76results 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.

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.

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.

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 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).

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.

A composite structure and methods of forming are presented. A composite structure comprises: structural supports and a structural skin formed by a plurality of composite panels fastened to the structural supports. Each composite panel of the plurality of composite panels comprises a first titanium end; a second titanium end; and a composite skin joined to and extending between the first titanium end and the second titanium end.

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 giant passenger aircraft structure includes a fuselage and at least two layers of wings located on both sides of the fuselage; a steering rudder is provided between adjacent upper and lower wings. This invention boasts excellent environmental friendliness and economic efficiency, specifically including: the use of a core panel structure, significantly reducing labor costs during production; a simple structural design, reducing maintenance difficulty and costs; an extremely low engine-to-weight ratio, resulting in significantly lower costs; an extremely low weight-to-fuel ratio, leading to substantial fuel savings; and a short wing design, facilitating wide airport applicability and increasing business opportunities.

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.

A core material has a structure including a collective body of hollow columnar cells compartmentally formed by a cell wall. The cell wall includes a fiber-reinforced composite material including a fiber (Cf) and a resin (Cr), and at least a part of the cell wall is a porous structural portion composed of an aggregate of the fiber (Cf) covered by the resin (Cr).

A method for manufacturing a box-shaped structure (1) for aircraft (4), wherein the structure (1) longitudinally extends between a first axial end and a second axial end and has a first longitudinal wall (5); a second longitudinal wall (6); a plurality of spars (7) longitudinally extending between said first wall (5) and second wall (6); a reinforcing rib (21) transversal to the spars and inserted, between the first and second wall, in the area of a transversal sector (S) interposed between the first axial end and the second axial end; the second wall (6) comprises a main portion (17) and a closing portion (18), which are coplanar to one another and define two distinct pieces with discontinuity and separated in the area of the transversal sector (S), so as to facilitate the insertion of the reinforcing rib without jeopardizing the structural properties of the box-shaped structure (1).

The utility model relates to the technical field of aviation, in particular to an aircraft wing with a deicing function, which comprises a wing component, a cam and a driving component, the wing assembly comprises a first wing, a second wing and a rotating shaft assembly, the rotating shaft assembly is connected with the first wing and the second wing, the driving assembly is connected with the second wing, and the output end of the driving assembly is connected with the cam; a cam groove is formed in the first wing, and the cam is arranged in the cam groove; when the driving assembly drives the cam to rotate in the cam groove, the first wing moves around the rotating shaft assembly, so that an ice layer on the wing assembly is removed. The unmanned aerial vehicle has the advantages that the cam is driven by the driving assembly, so that the first wing generates fluctuating vibration, and shearing force and vibration force generated by the vibration can quickly break an ice layer attached to the wing assembly.