101results about "Spars/stringers" patented technology

Truss-braced wing aircraft engine mounting and associated systems are disclosed. An example aircraft includes a fuselage including a central structural section, a wing extending from the fuselage, a truss extending from the central structural section and coupled to the wing, a pylon coupled to the central structural section, and an engine coupled to the pylon, wherein the pylon is positioned at an angle substantially 45 degrees from a horizontal plane and a vertical plane.

The invention discloses a built-in main control actuator connecting structure of an aircraft wing box, and relates to the technical field of aircraft structure design, an actuator is located in the aircraft wing box structure, an actuating rod of the actuator penetrates through an actuator rod avoiding hole of the aircraft wing box structure, and the actuating rod stretches out and draws back to cooperate with the actuator to rotate up and down in the wing box to achieve deflection of a control surface. The invention is suitable for small airfoils and small and medium-sized airfoils with low airfoil height, solves the problems that the small airfoils of the small airfoils are small, the trailing edge space is limited, and a control surface actuating system cannot be mounted, and avoids the problem that the aerodynamic efficiency is influenced as an actuating structure is arranged outside the airfoils; the structure of the actuator connector is designed and arranged on the front main bearing beam of the wing box, the load transfer continuity of the wall plate structure is not broken, meanwhile, the actuating load can be directly diffused to the skin through the support connector, and the load transfer efficiency is higher. By separating a daily inspection and maintenance passage from a disassembly and assembly passage, the opening size of the airfoil wall plate is reduced, the load transfer capacity of the lower airfoil skin is guaranteed, and structural weight reduction is facilitated.

Apparatus and methods for ultrasonic inspection of elongated composite members in a single scan pass using pulse echo phased arrays (97, 98, 99) operating in a bubbler method. The system concept is fully automated by integrating an inspection NDI probe assembly (32, 34, 36) to a robot (100) and using the robot to move the inspection NDI probe assembly along the part (i.e., outside of an inspection tank); and by integrating tooling fixtures that move out of the way as the inspection NDI probe assembly travels along the length of the part during the inspection. In addition, the system allows for generally elongated composite members having lengthwise variation in shape, curvature and dimensions.

A composite laminated layup panel structure suitable for wing combined connection, comprising a main box section skin (100) and a plurality of stiffeners (200) which are integrally cured and formed. The main box section skin is sequentially provided with recesses (300) at a front edge and a rear edge, forming a front recessed edge (110) for overlapping with a leading edge skin (400) and a rear recessed edge (120) for overlapping with a trailing edge skin. A front spar stiffener (600) configured to be connected to a front spar web (500) is fixedly arranged at the bottom of a transition region between the front recessed edge (110) and a main box section skin main body. A rear spar stiffener (700) configured to be connected to a rear spar web is fixedly provided at the bottom of a transition region between the rear recessed edge (120) and the main box section skin main body. The panel structure simplifies connection forms of panels and related structures, reducing the number of parts, improving assembly efficiency, and also facilitating disassembly and replacement maintenance during aircraft use and maintenance.

The application provides a swept-back three-beam thin wing composite wing structure, which comprises a wing main body and an aileron installed in a wing main body mounting groove, the wing main body comprises a wing skeleton and a wing outer skin structure outside the wing skeleton; the wing skeleton comprises a longitudinal member unit and a transverse member unit, the longitudinal member unit comprises a front beam, a middle beam and a rear beam, the front beam and the middle beam and the middle beam and the rear beam are arranged at acute angles, the transverse member unit comprises a plurality of wing ribs arranged between the front beam and the middle beam, between the middle beam and the rear beam and on one side of the front beam towards the front of the fuselage; the front beam, the middle beam and the rear beam are all provided with a spar connecting head for connecting with the fuselage. The wing provided by the application adopts a three-beam structure, each wing beam has a mouth-shaped cross section, so that each wing beam forms a multi-wall form, can bear the bending moment and the torque brought by a large wing effective load, and the wing structure is light in weight, high in strength and good in integrity.

A method for manufacturing a structural element (2) of a composite material is described, the structural element extending in a straight or curved longitudinal direction (D), having a Z-shaped cross section and comprising a central web (2a) and two flanges (2b, 2c) in its final configuration, the two flanges extending from opposite ends of the web (2a) in their respective opposite directions at a given final angle; the method comprises the steps of: arranging a plurality of layers (3) of the composite material on a forming portion (4) of a molding die; laminating the layers (3) onto the forming portion (4) such that the web (2a) is at least partially arranged in its final configuration, such that a first flange (2b) is positioned relative to the web in the final configuration of the web. (2a) is arranged at a final angle to form the final configuration of the first flange, and the second flange (2c) is arranged at an initial angle different from and greater than the final angle relative to the portion of the web (2a) arranged as the final configuration of the web, to form the initial configuration of the second flange; the movable part (6) of the molding die is moved from a rest position to a bent position, the movable part being movable relative to the fixed part (5) of the molding die; the second flange (2c) is shifted from the initial configuration to the final configuration, in which the second flange is at the final angle relative to the portion of the web (2a) arranged as the final configuration of the web; the shifting step (d) is performed by the shifting step (c).

A wing for an aircraft include battery modules a plurality of ribs that are spaced apart from one another in a longitudinal direction of the wing, the plurality of ribs respectively defining a plurality of first through-holes through which the battery modules pass, a plurality of first flange portions that are respectively disposed at the plurality of ribs along edges of the first through-holes, the plurality of first flange portions defining battery seating surfaces that support the battery modules, and a spar that supports the plurality of ribs.

This invention relates to an integrated variable camber wing based on a piezoelectric-driven torsional cubic superstructure, comprising a wing skeleton and a skin. The wing skeleton includes multiple torsional cubic superstructures arrayed along the wing chord direction. Each torsional cubic superstructure is a cube formed by an array of multiple unit-cell torsional superstructures. Adjacent unit-cell torsional superstructures in the wing length and thickness directions share a straight rod, while adjacent unit-cell torsional superstructures in the wing width direction share four Z-shaped piezoelectric rods. Each unit-cell torsional superstructure is a cube formed by an array of four cell units, with adjacent cell units sharing a straight rod. Each cell unit includes Z-shaped piezoelectric rods and straight rods. Two Z-shaped piezoelectric rods are connected to the two ends of two straight rods. Each Z-shaped piezoelectric rod at both ends of the torsional cubic superstructure is connected to the skin. When all the Z-shaped piezoelectric rods of the unit-cell torsional superstructure undergo compressive deformation in the height direction, the unit-cell torsional superstructure undergoes torsional deformation, thus causing the torsional cubic superstructure to undergo torsional deformation. By converting the compressive deformation of the cellular unit into the torsional deformation of the compressive-torsional cubic superstructure, the wing camber is changed, and the drive components are eliminated, thus realizing the lightweight design of the morphing wing.

The present invention relates to a split rib for securing a leading edge assembly. The rib includes a first rib portion having a first rear mounting portion for attaching the first rib portion to a wing box assembly and further having a front mounting portion opposite the first rear mounting portion. The rib further includes a second rib portion having a second rear mounting portion for attaching the second rib portion to the front mounting portion of the first rib portion.

The invention belongs to the technical field of aircraft structure design, and particularly relates to a winglet structure. A winglet upper wall plate is installed on the upper side of a wing box section butt joint rib, a winglet upper wing surface covering cap is arranged at the position, close to a wing box section, of the winglet upper wall plate, a winglet lower wall plate is installed on the lower side of the wing box section butt joint rib, and a winglet front beam and a wing box section front edge partition plate are installed on the front side of the wing box section butt joint rib. The winglet rear beam and the wing box section rear edge partition plate are mounted on the rear side of the wing box section butt joint rib; the winglet front edge partition plate is connected with the winglet front beam, the front edge skin is connected with the winglet front edge partition plate, the winglet front beam and the wing box section front edge partition plate, and a front edge covering cap is arranged at the position, close to the wing box section, of the front edge skin. The winglet rear edge comprises a winglet rear edge partition plate, and the winglet rear edge partition plate is connected with the winglet rear beam; the winglet tip is located at the position away from the wing box section and connected with the winglet upper wall plate, the winglet lower wall plate and the leading edge skin.

Systems, devices, and methods including one or more rib mounting flanges, where each rib mounting flange comprises: a spar opening configured to receive a main spar of a wing panel; one or more holes for receiving cross-bracing cables; and one or more holes for receiving cross-bracing cables; and one or more holes for connecting the rib mounting flange to an adjacent rib mounting flange.

A method of stiffening the integral ribs of a panel to prevent crippling of the ribs during forming of the panel. The method provides that anti-crippling stiffeners are attached to the ribs, such as by using clamps. Then, the panel is formed while the anti-crippling stiffeners are attached to the ribs. The anti-crippling stiffeners add stiffness to the ribs and prevent the ribs from crippling during forming of the panel. The anti-crippling stiffeners could be C-shaped and be attached to the panel using C-shaped clamps. Both the panel and the anti-crippling stiffeners could be formed of aluminum. Another embodiment involves clamping at least one anti-crippling strap to the flange of a wing stringer to prevent the flange from crippling during forming of the wing stringer.

The invention relates to an aircraft assembly (1) comprising an upper fitting (104c) fixed to a wing structure (104a), a lower fitting (104d) fixed to a wing underside panel, an engine pylon (106) with two side panels hinged to the lower fitting (104d), a starboard connecting rod (252a) and a port connecting rod (252b) hinged between the side panel (208a-b) on the same side and the upper fitting (104c), and a mounting block having a proximal end integral with a rear rib of the engine pylon (106) and a distal end mounted in a window of the lower fitting (104d) via an annular linear connection. With such an assembly, the occurrence of out-of-plane forces is limited.

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.

A method of forming a structural component is provided. The method includes forming, by arranging at least one composite material to a first determined configuration, a first uncured composite laminate beam structure. The method also includes forming, by arranging at least one composite material to a second determined configuration, a second uncured composite laminate beam structure. The method further includes co-curing a first surface of the first uncured composite laminate beam structure to a first surface of the second uncured composite laminate beam structure to form a composite laminate beam structure, where the composite laminate beam structure is a single continuous structure. The method additionally includes coupling a surface of the composite laminate beam structure to an internal surface of a skin.

Profiled stiffening element for load-bearing bending stiffening on a planar fiber composite structure, wherein the stiffening element is formed from a fiber composite material comprising a fiber material and a matrix material embedding the fiber material, has a profiled cross-sectional shape and extends in a longitudinal direction, characterized in that the stiffening element has an electrically conductive metallic layer extending continuously in the longitudinal direction, wherein the stiffening element has at least two electrical contacts electrically connected to the electrically conductive layer.

The present disclosure relates to methods, systems, and non-transitory computer readable media for inspecting a wing panel. Systems and methods for inspecting a wing panel are provided. Some methods include the steps of advancing a wing panel through a non-destructive inspection (NDI) station and inspecting the wing panel at the NDI station with an inspection head. During inspection and / or advancement, the wing panel can be suspended beneath a strongback, such as by using a vacuum coupler and / or an adjustable length spring member of the strongback, and the step of suspending can include profiling the wing panel using the vacuum coupler and / or the spring member. Other methods include the steps of receiving a wing panel at an NDI station and inspecting a portion of the wing panel during movement through the NDI station. Some systems include a track, a strongback that suspends a wing panel beneath and advances the wing panel along the track, and an NDI station disposed at the track to inspect the wing panel while suspended.

Thermal protection systems for aerospace vehicles, such as hypersonic vehicles or space vehicles, are disclosed herein. In some embodiments, the disclosed thermal protection systems can be used to protect propellant tanks (e.g., composite propellant tanks) and vehicle primary structures (e.g., composite primary structures) from exceeding upper design temperature limits during vehicle ascent, reentry into Earth's atmosphere, and / or high velocity flight through an atmosphere. In some embodiments, a thermal protection system can include a framework of elongated composite members forming a plurality of cells. One or more layers of insulation material such as aerogel can be positioned in each of the cells. The cells can be covered by an external skin of the vehicle. In some embodiments, the framework is attached to a propellant tank and the insulation material is in contact with the external surfaced of the propellant tank.

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).

Aircraft comprising at least one flexible profile of composite material forming at least one element among a coupling shaft of a moving surface and a moving surface. The invention relates to an aircraft comprising at least one moving surface (20) connected to a fixed structure (22) by at least one articulation system (24) which has a pivot axis (A24), at least one actuation system configured to rotate the moving surface (20) around the pivot axis (A24) and at least one profile (32) of composite material comprising reinforcing fibers embedded in a resin matrix, at least 50% of the reinforcing fibers of the profile (32) forming with a longitudinal direction an angle between [-60° and -45°] and between [+45° and +60°], less than 40% of the reinforcing fibers of the profile (32) forming with the longitudinal direction an angle between [-30° and +30°].This profile (32) can form the moving surface (20) and / or a coupling shaft (28) connecting the moving surface (20) and the fixed structure (22) or the moving surface (20) and the actuation system. This arrangement of reinforcing fibers gives the profile very high torsional strength while allowing it to deform in bending. Figure 5.

Example aircraft and aircraft evacuation slides and related methods are disclosed herein. An example evacuation slide assembly for an aircraft disclosed herein includes an inflatable slide body attachable to the aircraft and configured to inflate in response to an evacuation event, and a readiness indicator gate supported by the inflatable slide body, wherein the readiness indicator gate is movable from an unready position during inflation of the inflatable slide body, thereby indicating the inflatable slide body is not ready for use, to a ready position thereby indicating readiness of the inflatable slide body for user egress from the aircraft onto the inflatable slide body.

Embodiments are directed to systems and methods for two or more cured composite assemblies that are bonded together to form a tubular composite structure, wherein each of the cured composite assemblies do not have a tubular shape. The tubular composite structure may form a spar for an aerodynamic component, for example. The two or more cured composite assemblies may comprise carbon or fiberglass composite materials or a combination of materials. Each of the cured composite assemblies may further comprise axial edges that are configured to be bonded to another of the cured composite assemblies, wherein the axial edges have a sloped shape. An adhesive agent may be applied on the axial edges for bonding two cured composite assemblies. Alternatively, or additionally, one or more fasteners may be used to attach the axial edges of at least two cured composite assemblies.

Method for manufacturing a Y-shaped stringer (100) made of composite material, the Y-shaped stringer (100) comprising a stringer web (110) having a cross-section in a I-shape, lower flanges (130a, 130b), a first opened triangular-shaped cross-section structure (120) comprising first and second lower vertices (120a, 120b) respectively joined to the lower flanges (130a, 130b) and an upper vertex (120c) connected to a first end of the stringer web (110), wherein the first opened triangular-shaped cross-section structure (120) and the stringer web (110) form a cross-section in a Y-shape.

An aircraft wing structure (100) comprising a spar (101) and a systems module (200) is disclosed. The systems module (200) can comprise electrical cables (205). The systems module (200) is releasably secured to a plurality of brackets (300) on a first side of the spar (101) and the second, opposite side of the spar is configured to define a wall of a fuel tank of an aircraft wing. Methods of installing and removing the systems module are also described.

The invention discloses a composite material wing inclined supporting rod and a forming method thereof, the inclined supporting rod comprises the following structures: a structure main body, the cross section of the structure main body is of an I-shaped structure, and the two opposite sides of the structure main body are respectively provided with clamping grooves which extend along the longitudinal direction and have opposite openings; the transverse section of the front edge plate is of a C-shaped structure, the two edges of the front edge plate are provided with sunken first clamping plates in the longitudinal direction, and the first clamping plates are clamped in the clamping grooves and fixed; the transverse section of the rear edge plate is of a C-shaped structure, the two edges of the rear edge plate are provided with sunken second clamping plates in the longitudinal direction, and the second clamping plates are clamped in the other clamping grooves and fixed; connecting holes are formed in the metal joints, and the metal joints are connected to the two ends of the structure body and used for being connected with the wings and the fuselage. The inclined supporting rod has the advantages of being light in weight, high in strength, resistant to fatigue and resistant to corrosion, the arrangement of the front edge plate and the rear edge plate can facilitate reduction of air resistance when an aircraft flies, and the fuel economy of the aircraft can be improved.

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 method for use in assembling an aerofoil structure. A torsion box has a first mounting feature and a second mounting feature. A fixed leading or trailing edge structure has a third mounting feature and a fourth mounting feature. The first mounting feature and the third mounting feature each have a fastener hole prior to aligning the leading or trailing edge structure to the torsion box. The second mounting feature and the fourth mounting feature are without fastener holes for joining the torsion box to the fixed leading or trailing edge structure prior to aligning the leading or trailing edge structure to the torsion box. The method includes aligning the fixed leading or trailing edge structure to the torsion box so as to align the fastener holes of the first mounting feature and the third mounting feature, drilling through the second mounting feature and the fourth mounting feature to form fastener holes in the second mounting feature and the fourth mounting feature when the fixed leading or trailing edge structure is aligned with the torsion box, and installing a fastener through the fastener holes of the first mounting feature and the third mounting feature, and installing a fastener through the fastener holes of the second mounting feature and the fourth mounting feature, so as to join the fixed leading or trailing edge structure to the torsion box.