft with the folding wing structure, through the design of the telescopic framework of a multi-connecting-rod structure, the aircraft can be easily folded when not used, the occupied space is small, carrying and storage are convenient, and the problems that a traditional single-person aircraft is large in size and difficult to carry are solved; the applicability and convenience of the aircraft are improved; meanwhile, the structural strength of the wings in the unfolded state is ensured through the multi-point supporting design of the telescopic framework, the stability and safety of the aircraft are improved, the air resistance of the aircraft is reduced through the design of the wing skin, the protective cover, the conical cover, the wind tunnel and the fairing, the aerodynamic performance is effectively improved, and the aircraft is more efficient in the flight process.

An aircraft flap comprises an upper wing surface and a lower wing surface which are opposite to each other, and a cavity is defined by the upper wing surface and the lower wing surface; the noise reduction device is arranged in the cavity and comprises a first noise reduction structure, the first noise reduction structure comprises a perforated plate and an upper part and a lower part, the upper part is embedded in the upper wing surface, and the lower part is embedded in the lower wing surface; a second noise reduction structure including foam metal, disposed between the upper portion and the lower portion of the first noise reduction structure, and attached to the first noise reduction structure; and the third noise reduction structure comprises a super-structure partition plate and is embedded into the second noise reduction structure. According to the composite superstructure design integrating the three noise reduction structures together, the structural stability of the noise reduction device can be guaranteed, and meanwhile, the jet flow installation noise of the aircraft engine can be effectively reduced.

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 variable-pore phase-change sweating cooling bionic structure based on fractal characteristics. The variable-pore phase-change sweating cooling bionic structure comprises a variable-pore porous sweating layer and a regulation and control unit, wherein the variable-pore porous sweating layer is composed of fractal flow equalizing plates which are sequentially arranged in a stacked mode in the heat flow direction; the fractal flow equalizing plate is of an N-stage bifurcated tree-shaped structure and is used for equally dividing one path of cooling working medium into 2N paths of cooling working medium, and N is an integer ranging from 4 to 6; the variable-pore porous sweating layer is divided into an upper-layer high-heat-flow area and a lower-layer low-heat-flow area in the thickness direction, the upper-layer high-heat-flow area is used for promoting phase change and efficient heat exchange of a cooling working medium, and the lower-layer low-heat-flow area is used for promoting liquid-phase permeation of the cooling working medium; and the regulation and control unit predicts the phase change leading edge position S (t) in real time based on a neural network, and dynamically regulates the cooling working medium flow Q (t) input into the inlet of the fractal flow equalizing plate according to the S (t). The invention further discloses a design method of the structure. By introducing the bionic variable porosity characteristic, coupling of flowing heat exchange is achieved, the cooling efficiency of thermal protection is improved, and consumption of a heat exchange working medium is reduced.

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.

The invention provides a flexible composite material skin and a preparation method thereof. The flexible composite material skin comprises a zero Poisson's ratio skeleton, a negative Poisson's ratio filler and an elastic surface layer, the zero-Poisson-ratio framework is provided with a plurality of filling areas; the negative Poisson's ratio filler is arranged in the filling area; the elastic surface layer is attached to the upper surface and / or the lower surface of the zero-Poisson-ratio framework with the negative-Poisson-ratio filler. According to the flexible composite material skin provided by the invention, the surface can be kept continuous, smooth and free of concave-convex fluctuation in the deformation process, and the stealth performance and aerodynamic performance of a corresponding aircraft structure prepared from the flexible composite material skin can be improved.

A winglet system is configured to be secured to a wing of an aircraft. The winglet system includes an upper blade, a lower blade, and a cover plate integrally formed with the upper blade and the lower blade. The cover plate covers an opening within a transition joint between the upper blade and the lower blade. The cover plate remains coupled to the upper blade if the lower blade is separated from the winglet system.

The invention relates to a heated leading edge structure for an aircraft, a slat, a wing and an aircraft. A heated leading edge structure (14) for an aircraft (10) includes a leading edge panel (20) having an outer surface (32) configured for contact with an ambient flow (A) and an inner surface (34) opposite the outer surface (32); a reinforcing member (22) connected to the inner surface (34) of the leading edge panel (20); and a heating device (24) configured to heat the leading edge panel (20); the heating device (24) comprises at least one indirect heating device (44) mounted on the reinforcing member (22), said indirect heating device (44) being configured for heating the leading edge panel (20) at least by means of conduction through the reinforcing member (22).

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

Disclosed is a wing (3) for an aircraft (1), comprising a main wing (5) and a trailing edge high lift assembly (9) movably arranged at a trailing edge of the main wing (5), the trailing edge high lift assembly (9) comprising a flap (11) and a connection assembly movably mounting the flap (11) to the main wing (5), such that the flap (11) is movable between at least one retracted position (12, 14) and at least one extended position (15), wherein in the at least one extended position (15) a gap (16) is present between a lower rear panel (17) of the main wing (5) and a leading edge section (19) of the flap (11), and wherein in the at least one retracted position (12, 14) the gap (16) is closed or minimized. The object of the invention, to provide a wing that might be used as a morphing wing capable of adjusting wing area, curvature, and angle of attack during the entire flight, while having optimized aerodynamic performance throughout all flap settings, is achieved in that the flap (11) is movable between at least one extended position (15), a first retracted position (12), and a second retracted position (14) where the flap (11) is moved beyond the first retracted position (12) in a forward direction (13), and the wing comprises a gap closing device (23) configured to close or minimize the gap (16) in both the first retracted position (12) and the second retracted position (14) of the flap (11), and to open the gap (16) in the at least one extended position (15) of the flap.

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 method includes obtaining (402) thermal energy from a structure (112) to be cooled, the structure including microchannels (208). The method also includes supplying (402) the thermal energy to a water-based polymer network (206), the water-based polymer network including a gel formed using a polymer and water. The method further includes generating (404) one or more gases (210) by heating the water-based polymer network, the generating one or more gases including releasing water in the water-based polymer network to produce steam. Additionally, the method includes passing (408) the one or more gases through the microchannels to remove at least a portion of the thermal energy from the structure.

The invention discloses a distributed thermoelectric element variant wing SMA driver and method and application, the SMA driver comprises an active cooling layer, an SMA wire driving structure and an elastic supporting layer which are sequentially arranged from top to bottom, each layer is embedded in a flexible material in a sealed mode, the active cooling layer is composed of a plurality of thermoelectric elements arranged in an array mode, and the elastic supporting layer is embedded in the active cooling layer. The thermoelectric element achieves refrigeration of the SMA wire driving structure through the Peltier effect. The cooling efficiency can be remarkably improved, heat accumulation is eliminated, the recovery time of the driver is shortened, the driving frequency is improved, and the service life of the shape memory alloy driver is prolonged. According to the algorithm, an adaptive thermoelectric element layout scheme can be automatically generated in combination with the structural characteristics of the driver, so that heat transfer is accelerated, the uniformity of a temperature field is improved, the application requirements under multiple working conditions and complex environments are met, the algorithm has good universality, the integration requirements of different types of wings are met, and the application prospect is wide. The design process is effectively simplified, and the manufacturing cost is reduced.

The present invention relates to an aircraft (1002) comprising a wing (1001) having a fixed wing (1005) and a wingtip device (1003) movably mounted at the outer end of the fixed wing. The wingtip device (1003) is capable of moving between a flight configuration and a ground configuration about a rotation axis. The wingtip device (1003) abuts the fixed wing (1005) along an interface cut line (1035). The interface cut line (1035) includes a curved portion (1039) centered on the rotation axis (1011), the radius of the curved portion increasing continuously as the cut line moves about the axis (e.g., a spiral shape). The present invention also relates to a method of designing a wing, a fixed wing, a wingtip device, and an interface.

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.

This application discloses a foldable, tilted wingtip with an aerodynamic device. The disclosed exemplary wing for use with an aircraft includes a fixed portion and a folding portion proximate a distal end of the wing. The folding portion includes an inclined surface. The wing also includes at least one of a feather or a winglet, and a hinge operably coupled between the fixed portion and the folding portion to enable the folding portion to fold relative to the fixed portion.

This invention proposes an aerodynamic drag device and layout for a tailless aircraft. The aerodynamic drag device comprises three drag plates arranged side-by-side on the aircraft wing: an inner plate, a middle plate, and an outer plate. One side of each of the inner, middle, and outer plates is rotatably connected to the wing. The middle plate is located between the inner and outer plates, with the inner plate closer to the aircraft's centerline. The layout includes a blended fuselage and wing design, with a pair of ailerons and the aforementioned aerodynamic drag device symmetrically arranged at the wing's trailing edge. The opening angle of the pair of aerodynamic drag devices is independently controlled. This invention increases the aircraft's directional stability, reduces supersonic drag, improves supersonic handling performance, enhances cruise and maneuverability, and does not introduce additional adverse effects on the airframe structure.

The present invention provides a flap vortex disturbing device and a wing assembly (10) comprising the flap vortex disturbing device, which can disturb vortex generated at the inner and outer side end edges of a flap, thereby suppressing the reduction of lift force and the advance of stall of an aircraft, and a wing assembly (10) comprising the flap vortex disturbing device. The flap vortex interference devices (120 and 140) interfere vortexes of trailing edge flaps (210 and 220) arranged on the trailing edge of a fixed wing (100) when an aircraft flies at a high angle of attack, and the flap vortex interference devices are correspondingly arranged at the inner and outer side end edges (211 and 221) of the trailing edge flaps respectively and are in an obtuse triangle shape. And gaps for air flow to circulate are respectively formed between the flap vortex interference devices (120 and 140) and the inner and outer side end edges (211 and 221) of the trailing edge flap.

An aircraft morphable aerodynamic surface structure including an elastomeric material having a glass transition temperature, a temperature management system for heating or cooling the elastomeric material, and a controller to control the temperature management system, the controller configured to control the temperature management system on the basis of a command received at the controller.

To provide fairings to support engines relative to wings of an aircraft, and methods for reducing drag.SOLUTION: A wing comprises a fairing 130 covering a pylon to connect to an engine. The fairing defines a horizontal plane, a first plane perpendicular to a longitudinal axis, and a second plane perpendicular to the longitudinal axis and aft of the first plane. A reference line is defined as the longitudinal axis when projected onto the horizontal plane. The fairing includes a fairing body defining an aerodynamic surface. The aerodynamic surface has an outboard portion and an inboard portion. The inboard portion is configured such that the first plane intersects the horizontal plane and the aerodynamic surface of the inboard portion at a first intersection point. The first intersection point is laterally displaced from the reference line by a first distance. The second plane intersects the horizontal plane and the aerodynamic surface of the inboard portion at a second intersection point. The second intersection point is laterally displaced from the reference line by a second distance, which is greater than the first distance.SELECTED DRAWING: Figure 2

The present application relates to a large-sized heat-resistant load-bearing integrated aircraft wing, belonging to the technical field of aircraft structures. The large-sized heat-resistant load-bearing integrated aircraft wing includes a fuselage shell. On both sides of the fuselage shell, a first installation groove and a second installation groove are provided. One end of the front wing facing the middle wing is provided with a first groove, and on the middle wing, there is a first lug that is inserted in cooperation with the first groove. One end of the middle wing facing the rear wing is provided with a second groove, and on the rear wing, there is a second lug that is inserted in cooperation with the second groove. The front wing, the middle wing, and the rear wing are connected between the first installation groove and the second installation groove through a front support and a rear support. Heat-resistant putty is arranged at the butt joint between the front wing, the middle wing, and the rear wing; in the present application, through the movable installation of the rear support, the connection structure between it and the fuselage shell can be thermally deformed and matched, avoiding the completely fixed installation method. When the wing generates thermal deformation, it cannot be extended synchronously, resulting in the wing being easily damaged.

The application discloses a hydrogen-electric hybrid energy supersonic passenger plane, and belongs to the field of engineering technology.The plane comprises an aerodynamic shape, a hydrogen-electric hybrid driven power device and an integral hydrogen tank for storing hydrogen fuel, wherein the aerodynamic shape comprises an elongated body, a wing containing a lifting surface, a tail wing and a canard; the power device comprises a gas turbine main engine and an electric propulsion ducted fan; the integral hydrogen tank comprises a pressure-bearing inner tank, an adiabatic layer and a support structure; the power device is connected with the body and the wing; and the hydrogen tank is connected with the body through the support structure.The scheme increases the lifting surface on the wing to ensure the high lift-drag ratio and low sound blast requirement of the plane, introduces hydrogen energy to realize zero carbon emission, and adopts the integral hydrogen tank to increase the hydrogen fuel storage of the plane and guarantee the voyage requirement of the plane.