48results about "Sandwich constructions" patented technology

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.

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 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.

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.

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.

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 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).

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.

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.

The invention provides a modular tool for manufacturing a multi-spar torsion box having different web heights, the modular tool comprising: a core module having a hollow beam geometry comprising a first base and a second base opposite the first base and two walls extending between the first and second bases; at least one spacer module configured to couple with the core module, wherein the web height of a multi-spar torsion box is defined by the coupling between the core module and at least one spacer module. The invention also provides a method for assembling a tool set for manufacturing a multi-spar torsion box and an aircraft comprising a multi-spar torsion box manufactured by the method.

A Method (100) for manufacturing a stiffened aerodynamic structure component (10) is provided, comprising the following steps: - Providing (102) an aerodynamic structure component (10), wherein the aerodynamic structure component (10) includes a plurality of spar elements (12) extending spanwise within the aerodynamic structure component (10); wherein the plurality of spar elements (12) is integrally formed with an upper skin section (14) and a lower skin section (16), thereby providing an integrally formed aerodynamic structure component (10), - Generating (104) at least one chordwise reinforcement element (18) by injecting a curable medium (32), preferably via at least one injection hole (20), in the upper skin section (14) and / or the lower skin section (16).

A composite part for a vehicle is provided. The composite part has a first surface (22a) and an opposing second surface (22b), wherein at least part of the first surface forms an outer aerodynamic surface of a vehicle. A region of the part has a curvature in two orthogonal directions. The part has an undulation (40) defining a channel (60) in one of the first surface or the second surface, and a corresponding ridge (50) in the other of the first surface and the second surface.

A sheet member 1 includes: a front surface 111 including a first area 130 and a second area 140; a back surface 122; and a through-hole 15 formed between the front surface and the back surface, the first area 130 is a surface of a convex structure 13 that is formed as at least a part of the riblet structure RB and that protrudes from the second area, and an opening 151 of the through-hole 15 is formed in the second area 140.

An aircraft with carbon fiber material comprising a blended wing body having a main body and wings with no clear demarcation between the wings and the main body along a leading edge of the aircraft, wherein the blended wing body comprises, at least an outer skin surface comprising a carbon fiber material, at least a structural element comprising the carbon fiber material, and at least a portion of the main body comprising the carbon fiber material and configured to contain a pressure vessel and at least a propulsor, the at least a propulsor attached to the blended wing body and configured to propel the aircraft.

An aircraft wing structure comprises first and second unitary shell structures 7, 8, preferably formed from a fibre-reinforced composite material. Each shell structure 7, 8 comprises a partial front spar, a partial rear spar and a wing skin. The partial front spars of each shell structure and the partial rear spars of each shell structure are arranged to overlap to form front and rear spars respectively. A wing box structure 6 constructed according to the invention is lightweight, strong and durable, with a lower overall part count than has been conventionally achievable. Aircraft systems equipment may be attached to the shell structures 7, 8 prior to, or after, assembly into a wing box.

The present invention relates to a core material that is air-permeable and compatibly satisfies both weight reduction and mechanical properties, and to a skin-core structural body using the same. The core material according to the present invention has a structure including a collective body of hollow columnar cells compartmentally formed by a cell wall, in which the cell wall includes a fiber-reinforced composite material including a fiber (Cf) and a resin (Cr), and at least a part of it is a porous structural portion composed of an aggregate including the resin (Cr) and the fiber (Cf).

To provide a composite structure and a method for manufacturing the composite structure.SOLUTION: A composite structure has a second layer 120 connected to a first layer 110 to form a layered structure. The second layer has a plurality of base portions 140 abutting the first layer and a plurality of beaded portions 130 protruding from the plurality of base portions. Each beaded portion of the plurality of beaded portions defines a channel 135 between the first layer and the second layer. A method for manufacturing the composite structure includes depositing composite materials over a tool to form the second layer. The second layer has the plurality of beaded portions and the 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 the composite materials over the second layer and the plurality of mandrels to form the first layer. The method includes bonding the first layer to the second layer.SELECTED DRAWING: Figure 1

Methods (100, 200) are disclosed for the manufacture of composite components having incorporated reinforcing structures machined into composite material substrates, and composite components manufactured according to disclosed methods, and assemblies and larger structures comprising the composite material components.

Systems and methods are provided for fabricating multiple segment spars for an aircraft. In one example the method includes fabricating preforms of fiber reinforced material for spar segments, hardening the preforms to form the spar segments, and bonding the spar segments together to form a completed spar detail. In addition to bonding, other examples include co-curing and fastening the spar segments. In additional examples, the spar segments include kinks or sub-kinks as described.

The invention discloses a tilting wing multi-rotor aircraft system based on self-adaptive pneumatic optimization, and relates to the technical field of aircrafts. A plurality of protruding wing roots are arranged on the outer wall of the aircraft body, tilting shafts are installed in the protruding wing roots through bearings, tilting rotor wings are installed at the other ends of the tilting shafts, and telescopic mechanisms are installed in the protruding wing roots. Through cooperative work of the telescopic rear wings, the magnetic coding sliding blocks, the sliding rails, the tilting rotor wings and other components, precise sliding and position feedback of the telescopic rear wings in the protruding wing roots are achieved, the telescopic rear wings meet the requirements for light weight and high strength, and in the flight mode conversion process, the telescopic rear wings can be fixed to the rear wings. The aerodynamic configuration of the wings is optimized through dynamic adjustment of the telescopic rear wings, energy loss is reduced, and autonomous switching between the vertical take-off and landing mode and the efficient level flight mode of the aircraft is achieved through the tilting rotor wings.

An aircraft structural component, for example, a wing spar that has an I-beam shape having upper and lower spar caps coupled by a web therebetween and that provides structural support for an aircraft wing. The wing spar may be fabricated as a graphite composite that is thermally cured to have a certain stiffness. In one embodiment, the wing spar is fabricated so that the upper spar cap has a higher CTE than the web, which creates tension forces in the spar cap when the spar is thermally cured and then cooled. Therefore, when the wing spar is mounted to the wing and the aircraft is in flight, compression forces on the wing skin act to relieve the tension forces in the spar cap, which reduces the compression buckling load on the wing.

A composite part for a vehicle is provided. The composite part has a first surface and an opposing second surface, wherein at least part of the first surface forms an outer aerodynamic surface of a vehicle. A region of the composite part has a curvature in two orthogonal directions. The composite part has an undulation defining a channel in one of the first surface or the second surface, and a corresponding ridge in the other of the first surface and the second surface.

The present invention provides an aerodynamic structure (1), in particular for an aircraft (10), comprising a leading edge (2) having an outer structure, wherein the outer structure comprises: a first layer (3); and a second layer (4) geometrically corresponding to the first layer (3); wherein the first layer (3) is positioned on an outside (5) of the outer structure, the outside (5) being exposed to a flowing fluid, and the second layer (4) is positioned on an inside of the outer structure; and wherein the second layer (4) has a higher ductility than the first layer (3). Further the present invention provides an aircraft comprising such an aerodynamic structure.

Co-curable and co-cured UV / visible light-resistant composite materials are disclosed with UV / visible light protection exclusively imparted to a co-curable composite material substrate assembly by a co-curable UV / visible light-resistant layer disposed to cover the co-curable and co-cured composite material substrate.