126results about "Sandwich constructions" patented technology

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.

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.

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.

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.

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.

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.

Disclosed is an aircraft wing assembly 500, comprising: a stringer 300, the stringer comprising a first flange 320, a second flange 330, and a web (310 fig.4A) extending between, and connecting, the f

A leading-edge arrangement (2) for a flow body has a curved skin panel (4) having a skin-panel outer side and a skin-panel inner side (8), and a three-dimensional reinforcing lattice (16) having a lattice outer side. The curvature of the lattice outer side corresponds to the curvature of the skin-panel inner side (8), wherein the lattice outer side has a multiplicity of attachment points (32), which are connected to the skin-panel inner side (8), and wherein the reinforcing lattice (16) is constructed as a three-dimensional framework having a multiplicity of interconnected framework members (18), which form at least one layer of interconnected three-dimensional bodies.

The invention relates to a monolithic structure in composite material manufactured starting from a fiber-reinforced prepreg material and comprises two walls facing each other and at least one interconnection element extending transversely between the walls, connected to them and delimiting with the latter respective elongated cavities; wherein the walls extend symmetrically at opposite sides of a direction; the interconnection element is a rib extending transversely to the aforesaid direction; at least one wall has a sandwich configuration and comprises two panels facing each other and at least one spar member, which extends transversely between the panels, is connected to them, delimits with the latter respective elongated cavities and extends transversally to the rib.

The application discloses a rear-swept wing distributed non-equal-diameter rotorcraft and a control method thereof, which comprises a fuselage, a wing, a tilt-rotor group, a landing gear, a horizontal tail and a vertical tail; the tilt-rotor group comprises four tilt-rotors which are connected to the leading edge of the wing through a support rod; the rotor disc area near the wing tip is larger than the rotor disc area near the wing root, so as to balance the flight efficiency at low speed and the flight speed at high speed, and the rear sweep of the wing is beneficial to improving the stability of the tilt-rotor aircraft in the fixed-wing mode flight, and the environmental adaptability of the aircraft is improved. The application balances the efficiency of the tilt-rotor aircraft in the hovering and cruising states, the aircraft structure is relatively simple, is beneficial to the attitude control of the aircraft and the flight control software design.

The application provides an electric-duct vertical-lift co-rotating-wing aircraft, which comprises a fuselage, a main wing arranged in the middle of the fuselage, and engines connected to the fuselage through engine flight attack angle electric adjusting mechanisms respectively arranged on the front and rear sides of the main wing; the engine flight attack angle electric adjusting mechanisms are provided with ailerons; the front and rear engines are staggered vertically; and the front aileron, the main wing and the rear aileron are arranged in sequence from bottom to top. The electric-duct vertical-lift co-rotating-wing aircraft provided by the application is arranged at three heights, i.e. the lower, middle and upper heights, so that the engines and the wings are prevented from affecting each other when exerting their respective functions, and the stability of flight is improved.

The present invention provides a bladeless aircraft, comprising a fuselage (4), two front wings (3), two rear wings (5), two front propellers (2) and two rear propellers (1), wherein the fuselage (4) is located in the middle of the entire aircraft, the two front wings (3) are arranged at the front end of the aircraft and rotate together through a transmission shaft, the two rear wings (5) are arranged at the end of the aircraft and rotate together through a transmission shaft, the two front propellers (2) and the two front wings (3) are integrated, the two rear propellers (1) and the two rear wings (5) are integrated, and both the front propellers (2) and the rear propellers (1) are designed without exposed blades. The aircraft of the present invention adopts a bladeless design, which can prevent foreign objects from being sucked in, enhances safety during flight, and improves flight efficiency by utilizing an airflow multiplication effect.

The invention relates to a monolithic structure (1) made of composite material, manufactured starting from a fiber-reinforced prepreg material. The structure comprises two walls (6, 7) facing each other and at least one interconnecting element (8) extending transversely between the walls (6, 7) and connected to them, with which each elongated cavity (9) is defined. The walls (6, 7) extend symmetrically on both sides of a direction (B). The interconnecting element (8) is a rib (10) extending transversely to said direction (B), at least one wall (6, 7) having a sandwich structure and comprising two panels (11, 12) facing each other and at least one spar element (13) extending transversely between the panels (11, 12) and connected to them, with which each elongated cavity (14) is defined, extending transversely to the rib (10).

In some examples, a wing for an aircraft, the wing having a passive variable aeroelastic geometry for accommodating efficient flight at multiple aircraft speeds, comprises a wing skin comprising a composite structure comprising a lamination of multiple layers deposited for the wing skin according to a predetermined lamination profile, at least one of the layers comprising a continuous non-linear deposition of material, the wing skin comprising a stiffness distribution profile defined by the lamination profile.

A flying object having a shell (11) that separates an interior (12) of the flying object from an external environment, wherein the shell (11) of the flying object has, at least in sections, an inner wall (17) pointing toward the interior (12) and an outer wall (18) pointing toward the external environment to form at least one double-walled shell section such that a cavity is formed between the inner wall (17) and the outer wall (18), in which cavity at least one battery (20) with a plurality of battery cells for supplying electrical energy to the flying object is arranged, characterized in that the flying object has a fuselage on which lift-generating wings (10) are arranged, wherein the lift-generating wings (10) have a leading edge that has the double-walled shell section with at least one battery (20), wherein the leading edge is detachably arranged on the wing (10).

A method of manufacturing an aircraft aerofoil such as a winglet. The aircraft aerofoil comprises: a first cover, a second cover, an aerofoil leading edge, and an aerofoil trailing edge where the first and second cover are joined at a trailing edge interface. The first cover is formed by: a lay-up step in which a plurality of fibre plies are laid up to obtain a preform, wherein the preform has a thickness and a preform trailing edge, and the preform is laid up during the lay-up step with a ramp of decreasing thickness where the thickness of the preform decreases towards the preform trailing edge; and a curing step in which the preform is cured to form the first cover, wherein the ramp of decreasing thickness cures during the curing step to form a contact surface of the first cover. The trailing edge interface is assembled by engaging the contact surface of the first cover with a contact surface of the second cover.