188results about "Sandwich constructions" patented technology

An aircraft component including a structural foam having a predetermined shape, an outer flange located near a first end of the structural foam, an inner flange located near a second end of the structural foam, a web skin encapsulating the inner flange and at least a portion of the structural foam wherein the structural foam is affixed to the web skin, wherein the aircraft component comprises a portion of a blended wing body aircraft (BWB), the BWB having no demarcation between a wing and a main body of the BWB.

The invention discloses a distributed electric propulsion box type wing ground effect aircraft which comprises a main fuselage, the two sides of the main fuselage are connected with one ends of sweepback wing main wings, one ends of two empennages above the tail of the main fuselage are connected with the main fuselage through vertical tails, and the other ends of the two empennages are connected with the other ends of the two main wings through end plates respectively. The empennage is connected with a plurality of high ducted fans directly driven by a motor, an elevator and an aileron are arranged at the rear part of the empennage, a rudder is arranged at the rear part of the vertical fin, and the elevator, the aileron and the rudder are controlled by a flight control system and are mutually independent. The ground-effect aircraft provided by the invention adopts a box-type wing structure formed by combining the sweepback main wing and the sweepforward empennage, has better aerodynamic performance in a take-off and climbing low-speed flight state, and can provide larger lift force and better maneuvering characteristics.

The invention relates to the field of aircraft wing design, in particular to a multi-stage telescopic wing for improving the flight performance of an aircraft, the extension length and the area of the wing are changed through multi-stage stretching of the wing, and different requirements of the aircraft for the wing in low-speed and high-speed flight states are met. When the aircraft takes off, lands and flies at a low speed, the telescopic wings stretch out to increase the area of the wings and enable the aircraft to obtain enough lift force, and when the aircraft flies at a high speed, the telescopic wings retract to reduce the extension length and area of the wings and enable the aircraft to obtain enough lift force and reduce the resistance of the wings, so that the extension length and area of the wings are changed through extension of the wings; and the aircraft can simultaneously meet different requirements on the extension length and the area of the wings in low-speed and high-speed flight states.

Aircraft ice protection systems and methods include an aircraft component having an internal layer and an external layer, a piezoelectric actuator element embedded within the aircraft component between material of the internal layer and the external layer, and a control system operably connected to the piezoelectric actuator element and configured to drive actuation of the piezoelectric actuator element to at least one of remove ice formed on the aircraft component and prevent the formation of ice on the aircraft component.

A blended wing body (BWB) aircraft comprising a first composite part comprising, a first dry carbon fiber sheet having a width greater than 3 inches and comprising a first layer comprising dry carbon fibers, a second layer disposed on top of the first layer comprising dry carbon fibers, a stitching extending through the first dry carbon fiber sheet, wherein the stitching is configured to bind the first dry carbon fiber sheet together, and a resin matrix, wherein the stitching and the first dry carbon fiber sheet are embedded within the resin matrix to create a single unified structure. BWB aircraft further comprises a blended wing body, having no clear demarcation along a leading edge of the BWB aircraft from a wing to a nose of the blended wing body, and wherein the first composite part comprises a portion of the blended wing body.

A composite assembly for a blended wing body aircraft, the composite assembly comprising: a first assembly comprising a first molded part, wherein the first molded part comprises a first plurality of fibers, the first molded part partially infused with a first resin, a second assembly oriented relative to the first assembly, the second assembly comprising a second molded part wherein the second molded part comprises a second plurality of fibers, the second molded part partially infused with a second resin, and a joining region, the joining region comprising at least a portion of first plurality of fibers and at least a portion of second plurality of fibers, the at least a portion of first plurality of fibers and the at least a portion of second plurality of fibers substantially adjacent and infused with a third resin.

A giant airliner structure comprises an airliner body and at least two layers of wings arranged on the two sides of the airliner body, a steering rudder is arranged between the upper and lower adjacent wings. The method has excellent environmental protection property and economical efficiency, and specifically comprises the steps that the core plate structure is adopted, and the labor cost needed in the production process is remarkably reduced; the structural design is simple, the maintenance difficulty is reduced, and the maintenance cost is reduced; the passenger-engine ratio is extremely low, and the manufacturing cost is greatly reduced; the weight ratio of the passenger plane is extremely low, and oil is greatly saved; the wings are short, the airport applicability is wide, and the opportunities of business are increased.

Stamp-forming a composite assembly includes applying a low-melt film on an interface surface of at least one spar, loading the spar onto at least one tooling segment such that the interface surface is exposed, and heating the tooling segment to a first temperature that is above a crystallization temperature of the low-melt film and below a melt temperature of a resin of the spar. A second low-melt film may be applied on a fiber-reinforced thermoplastic skin blank, and the skin blank is heated to a second temperature above a melt temperature of a resin of the skin blank. The skin blank is then draped around the tooling segment such that at least some of any second low-melt film contacts the low-melt film on the interface surface of the spar. The stamp press is then closed to co-consolidate the skin blank and the spar, thereby stamp-forming the composite assembly.

Aircraft ice protection systems and methods include an aircraft component (200) having an internal layer (202) and an external layer (204), a piezoelectric actuator element (208) embedded within the aircraft component (200) between material of the internal layer (202) and the external layer (204), and a control system operably connected to the piezoelectric actuator element (208) and configured to drive actuation of the piezoelectric actuator element (208) to at least one of remove ice formed on the aircraft component (200) and prevent the formation of ice on the aircraft component (200).

A backscatter x-ray system has an x-ray device, a movable platform, and a processor. The x-ray device includes an x-ray emitter and an x-ray detector. The movable platform positions the x-ray device relative to a frontside of a structure. The processor is communicatively couplable to the x-ray device. The x-ray emitter emits x-rays that penetrate the structure from the frontside. The x-ray detector detects a backscatter of the x-rays reflected from the structure. The processor generates x-ray images of the structure based on the backscatter, and analyzes the x-ray images in a manner facilitating at least one of manufacturing and inspection of the structure.

A method for producing a torsion box for a structure of an airplane. The method includes providing a first component made of a fiber composite material, the first component has a first planar base having a first inner side and a first outer side, first stiffening elements on the first inner side forming a composite with the first base. A second component is provided of a fiber composite material and has a second planar base having a second inner side and a second outer side. Second stiffening elements are on the second inner side and form a composite with the second base. The method includes superimposing the first component and the second component such that the first stiffening elements lie, at least in some areas, on the second inner side and the second stiffening elements lie, at least in some areas, on the first inner side. The methods includes connecting the first stiffening elements to the second base and connecting the second stiffening elements to the first base.

Composite multi-spar aircraft lifting surface comprising an integrated torsion box (1) comprising: - a first portion of a skin (3.1) continuous on the second face (1.2) from a right tip (4.1) to a left tip (4.2), - a second portion of a skin (3.2) continuous on the lateral right side (1.3) of the first face (1.1) from a right tip (4.1) to a first central portion (5), - a third portion of a skin (3.3) continuous on the lateral left side (1.4) of the first face (1.1) from the left tip (4.2) to a second central portion (6), - a cover (7) spanning between the first central portion (5) and the second central portion (6). The spars (2), the first portion of the skin (3.1), the second portion of the skin (3.2) and the third portion of the skin (3.3) being an integrated multi-cell part.

A composite structure has a first layer connect to a second layer to form a layered structure. The second layer has a plurality of base portions abutting the first layer and a plurality of beaded portions protruding from the plurality of base portions. Each beaded portion of the plurality of beaded portions defines a channel between the first layer and the second layer. A method for manufacturing a composite structure includes depositing composite material over a tool to form a second layer. The second layer has a plurality of beaded portions and a plurality of base portions. The method includes distributing a plurality of mandrels over the second layer to define channels in the plurality of beaded portions. The method includes depositing composite material over the second layer and the plurality of mandrels to form a first layer. The method includes bonding the first layer to the second layer.

The invention relates to an aerodynamic profile body with an outer flow surface for flow through a fluid, wherein the aerodynamic profile body comprises:a first surface layer and a second surface layer, which form at least part of the outer flow surface and are designed to be flexible in an end section of the aerodynamic profile body,a connection arrangement that flexibly connects the first surface layer to the second surface layer of the profile body in the end section, andan actuator device with at least one actuator provided in an interior of the aerodynamic profile body which interacts with at least one of the surface layers in the flexible end section in such a way that the cross-section of the aerodynamic profile body in the flexible end section can be changed by the at least one actuator,characterized in thatthe outer flow surface of at least one of the surface layers in the flexible end section has, at least in sections, a surface profile that extends parallel to the direction of flow.

The invention provides a non-exposed blade type aircraft which comprises an aircraft body (4), two front wings (3), two rear wings (5), two front propellers (2) and two rear propellers (1), the aircraft body (4) is located in the middle of the whole aircraft, the two front wings (3) are arranged at the front end of the aircraft and rotate together through a transmission shaft, and the two rear wings (5) are arranged at the rear end of the aircraft. The two rear wings (5) are arranged at the tail end of the aircraft and rotate together through a transmission shaft, the two front propellers (2) and the two front wings (3) are integrally designed, the two rear propellers (1) and the two rear wings (5) are integrally designed, and the two rear propellers (1) and the two rear wings (5) are integrally designed. The front thruster (2) and the rear thruster (1) are both designed to be free of exposed blades. According to the aircraft, the design without exposed blades is adopted, foreign objects can be prevented from being sucked in, the safety in the flight process is enhanced, and the flight efficiency is improved through the airflow multiplication effect.

Disclosed is an aircraft wing including a skin and a framework for supporting the skin. The framework comprises a spar extending along a length of the wing, the length extending between a root end and a tip end of the wing, a plurality of ribs extending chordwise between a leading edge of the wing and a trailing edge of the wing; and a plurality of stringers attached to the skin. The plurality of stringers comprises a first stringer extending along the length of the wing, wherein a proximal end of the first stringer is arranged nearer the root end and a distal end of the stringer is arranged nearer the tip end; and an auxiliary stringer extending non-parallel with the first stringer and arranged such that a portion of the auxiliary stringer is proximal to the distal end of the first stringer.

A panel is disclosed, including a first facesheet, a second face sheet, and a plurality of pultrusion-formed web structures. Each web structure has a middle support portion, a first end portion, and a second end portion. The first end portion of each web structure is attached to the first facesheet and the second end portion of each web structure is attached to the second facesheet. The middle support portion, first end portion, and second end portion of each web structure form a single monolithic structure.

The invention relates to an aircraft panel, an aircraft assembly, an aircraft wing, an aircraft and a method of manufacturing an aircraft panel. In one aspect, an aircraft panel is disclosed that includes a layer of composite material that is curved about an axis such that a free end of the layer extends greater than 180 degrees about the axis and defines a wall of a channel for receiving a hinge member of an aircraft component.

Systems, devices, and methods including a leading edge tubular member; an upper tubular member; a lower tubular member; one or more upper rib members connected between the leading edge tubular member and the upper tubular member; one or more lower rib members connected between the leading edge tubular member and the lower tubular member; a rigid sandwich shell disposed between the upper tubular member and the leading edge tubular member; and a sandwich shear web disposed between the upper tubular member and the lower tubular member; where the rigid sandwich shell and the sandwich shear web form a D-shape.

To provide an apparatus and method for producing skinned cell structures.SOLUTION: The device includes a placer to place foamable structures together to define a bundle, a restrainer to restrain the bundle, an activator to activate the foamable structures to expand and form a cell structure within the restrainer, a trimmer to trim the cell structure to define a core, and an assembler to couple a skin to the core to define the skinned cell structure.SELECTED DRAWING: Figure 2

The present invention relates to a conductive composite material and a method for manufacturing the conductive composite material. The conductive composite material includes a first elastic polymer layer, a conductive paste layer on the first elastic polymer layer, and a second elastic polymer layer on the conductive paste layer. A reinforcing mesh contacts the conductive paste layer.

A composite structure comprises stacked sets of laminated fiber reinforced resin plies and metal sheets. Edges of the resin plies and metal sheets are interleaved to form a composite-to-metal joint connecting the resin plies with the metal sheets.

A method of forming an aircraft component 800 comprising a curved composite panel 700 with a stiffener support region 709, the method comprises placing first (figure 5, 200) and second (figure 5, 300) fibre preforms in a mould to form a lay-up (figure 5, 400), the first preform comprises curved first surface (figure 5, 201), the second preform arranged against the first surface, infusing the preforms with resin 600 and curing to form the panel with the stiffener support region arranged contiguously with a curved region 707, wherein the support region is substantially planar 107’ along a first direction of the panel and the curved region curved along the same direction. The first surface of the first preform may be concave in the first direction and a first surface of the second preform may be convex. The panel may form a wing skin, the stiffener support region may be planar spanwise and curved in a chordwise direction. A foot 105 of a stiffener, such as a rib or stringer, may be attached to the support region. An aircraft component comprising a curved composite panel is also provided.

The invention provides a particle damper for vibration reduction of a wing structure, which comprises a bottom plate, a cover plate and a box body symmetrically arranged on the bottom plate, a plurality of damping particles and metal grids parallel to each other are placed in the box body, and a damping layer is attached to the inner wall of the box body. When the main beam vibrates due to external disturbance of the wing, the damping particles vibrate along with the main beam and collide and rub with the metal grating and the damping layer in the vibration process, vibration energy is effectively converted into heat energy, and therefore dissipation of the vibration energy is achieved, the vibration amplitude of the main beam of the wing is remarkably reduced under the condition that the structure of the wing is not changed, and the service life of the wing is prolonged. Not only is the influence of vibration on the performance of the wings effectively reduced, but also the purpose of noise reduction is achieved, and the mute performance and the riding comfort of the aircraft are improved.

The invention provides a method of punch forming a fiber reinforced thermoplastic composite component and a related system. Punch forming the composite assembly includes applying a low melt film on an interface surface of at least one spar, loading the spar onto at least one tooling section such that the interface surface is exposed, and heating the tooling section to a first temperature that is higher than a crystallization temperature of the low melt film and lower than a melting temperature of the spar resin. A second low melt film may be applied on the fiber reinforced thermoplastic skin blank and the skin blank is heated to a second temperature that is higher than the melt temperature of the skin blank resin. The skin blank is then covered around the tooling section such that at least some of the second low melt film contacts the low melt film on the interface surface of the spar. And then the punching machine is closed, so that the skin blank and the spar are jointly consolidated, and the composite assembly is formed through punching.

In order to enable influencing the temperature of an aircraft system with enlarged functionality while saving weight, the invention provides an aircraft structural component (14) comprising a load bearing foldcore sandwich element (22), e.g., made of a metal, a ceramic or a fibre reinforced composite material or hybrid materials, wherein the foldcore sandwich element (22) comprises a foldcore (24, 24.1, 24.2), e.g., of metal, of ceramics or of fibre reinforced composite material, fixed between a first and second plate (26, 26. i, 28) and is configured as a heat exchanger (16) including a first fluid conduct (30) for conducting a first fluid through the sandwich element (22) and a second fluid conduct (32) for conducting a second fluid through the sandwich element (22) wherein at least the first fluid conduct (30) comprises channels (34) defined by the foldcore (24, 24.1, 24.2). The foldcore (24, 24.1, 24.2) is comprised of a flat base material (42) folded at first meandering fold lines (60) and second meandering fold lines (62), the first fold lines crossing the second fold lines, so that the channels (34, 36) have a zigzag form.

An aircraft wing including a skin and a framework for supporting the skin. The framework includes a spar extending along a length of the wing, the length extending between a root end and a tip end of the wing, a plurality of ribs extending chordwise between a leading edge of the wing and a trailing edge of the wing; and a plurality of stringers attached to the skin. The plurality of stringers includes a first stringer extending along the length of the wing, wherein a proximal end of the first stringer is arranged nearer the root end and a distal end of the stringer is arranged nearer the tip end; and an auxiliary stringer extending non-parallel with the first stringer and arranged such that a portion of the auxiliary stringer is proximal to the distal end of the first stringer.

The name of the present invention is a conductive composite material and a method for preparing a conductive composite material. The conductive composite material includes a first elastic polymer layer, a conductive fluorinated fluid layer on the first elastic polymer layer, and a second elastic polymer layer on the conductive fluorinated fluid layer.

Stamp-forming a composite assembly includes applying a low-melt film on an interface surface of at least one spar, loading the spar onto at least one tooling segment such that the interface surface is exposed, and heating the tooling segment to a first temperature that is above a crystallization temperature of the low-melt film and below a melt temperature of a resin of the spar. A second low-melt film may be applied on a fiber-reinforced thermoplastic skin blank, and the skin blank is heated to a second temperature above a melt temperature of a resin of the skin blank. The skin blank is then draped around the tooling segment such that at least some of any second low-melt film contacts the low-melt film on the interface surface of the spar. The stamp press is then closed to co-consolidate the skin blank and the spar, thereby stamp-forming the composite assembly.