33results about "Heat reducing structures" patented technology

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.

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 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.

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.

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 invention provides a wingtip vortex active control method and system based on surface pressure feedback and synthetic jet and a medium, and the method comprises the steps: collecting pressure data in real time based on a pressure sensor array arranged on the surface of a wing; performing time-frequency analysis on the pressure data, and extracting the shedding frequency and pressure pulsation characteristics of the current wingtip vortex; retrieving flow field characteristics based on the wingtip vortex optimal disturbance modal database of the pre-solution analysis to obtain dimensionless information; converting the dimensionless information into driving frequency, driving voltage and a phase instruction of a synthetic jet actuator to generate periodic jet; continuously monitoring the power spectrum density and suction peak change of a pressure sensor at a specific position of a wing, and judging whether a flow field meets a rapid instability condition or not; and if the rapid instability condition is not met, the frequency or phase difference of the synthetic jet is adjusted, rapid capture and closed-loop control from flow field data to optimal disturbance application can be achieved, and instability and rapid attenuation of the wingtip vortex are effectively promoted.

A heated leading edge structure (14) for an aircraft (10) including a leading edge panel (20) having an outer surface (32) configured to contact 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 arrangement (24) configured to heat the leading edge panel (20); wherein the heating arrangement (24) includes an indirect heating device (44) mounted on the reinforcing member (22) and configured to heat the leading edge panel (20) by conduction through the reinforcing member (22).

The invention discloses an ultrathin composite titanium-based material high-temperature-resistant and high-strength connecting structure and application. The connecting structure comprises a main bearing core body, a conical cushion block, an indexing thread locking block, an indexing inclined cushion block, a locking nut and an elastic gasket. The two ends of the main bearing core body are of 60-degree cone structures and are matched with the conical face of the main bearing core body through a conical cushion block and an indexing cushion block assembly formed by splicing an indexing thread locking block and an indexing inclined cushion block, clamping of the ultra-thin-wall-thickness interlayer is achieved under the pretightening force effect of an elastic gasket, and finally locking and fixing are conducted through a locking nut. All the components are made of continuous SiC fiber reinforced titanium-based composite materials through hot press molding. Through the synergistic effect of conical surface matching and elastic pre-tightening, uniform distribution of stress at the connecting position is achieved, damage to an ultrathin interlayer is effectively avoided, and the connecting piece has the advantages of being resistant to high temperature (900-1000 DEG C), high in strength (sigma b is larger than or equal to 700 MPa-900 MPa), capable of preventing falling, locking, high in vibration resistance and the like, and is particularly suitable for connection of a high-speed aircraft composite material interlayer structure.

This invention relates to a self-driven active cooling system for aircraft control surfaces, belonging to the field of aircraft thermal protection technology. It includes: a pressurization pipeline and a jet pipeline disposed on the control surface; the pressurization pipeline contains a phase-change working fluid, and the jet pipeline has nozzles facing the high heat flux region of the control surface; a vapor collection chamber and a liquid storage chamber disposed on the fuselage, separated by a partition; an elastic reset member disposed in the liquid storage chamber, acting on the partition; the vapor collection chamber is connected to the outlet of the pressurization pipeline and the inlet of the jet pipeline; the liquid storage chamber is connected to the inlet of the pressurization pipeline through a return pipeline. The system of this invention can achieve self-driven, high-efficiency heat exchange.

The application discloses a transonic buffet control structure based on a wing trailing edge air vent, and belongs to the technical field of aircraft flow control. The buffet control structure is an air vent arranged in the wing trailing edge. One end of the air vent is communicated with the upper surface of the airfoil, and the other end is communicated with the blunt trailing edge of the airfoil. The aperture of the air vent and the thickness of the trailing edge of the airfoil are of the same order of magnitude. The application can control the buffet while reducing the influence on the lift-drag characteristics of the original wing.

The application discloses an aircraft wing anti-chattering device and system and belongs to the field of aircraft aeroelastic control. The device comprises a hollow tube vertically and penetratively installed in the inside of the wing, a counterweight arranged in the tube, upper and lower springs respectively abutting the upper and lower ends of the counterweight and having a pre-compression amount, a damping mechanism and a limiting mechanism. The counterweight and the spring constitute a tuned mass system, the total stiffness of which satisfies k total =m×(2πf w ) 2 , accurate alignment with the wing chattering frequency is realized, the damping mechanism dissipates vibration energy, and the limiting mechanism prevents damage caused by the limit stroke. Furthermore, a plurality of devices can be arranged along the spanwise direction and the chordwise direction to form an array system, multi-order mode and bending-torsional coupling chattering are suppressed through differential frequency tuning and staggered installation.

An aircraft wing, including a main wing and a trailing edge high lift assembly movably arranged at a main wing trailing edge. The trailing edge high lift assembly includes a flap and a connection assembly movably mounting the flap to the main wing, the flap being movable between retracted and extended positions. In the extended position, a gap is present between a main wing lower rear panel and a flap leading-edge section. In the retracted position, the gap is closed or minimized. The flap is movable between the extended position, a first retracted position, and a second retracted position where the flap is moved beyond the first retracted position in a forward direction, and the wing includes a gap closing device configured to close or minimize the gap in both the first and second retracted flap positions, and to open the gap in the extended flap position.

The invention relates to a flying wing unmanned aerial vehicle which comprises a vehicle body and wings arranged on the two sides of the vehicle body, the section airfoil of each wing is a laminar flow airfoil, the torsion angle of each wing is negative torsion, and the angle range of the negative torsion is 5-10 degrees. According to the flying-wing unmanned aerial vehicle disclosed by the invention, the cruise lift coefficient and lift-drag ratio of the flying-wing layout are greatly improved, and the endurance performance of the low-speed flying-wing unmanned aerial vehicle is improved.

The utility model belongs to the technical field of aircraft structure components, and relates to aircraft wing wallboard forming equipment with a reinforcing rib structure. The wing wall plate comprises a wing wall plate main body, the wing wall plate main body is provided with a buckle connecting piece and a screw joint fixing piece, a wall plate positioning part and an installation alignment part are arranged in the wing wall plate main body, the wall plate positioning part and the installation alignment part are arranged in a staggered mode, and a splayed reinforcing rib assembly is arranged at the bottom of the wing wall plate main body. And the wall plate positioning part and the mounting alignment part are respectively staggered with the splayed reinforcing rib assembly. By means of the splayed reinforcing rib assembly at the bottom of the wing wall plate body, especially the T-shaped reinforcing structure, bending loads can be effectively resisted, the overall bending rigidity of the wall plate is remarkably improved, deformation caused by bearing complex aerodynamic loads in the flying process is reduced, and the aerodynamic performance of an aircraft wing is guaranteed.

An aircraft includes a wing having a longitudinal centerline axis, a leading edge, a trailing edge aft of the leading edge, a wing thickness, and a wing chord, an engine having a longitudinal axis vertically aligned with the wing chord and parallel to the longitudinal centerline axis, a pylon connecting the wing to the engine, and a fairing received over the pylon. The fairing defines a horizontal plane, a first plane perpendicular to the longitudinal axis, and a second plane perpendicular to the longitudinal axis, the second plane being aft of the first plane, the longitudinal axis defining a reference line when projected onto the horizontal plane. The fairing includes a fairing body defining an aerodynamic surface having an outboard portion and an inboard portion 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. The second distance is greater than the first distance.

The invention relates to an aircraft (1002) comprising a wing (1001) having a fixed wing (1005) and a wing tip device (1003) movably mounted at an outer end of the fixed wing. The wing tip device (1003) is movable between a flight configuration and a ground configuration. The wing tip device (1003) and the fixed wing (1005) are separated along a tilted main cut plane (1013). The wing tip device (1003) and the fixed wing (1005) meet along an interface cut line (1035). The wing tip device and the fixed wing comprise a wing skin having a thickness and an end face extending across the thickness of the wing skin provides an interface surface corresponding to the interface cut line, wherein the interface surface is tilted in a first orientation towards a front of the wing and in a second, opposite orientation towards a rear of the wing. The invention also relates to a wing of an aircraft, a fixed wing and a wing tip device.

A functionally improved fairing arrangement (26) for a high-lift mechanism (20) of an aircraft (10) has been described. The high-lift mechanism (20) includes a flap (22) to be arranged at a trailing edge (16) of an aircraft wing (12) and a mounting and guiding mechanism (24) for the flap (26). The fairing arrangement (26) comprises a flap side fairing unit (28) and a wing side fairing unit (30). The flap side fairing unit (28) comprises a flap side fairing (32) for covering an aft part of the mounting and guiding mechanism (24) and a flap side fairing mount (34) for mounting the flap side fairing (32) to the flap (22). The wing side fairing unit (30) comprises a wing side fairing (36) for covering a forward part of the mounting and guiding mechanism (24) and a wing side fairing mount (38) for mounting the wing side fairing (36) to the wing (12). The wing side fairing mount (38) is configured for movably connecting the wing side fairing (36) to the wing (12), preferably such that the wing side fairing (36) is rotatable around an axis (39) directed at least partially in a vertical direction. A fairing joint (80) couples the flap side fairing (32) and the wing side fairing (36) in order to transfer movements of the flap side fairing (32) to the wing side fairing (36). Hence, a lateral or rotational movement of the wing side fairing (36) is controlled by the flap side fairing (32).

A wing system for an aircraft is provided, the wing system having a stowed configuration, a pre-deployed configuration, and a deployed configuration. The wing system includes two wings, each having an airfoil shape and deployable to pivot about a respective pivot axis between a pre-deployed configuration and a deployed configuration. In the stowed configuration, the two wings are in a first substantially overlapping spatial relationship with each other and are retractable within an envelope having an envelope cross-sectional shape and a corresponding envelope cross-sectional area. In the pre-deployed configuration, the two wings are in a second substantially overlapping spatial relationship with each other and are deployable to the deployed configuration. In the deployed configuration, each of the wings is capable of generating aerodynamic lift in an airflow. Each airfoil shape of each wing is a slotted airfoil having a primary element, a secondary element, and a chord, the secondary element being pivotable relative to the primary element and spaced apart by a gap. Each airfoil shape has a respective maximum thickness and a respective maximum absolute thickness. In the stowed configuration, each secondary element of each airfoil on one wing is set at a different flap angle compared to each secondary element of each airfoil on the other wing.

The application relates to a high-speed aircraft thermal management system design method based on a pneumatic heat utilization technology. The system comprises a thermal control box arranged in a control surface, a cooling channel arranged at the bottom end of the control surface and a cooling assembly arranged in a fuselage. A gap is arranged between the bottom end of the control surface and the fuselage, and the cooling channel is provided with an exhaust port facing the gap. One end of the cooling assembly is communicated with the thermal control box through a pipeline, the pipeline is provided with a pressure control member, and the other end is communicated with the cooling channel. The thermal control box is provided with a cooling working medium, and the cooling working medium absorbs the heat of the control surface to generate phase change. The application can realize the multiple functions of leading edge drag reduction, heat protection, semi-active thermal control of the gap and connecting members and the like, and effectively ensure the safe and stable working temperature of the servo mechanism.

The present invention relates to a drive mechanism for changing the tilt angle of a winglet, the drive mechanism comprising a linear actuator, one end of which is fixed to the wingtip and configured to output a linear load; a conversion mechanism connected to the linear actuator and configured to rotate in response to the linear load to output torque; and a transmission mechanism connected to the conversion mechanism and the winglet to transmit torque to the winglet to change the tilt angle of the winglet.

An airfoil profile for an aircraft, such as an unmanned aerial system. The airfoil profile includes a leading edge portion and a trailing edge portion that are spaced apart along a chordwise direction, an airfoil centroid; and an upper airfoil surface and a lower airfoil surface. The airfoil surfaces are shaped such that the pressure center of the lifting force is arranged at the same chord location as the airfoil centroid or closer to the trailing edge portion along the chordwise direction than the airfoil centroid. With this, a pitch-down momentum is generated that urges the leading edge portion towards a lower angle of attack.

The invention relates to the field of aeronautical engineering and aerodynamic noise control, and provides a noise reduction structure for a flap side edge. The noise reduction structure for the flap side edge comprises a three-dimensional gradient front edge flow guide structure which is arranged at a front end inlet of the flap side edge; the side edge vortex blocking baffle is of a sheet-shaped structure and is arranged on the downstream of the three-dimensional gradual change front edge flow guide structure; the sawtooth-shaped upper edge vortex control structure is arranged on the upper edge, away from the flap airfoil, of the side edge vortex blocking baffle. According to the noise reduction structure for the flap side edge, the speed difference between the noise reduction structure and a root low-energy shearing layer can be reduced, and the forming condition of a primary large vortex is destroyed; the core source of the broadband noise is solved, the high-frequency noise peak is avoided, and the broadband noise intensity is greatly reduced.

The invention provides a winglet (22) capable of suppressing a wingtip vortex and slowing down the reduction of lift. A wingtip winglet mounted to a wingtip of an aircraft wing (2), comprising a winglet body (220) in which at least one stepped groove (2231, 2232, 2233) is formed on a surface on a trailing edge side of the winglet body, the stepped groove comprising: a plurality of short discontinuity surfaces (22311-22315) arranged along a first direction (X1) from a wing root of the wing toward a wingtip, the plurality of blades are arranged in a staggered manner one by one in a second direction (Y1) from the trailing edge side to the leading edge side of the wing; and a long discontinuity surface (2231B) extending from the short discontinuity surface closest to the wing tip side in the first direction toward a third direction (Y2) opposite to the second direction, the length of the long discontinuity surface being not less than the total length of the plurality of short discontinuity surfaces in the second direction or the third direction.

An aircraft includes a wing having a longitudinal centerline axis, a leading edge, a trailing edge aft of the leading edge, a wing thickness, and a wing chord, an engine having a longitudinal axis vertically aligned with the wing chord and parallel to the longitudinal centerline axis, a pylon connecting the wing to the engine, and a fairing received over the pylon. The fairing defines a horizontal plane, a first plane perpendicular to the longitudinal axis, and a second plane perpendicular to the longitudinal axis, the second plane being aft of the first plane, the longitudinal axis defining a reference line when projected onto the horizontal plane. The fairing includes a fairing body defining an aerodynamic surface having an outboard portion and an inboard portion 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. The second distance is greater than the first distance.