605results about "Fuselage bulkheads" patented technology

The invention discloses a multifunctional aircraft skin bionic heterostructure and a preparation method thereof, belongs to the technical field of aerospace, and aims to solve the problem that a traditional aircraft skin structure is difficult to synergistically realize impact resistance, sound absorption, noise reduction and electromagnetic shielding, and a multifunctional composite skin protection structure is designed in a breakthrough manner. Through collaborative innovation of a bionic framework and a functional material, triple optimization of mechanical-acoustic-electromagnetic characteristics is realized. A bionic skeleton structure with outstanding mechanical properties is obtained through a layered annular structure of a glass sponge skeleton and a hierarchical structure design of a thighbone, and the bionic skeleton structure is combined with modified polyurethane foam with excellent sound absorption performance and electromagnetic shielding performance to obtain a multifunctional heterostructure with impact resistance, noise reduction and electromagnetic shielding.

Methods and apparatus for determinant assembly support are disclosed. A disclosed example apparatus to support a panel component for determinant assembly thereof includes a support frame that is releasably couplable to a jig, a first clamp to support the panel component at a first position of the panel component, a first body of the first clamp or the support frame including a nominal clearance opening to receive a first fastener to operatively couple the first clamp to the support frame, and a second clamp to support the panel component at a second position of the panel component different from the first position, a second body of the second clamp or the support frame including a slotted opening to receive a second fastener to operatively couple the second clamp to the support frame, the second fastener movable about the slotted opening to enable the second clamp to reduce forces imparted to the panel component while held by the support frame.

A hydrogen aircraft includes: an airframe including a fuselage and a wing; at least one propulsion system fixed to the airframe; a pressurized chamber disposed inside the fuselage; a pressure bulkhead that is disposed at a rear part of the pressurized chamber and has strength to withstand pressurization of the pressurized chamber; a hydrogen tank that is disposed in an accommodation section and stores liquid hydrogen, the accommodation section being installed behind the pressure bulkhead, outside the pressurized chamber, and inside the fuselage; and a supply line that supplies the liquid hydrogen stored in the hydrogen tank to the propulsion system.

A partition including, but not limited to, a first body comprising a planar, foldable material. The partition further comprises a plurality of flat substrates resistant to folding. The partition further comprising an adhesive. The substrates are adhered to the first body by the adhesive and arranged in a repeating pattern that defines a plurality of columns and rows. Each column and each row is spaced apart from its neighbor, thus defining longitudinal gaps between the columns and lateral gaps between the rows. The first body folds laterally along the longitudinal gaps in an accordion-like manner and contemporaneously folds longitudinally along the lateral gaps in an accordion-like manner. The partition contracts both laterally and longitudinally when first body contemporaneously folds laterally and longitudinally. The partition extends both laterally and longitudinally when the first body contemporaneously unfolds laterally and longitudinally.

The invention belongs to the technical field of grating stiffened wall plates, and particularly relates to a grating stiffened wall plate for an aircraft structure and a forming mold and method.The grating stiffened wall plate comprises a grating frame, a skin plate is arranged on the grating frame, and the skin plate comprises an outer skin arranged on the bottom face of the grating frame and an inner skin arranged on the grating frame; the grille frame comprises a transverse rib core material and a longitudinal rib core material which are connected with each other, the inner skin is located between the transverse rib core material and the longitudinal rib core material, the connecting position between the transverse rib core material and the longitudinal rib core material is called as a connecting node, and each connecting node is provided with a forming block; the strength and the stability of the grating frame are effectively enhanced through the forming blocks. In addition, the transverse rib core materials and the longitudinal rib core materials are tightly fixed through flanging wrapping of the inner skin, and therefore the strength and the structural efficiency of the grating stiffened wall plate are remarkably improved. And the unique requirements of aircraft wall plate structures such as airplanes and the like on the aspects of transmitting and bearing large loads are met.

Aerogels comprising a hydrophobic polyimide moiety, including hydrophobic polyimide aerogels, as well as methods of manufacture and applications thereof, are generally described.

An aerospace panel includes a first skin, a second skin spaced apart from the first skin, and a first truss structure connecting the first skin to the second skin. The first truss structure includes a plurality of truss members. Each truss member is integral with the first skin and the second skin, such that the first skin, the second skin, and the first truss structure collectively form a single monolithic joint-free structure. At least one of the skins also includes at least one lattice region that includes a lattice grid and an array of openings.

The present disclosure provides a battery load support structure of a fuselage, the battery load support structure including a battery mount support formed in a longitudinal direction of a floor frame of the fuselage, a battery unit inputted into an opened lower side of the floor frame and fastened to the battery mount support, and a joint unit configured to support the battery mount support and connected to a plurality of framework members that constitutes the floor frame.

The invention belongs to the field of composite material wallboard design, and provides a stiffened wallboard which comprises a first skin, a second skin and a connecting piece, a first stringer is arranged at the top of the first skin, a second stringer is arranged at the top of the second skin, the first skin is connected with the bottom of the second skin through the first stringer, and the second skin is connected with the connecting piece through the second stringer. The first stringers and the second stringers are distributed in a staggered mode, the connecting pieces are connected with the first skin through the first transverse connecting band plates or connected with the second skin through the second transverse connecting band plates, and connection of the first skin or the second skin is achieved. According to the scheme, the two connected wall plate layers are designed, so that larger out-of-plane loads can be borne in the thickness direction, and the structural stability is improved; and the stringers are distributed in a staggered manner, so that bearing force can be more uniformly distributed in a limited space, the stress caused by out-of-plane load is uniformly transmitted to the two wall plate layers, and the structural stability is further improved.

An airframe part for an airframe for an unmanned aerial vehicle is disclosed. The UAV may be a tail-sitting fixed-wing vertical takeoff and landing UAV. The airframe part has an expanded polystyrene liner and a polycarbonate shell. The shell is arranged on the outer side of the airframe part. The outer face of the liner is mainly or completely covered by the shell. Adhesive bonding is established between the shell and the liner over essentially their entire contact area. The same mold can be used for forming the liner and for bonding the liner and the shell.

An aircraft structure part having a stacked arrangement of layers with the layers being one of an at least first doubler and a basic layer of the outer skin of an aircraft fuselage, with the first doubler and the basic layer being bonded by a first intermediate bond film layer to form a step-shaped configuration of the aircraft structure part, wherein the first doubler includes a first fiber prepreg layer, in particular glass fiber prepreg layer, arranged at the perimeter of the first doubler and wherein the first fiber prepreg layer is bonded to the basic layer by at least one first bond-film segment deposited adjacent to the first fiber prepreg layer in a first wedge-shaped element at least partially incorporating the first fiber prepreg layer. A method is disclosed for providing an aircraft structure part and an aircraft fuselage provided with the aircraft structure part.

The utility model belongs to the technical field of aviation structure design, and relates to a fuselage external protruding component reinforcing structure which is characterized in that two combined frames and two combined longitudinal pieces form a box section combined piece; and the box section assembly is fixedly connected with the fuselage skin. The reinforcing structure is used for the fuselage external protruding component fuselage, and overcomes the defects that a traditional supporting structure is poor in bearing capacity, low in rigidity, large in weight, poor in manufacturability and the like.

An aircraft structure including an outer mold line skin, one or more frames located on an inner surface of the outer mold line skin having an outer flange attached to an inner surface of the outer mold line skin, a first core filler in contact with the outer flange, an inner flange in contact with the first core filler and a web skin, wherein the web skin and the outer flange encapsulate the inner flange and at least a portion of the first core filler, and an arched inner mold line skin located between each frame of the one or more frames, wherein the arched inner mold line skin is in contact with each web skin of the one or more web skins.

A variable fuselage fuel tank structure of a high-speed aircraft comprises a fuselage fixing structure, an external flexible heat insulation skin and an internal skin bearing framework. The machine body fixing structure comprises a lower wall plate and an upper wall plate; the edges of the lower wall plate and the upper wall plate are connected, and the upper wall plate is provided with a fuel circulating hole; the outer flexible heat insulation skin is connected with the edge of the upper wall plate; the external flexible heat-insulating skin comprises a flexible fabric heat-proof outer layer and a microporous aerogel heat-insulating inner layer; the inner skin bearing framework is connected to the inner side of the outer flexible heat insulation skin and comprises reinforcing longitudinal beams, elastic elements and high-temperature-resistant rubber. The longitudinal beams are connected to the inner side of the outer flexible heat insulation skin in the longitudinal direction. A plurality of groups of elastic elements are respectively connected between two adjacent longitudinal beams, are distributed along the longitudinal direction, and are connected to the inner side of the external flexible heat insulation skin; and the high-temperature-resistant rubber is filled between the longitudinal beam and the elastic element, is filled in the elastic element and is adhered to the inner side of the external flexible heat-insulating skin.

The utility model discloses an eVTOL aircraft wing skin structure which comprises an upper skin, the lower portion of the upper skin is connected with a lower skin, and the front end and the rear end of the butt joint position of the upper skin, the lower skin and a motor arm appearance area are connected with a front splicing piece and a rear splicing piece respectively. According to the structure, the product quality is good, the appearance is complex, the wing appearance area is in lap joint, the motor arm appearance area is in butt joint forming at the combined part of the wing appearance and the motor arm appearance, the defects that the two areas are sunken and the combined part is prone to generating fiber bridging are avoided, and the product quality is improved.

The invention relates to a self-adaptive skin structure suitable for a cross-airspace speed domain morphing aircraft, and belongs to the technical field of intelligent skin. The skin structure is of a multi-layer form and sequentially comprises a heat insulation protection layer, a flexible skin base body, a sensing layer and a bearing and driving layer from the outer side to the inner side. Wherein the outermost side adopts a fiber sandwich structure formed by coating aerogel with an alumina fiber reinforced alumina composite material as a heat insulation protective layer; a high-temperature-resistant carbon fiber elastic matrix composite material is used as a flexible skin matrix; a bearing layer and a driving layer on the inner side are integrally designed, and a honeycomb cell element structure driven by an integrated high-temperature shape memory alloy wire is adopted; an optical fiber sensor is embedded between the flexible skin substrate and the bearing and driving layer, the optical fiber sensor realizes direct strain measurement, and the thickness change of the skin is calculated through a strain measurement value; the method can realize self-adaptive active flexible deformation of the skin when the cross-airspace speed domain aircraft is in a hypersonic flight stage, and has important application value in the field of skin design of the cross-airspace speed domain aircraft.

A rotary wing aircraft with a fuselage that forms an aircraft interior region, the fuselage comprising an upper primary skin that separates the aircraft interior region from an aircraft upper deck arranged above the fuselage, the upper primary skin being mounted on top of a predetermined number of supporting beams; wherein the aircraft upper deck comprises an engine accommodating region with a firewall arrangement, the engine accommodating region accommodating at least one aircraft engine within the firewall arrangement such that the firewall arrangement defines a fire proof separation at least between the at least one aircraft engine and the aircraft interior region; and wherein the firewall arrangement comprises at least a funnel-shaped lower firewall arranged between the at least one aircraft engine and the upper primary skin of the fuselage.

The invention discloses a laser processing microstructure electric heating anti-icing and de-icing skin and a using method thereof.The laser processing microstructure electric heating anti-icing and de-icing skin comprises a skin body arranged on the surface of an airplane, the skin body comprises an electric heating microstructure anti-icing and de-icing skin body, a microstructure anti-icing and de-icing skin body and a common skin body, and the input end of the electric heating microstructure anti-icing and de-icing skin body is connected with an airborne power source through a wire; the electric heating micro-structure anti-icing and deicing skin is arranged on the windward side, with the maximum airflow collection coefficient, of the outer surface of the aircraft, the micro-structure anti-icing and deicing skin is arranged on the upstream and downstream of the windward side of the outer surface of the aircraft, and the common skin is arranged on other parts, except the electric heating micro-structure anti-icing and deicing skin and the micro-structure anti-icing and deicing skin, of the outer surface of the aircraft. According to the laser processing microstructure electric heating anti-icing and deicing skin and the using method thereof, through matched arrangement of the electric heating microstructure anti-icing and deicing skin and the microstructure anti-icing and deicing skin, the purposes of being good in anti-icing and deicing effect, low in energy consumption, light in weight and good in maintenance performance can be achieved.

To provide a method of manufacturing structural components that is reliable, cost-effective, and solves the drawbacks associated with known manufacturing methods.SOLUTION: A structural component (1) having a plurality of subpanels (5) joined to each other, each including a subskin defined by a first layer of composite material, at least one stringer, and at least one rib fixed to an inner surface of the subskin and continuous subwalls defined by the bonded outer surfaces of the subskins of the subpanels (5) connected each other, a second layer (10) of composite material is laminated over the first layer to define a continuous external overskin layered over the subwalls, and a skin (4) is defined by a set of the subskin and overskin, the first layer and a second layer (10).SELECTED DRAWING: Figure 8

This document describes a system that can be attached to the outer skin of a vehicle, such as an aircraft. A disclosed version of the system comprises an antenna component, a support frame, brackets, and an aerodynamic fairing. The antenna component is connected to the support frame, which has a plurality of clevis mounting features. Each support fitting can be coupled between one of the clevis mounting features and the outer skin of the vehicle. The fairing consists of an upper and a lower part. The upper part has a defined opening. The lower part is configured and arranged to connect to the outer skin of the vehicle. The opening is shaped and dimensioned to accommodate at least a portion of the support frame that is connected to the upper part of the fairing.

Methods and apparatus for determinant assembly support are disclosed. A disclosed example apparatus to support a panel component for determinant assembly thereof includes a support frame that is releasably couplable to a jig, a first clamp to support the panel component at a first position of the panel component, a first body of the first clamp or the support frame including a nominal clearance opening to receive a first fastener to operatively couple the first clamp to the support frame, and a second clamp to support the panel component at a second position of the panel component different from the first position, a second body of the second clamp or the support frame including a slotted opening to receive a second fastener to operatively couple the second clamp to the support frame, the second fastener movable about the slotted opening to enable the second clamp to reduce forces imparted to the panel component while held by the support frame.

A panel assembly (such as for an internal cabin of an aircraft) includes at least one liquid flow path formed on at least one surface through surface treatment. The at least one liquid flow path is formed in relation to at least one bank on the at least one surface. The at least one liquid flow path is configured to divert liquid to a desired location.

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.

The invention relates to the technical field of layered products, and particularly discloses a heat-insulating and wave-absorbing skin material as well as a preparation method and application thereof. The heat-insulating and wave-absorbing skin material is of an integrated layered structure based on an aluminum oxide-mullite system, comprising a functional gradient layer with double gradients of fiber density and aperture, an impedance matching layer with a gradient dielectric constant, a resonance wave absorbing layer with a periodic dot matrix formed by silicon carbide and aluminum oxide core-shell microspheres, and a heat insulation layer with a nano-micron double-scale pore structure. According to the preparation method, a green body is constructed mainly through tape casting and silk-screen printing, and then co-sintering integrated forming is carried out. The material has the excellent characteristics of high temperature resistance, toughness, efficient heat insulation, broadband wave absorption and the like, and is suitable for extreme environments such as hypersonic aircraft skin and the like.

The invention discloses a stringer assembly, a stiffened wall plate, a stiffened wall plate assembling method and an aircraft, and belongs to the technical field of aircraft manufacturing. The stringer assembly comprises a cap-shaped stringer and a forming die, the cap-shaped stringer comprises a stringer main body and a lap joint part, the stringer main body extends in the first direction and is provided with a cavity, the forming die comprises a containing part extending in the first direction, the containing part is provided with a containing groove, and the lap joint part is provided with a lap joint part. The containing groove extends in the first direction and is opened in one side of the second direction, the containing groove and the opening of the cavity face the same direction, the stringer body penetrates through the opening of the containing groove and is installed in the containing groove, the end, in the first direction, of the containing groove is closed, and the lap joint part is in lap joint with the forming mold. The cap-shaped stringer is blocked by the closed end of the containing groove of the containing part, it is guaranteed that the cap-shaped stringer cannot penetrate through the containing part and pass through the net size area in the skin, the influence on the surface quality of the skin is reduced, and the forming die cost is reduced.

An aircraft including a blended wing body aircraft (BWB) having a main body and wings with no clear demarcation between the wings and the main body along a leading edge of the aircraft, a rear spar located within the main body, one or more gear wells located on a lower surface of the rear spar, a landing gear system having at least a main gear the at least a main gear situated in one of the one or more gear wells, and one or more propulsors located directly above the one or more gear wells, wherein at least one propulsor is located over the at least one gear well of the one or more gear wells.

The invention provides a composite fuselage wall plate for an aircraft, which comprises a skin wall plate, a plurality of annular reinforcing frames distributed on the inner surface of the skin wall plate along the heading direction of the aircraft, and a plurality of segmented stringers positioned between the adjacent annular reinforcing frames and distributed along the annular direction of the aircraft, an annular local reinforcing part with a preset thickness is formed at the intersection position of the segmented stringer and the annular reinforcing frame on the inner surface of the skin wallboard, and the end part, extending in the heading direction, of the annular local reinforcing part is overlapped with the end part, close to the annular local reinforcing part, of the segmented stringer.

An aircraft including first and second elements spaced apart from one another and at least one device for absorbing energy by compression, positioned between the first and second elements, including at least one module for absorbing energy by compression and at least one attachment system removably connecting each module for absorbing energy by compression to at least one of the first and second elements, at least one module for absorbing energy by compression including several first conduits oriented in a compression direction and being independent of the first and second elements.