477results about "Fuselage frames" patented technology

The utility model belongs to the technical field of flexible structure design, and particularly relates to a perimeter-variable fuselage frame structure which comprises sliding groove sections, sliding key sections, corrugated beams, height-variable supporting columns and length-variable supporting rods. The corrugated beams are arranged on the outer sides of the sliding key sections and the sliding groove sections and are provided with wave crest and wave trough structures; the sliding key sections and the sliding groove sections are located on the inner sides of the corrugated beams, arranged at intervals and distributed in a staggered mode. The corrugated beams, the sliding key sections and the sliding groove sections are all of arc-shaped structures, and wave crests and wave troughs of the corrugated beams are arranged in the radial direction of the arc-shaped structures; the sliding key section and the sliding groove section are mutually nested to form an arc-shaped supporting sliding way, and the wave trough of the corrugated beam slides on the arc-shaped supporting sliding way; the height-variable supporting column and the length-variable supporting rod are hinged to the two ends of the sliding key section respectively. The shape of the fuselage can be greatly changed, the large-size fuselage can provide a large space to carry more fuel, and the small-size fuselage is beneficial for reducing flight resistance and improving the fuel utilization rate.

The invention relates to the technical field of layered composite fabrics, in particular to a flame-retardant-fiber-based multilayer flame-retardant composite fabric for an aircraft cabin and a preparation method of the flame-retardant-fiber-based multilayer flame-retardant composite fabric. The preparation method comprises the following steps: reacting acrylic acid with DOPO to prepare modified DOPO; reacting the epoxy POSS with the modified DOPO to prepare a composite flame retardant; reacting the hydrophilic grafted polyamide with a composite flame retardant and amino-modified nano-montmorillonite to prepare modified polyamide; carrying out melt spinning and chopping on the modified polyamide to obtain flame-retardant fibers; the flame-retardant fibers are spun into flame-retardant yarns, then the flame-retardant yarns are spun into a fabric, and a surface fabric is obtained; the flame-retardant fibers and the polyester fibers are blended to obtain blended yarns, and then the blended yarns are spun into a fabric to obtain a bottom layer fabric; and compounding the surface layer fabric and the bottom layer fabric to obtain the composite fabric. The prepared composite fabric has excellent flame retardant property and antistatic property, and can be applied to aviation cabins.

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 invention belongs to the technical field of aircraft structure design, and discloses a fuselage frame preparation process and a fuselage frame. According to the fuselage frame preparation process, preparation of the first woven preformed body and the second woven preformed body is achieved through the weaving operation of the weaving machine, the preparation efficiency and the structural integrity of the first woven preformed body and the second woven preformed body are improved, and the influence of manual operation on the fuselage frame forming quality is avoided. The preparation of the first weaving body and the second weaving body is realized by cutting operation of cutting guide parts arranged on the first weaving auxiliary core mold and the second weaving auxiliary core mold in advance, so that the preparation precision is guaranteed, and the forming quality of the fuselage frame is improved. The invention further provides the machine body frame, the machine body frame is prepared through the machine body frame preparation technology, the manual operation participation degree is reduced in the machine body frame preparation process, the influence of manual operation errors on the forming quality of the machine body frame is avoided, the forming quality and forming efficiency of the machine body frame are improved, and the manual labor intensity is relieved.

An assembly is provided for an aircraft. This aircraft assembly includes an airframe structure, a propulsion system and a vane. The airframe structure includes a fuselage and at least two wings. The fuselage extends longitudinally along a longitudinal centerline from a forward end of the fuselage to an aft end of the fuselage. The wings extend laterally from opposing sides of the fuselage. The propulsion system is mounted to the airframe structure. The propulsion system includes an airflow inlet into the propulsion system. The airflow inlet is spaced vertically out from a top side of the airframe structure. The vane projects vertically out from the top side of the airframe structure to a distal end of the vane. The vane laterally overlaps and is longitudinally forward of the airflow inlet.

The invention relates to a cross-medium aircraft variant nose cone with dynamically adjustable surface wettability, and belongs to the field of cross-medium aircrafts. Comprising a variant skeleton and a flexible skin. Under the heating condition, the variant framework can drive the flexible skin to deform, the flexible skin keeps the deformed state after normal temperature is recovered, and therefore curvature distribution and section characteristics of the nose cone structure are changed. The surface of the flexible skin is covered with a wettability intelligent response coating. After the wettability intelligent response coating is irradiated by ultraviolet rays, the surface wettability can be converted into a super-hydrophilic state, and the wettability intelligent response coating can be recovered into a super-hydrophobic state after being heated. The head configuration and the surface wettability layout of the cross-medium aircraft are adjusted according to the specific water entry working condition, the hydrodynamic force and the hydrodynamic moment borne by the aircraft in the water entry process are changed through the influence effect of the configuration and the surface wettability on water entry cavitation evolution, the impact load is reduced as needed, and the attitude and trajectory stability is improved; the purposes of being compatible with various launching platforms and guaranteeing the water entry efficiency and safety are achieved.

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.

A craft is provided comprising: a propulsion system; a fuselage coupled with the propulsion system, wherein the fuselage is configured to carry a payload and fuel for the propulsion system; and a pre-assembled wing coupled with the fuselage; wherein the craft is designed to take-off without using a runway and does not have landing gear to land.

A method of aligning an aircraft with a landing platform in motion comprises measuring a GPS heading with at least one GPS sensor positioned at a known location relative to the landing platform while the landing platform is in motion, measuring an orientation of the aircraft with an orientation sensor fixed relative to the aircraft, calculating an orientation of the landing platform from the GPS heading, calculating an orientation offset between the measured orientation of the aircraft and the calculated orientation of the landing platform, and changing an orientation of the aircraft or the landing platform to reduce the orientation offset. A system for landing and securing an aircraft in an enclosure, a system for disconnecting a tether from an aircraft, and a system for landing an aircraft in an enclosure are also described.

The invention discloses a variable-pore phase-change sweating cooling bionic structure based on fractal characteristics. The variable-pore phase-change sweating cooling bionic structure comprises a variable-pore porous sweating layer and a regulation and control unit, wherein the variable-pore porous sweating layer is composed of fractal flow equalizing plates which are sequentially arranged in a stacked mode in the heat flow direction; the fractal flow equalizing plate is of an N-stage bifurcated tree-shaped structure and is used for equally dividing one path of cooling working medium into 2N paths of cooling working medium, and N is an integer ranging from 4 to 6; the variable-pore porous sweating layer is divided into an upper-layer high-heat-flow area and a lower-layer low-heat-flow area in the thickness direction, the upper-layer high-heat-flow area is used for promoting phase change and efficient heat exchange of a cooling working medium, and the lower-layer low-heat-flow area is used for promoting liquid-phase permeation of the cooling working medium; and the regulation and control unit predicts the phase change leading edge position S (t) in real time based on a neural network, and dynamically regulates the cooling working medium flow Q (t) input into the inlet of the fractal flow equalizing plate according to the S (t). The invention further discloses a design method of the structure. By introducing the bionic variable porosity characteristic, coupling of flowing heat exchange is achieved, the cooling efficiency of thermal protection is improved, and consumption of a heat exchange working medium is reduced.

The invention discloses an electric vertical take-off and landing manned aircraft which comprises an aluminum alloy rack, a second carbon fiber tube and supporting legs, one side of the aluminum alloy rack is connected with a right guardrail, the other side of the aluminum alloy rack is rotationally connected with a left guardrail through a winding shaft seat, and a vibration reduction assembly comprises a vibration reduction seat arranged on the first carbon fiber tube in a sleeving mode. The upper vibration reduction seat and the lower vibration reduction seat are fastened through bolts at the penetrating positions of the first carbon fiber pipes and the aluminum alloy rack, the rubber vibration reduction pads absorb high-frequency vibration energy through elastic deformation of the rubber vibration reduction pads, direct transmission of vibration to the rack is effectively blocked, the second carbon fiber pipes can be rotationally folded through the hinge seats, the transverse size of the whole machine is greatly reduced after folding, and the whole machine is more compact in structure. The occupied area is remarkably reduced, storage in a narrow space is facilitated, pushing can be achieved only by pressing down the tail end during transition on the flat ground, extra tools or complex operation procedures are not needed, the gravity center of the machine body is naturally transferred to the supporting wheels along with posture adjustment in the pushing process, and transition flexibility and efficiency are improved.

A connecting rod coupling device for attaching a component to an aircraft fuselage and / or stiffening the fuselage has a first coupler, and a second coupler couplable to the first coupler. The first coupler has an elongate body with a first channel on a lateral surface along a first longitudinal axis. The second coupler has a hollow-cylindrical portion, which, on an inner face along a second longitudinal axis, has a second channel complementing the first channel. The hollow-cylindrical portion has a radial cutout through which the elongate body is radially insertable into the second channel from outside the hollow-cylindrical portion. The second coupler has a closure body, which closes the radial cutout, and can be positioned removably in the cutout to complete the second channel on a radially internal face. The second coupler also has an element, pushable onto the hollow-cylindrical portion to secure the closure body on the hollow-cylindrical portion.

The invention relates to a preform of an aircraft bulkhead and a manufacturing method thereof. According to the invention, the preform of the bulkhead has edge strips located at both ends thereof and a web located in the middle of the preform and connecting the edge strips, the web has a radian in the longitudinal direction of the web, and a 0-degree yarn extending at an angle of 0 degree relative to the longitudinal direction of the web without wrinkles is laid on the web. According to the invention, the 0-degree yarn can be arranged on the web plate of the bulkhead preform in the longitudinal direction without wrinkles, so that the 0-degree yarn has tension in the longitudinal direction of the web plate, thereby meeting the stress requirement on the bulkhead in the longitudinal direction, ensuring the smoothness of preform fibers, improving the adaptation degree of the bulkhead preform to a curved surface, and improving the production efficiency of the bulkhead preform. The manufacturing efficiency of the bulkhead preformed body is improved, and the manufacturing quality and reliability are ensured.

The application provides a flat threshold maintenance passage structure and an operation method. The structure comprises a large opening reinforcing structure assembly 100, a quick-release floor assembly 200, a footboard assembly 300, a door frame assembly 400, a stepping reinforcing assembly 500, a mechanical support rod assembly 600, a bolt locking mechanism 700, a maintenance passage door 800, an external hinge 900 and a boarding handle 1000. The large opening reinforcing structure assembly 100 comprises an upper reinforcing horizontal piece 101, a lower reinforcing horizontal piece 102, a reinforcing front partition frame 103, a reinforcing rear partition frame 104, a horizontal piece outer flange 105, a stretching belt plate 106, an upper horizontal piece inner flange 107, a stretching belt plate 108, a lower horizontal piece inner flange 109, a counter-pulling angle box 110 and a counter-pulling angle piece 111. The quick-release floor assembly 200 comprises an upper panel 201, an aluminum honeycomb 202, a lower panel 203 and a bowl-shaped piece 204. The footboard assembly 300 comprises a foldable cantilever footboard 301, a locking mechanism 302, a bolt 303, a partition frame T-shaped outer flange 304, a support lug 305, a rib 306 and a stepping pedal 307. The door frame assembly 400 comprises a door frame 401 and a skin 402.

The invention discloses a multi-layer combined unmanned aerial vehicle and a control method, and relates to the field of high-altitude long-endurance unmanned aerial vehicles, the multi-layer combined unmanned aerial vehicle comprises a plurality of sub-aircrafts, each sub-aircraft comprises a fuselage, an empennage, wings, wingtips and protrusions, the wingtips are arranged at the ends, away from the fuselage, of the left and right wings, and the wingtips can be in butt joint with the wingtips of the other aircraft; protrusions are arranged on the upper surface and the lower surface of each wing, the end faces of the protrusions can be in butt joint with the end faces of the protrusions of another airplane, and the protrusions can be used for containing an aircraft undercarriage at the same time. Through aggregation and recombination of a plurality of small unmanned aerial vehicles, a multi-layer combined unmanned aerial vehicle with a high aspect ratio is formed, and the combined unmanned aerial vehicle can effectively overcome the defect that the rigidity of wings of a traditional unmanned aerial vehicle with a high aspect ratio is insufficient and can carry an oversized radar conformal antenna with long wave bands such as meter waves and the like. And long-time air-hang detection of high-value targets such as stealth aircrafts and the like is realized.

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 system and method for an adaptive hardware system. The adaptive hardware includes rail segments and reconfigurable anchor assemblies. The anchor assemblies are secured onto different portions of the rail segments at anchor installation ports. The anchor assemblies are adaptable for accepting common fastening elements (e.g., bolts), such that interior monuments and fixtures may be accommodated without interior aircraft design changes.

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.

The invention relates to the technical field of thermal protection composite materials, in particular to a thin-wall gradient thermal protection structure and application. The thermal protection structure sequentially comprises, along a thickness direction, an outer surface layer; the ablation-resistant layer is composed of at least one layer of quartz fiber cloth, and the distance d between the outer side face of the quartz fiber cloth and the outer surface layer is larger than 0 mm and smaller than or equal to 0.3 mm; the heat insulation layer is positioned on the inner side of the ablation-resistant layer; the total thickness of the thermal protection structure is smaller than or equal to 5 mm, and the overall density of the thermal protection structure is smaller than or equal to 0.6 g / cm. According to the thin-wall gradient thermal protection structure, under the extreme lightweight constraint that the thickness is smaller than or equal to 5 mm and smaller than or equal to 0.6 g / cm < 3 >, the extra thermal insulation gain larger than or equal to 5% is obtained through the principle of gradient layering and millimeter-level near-wall ablation resistance.

The disclosure provides a battery assembly (116). The battery assembly (116) includes a plurality of stiffener assemblies (118). Each stiffener assembly (118) in the plurality of stiffener assemblies (118) comprises a stiffener body (130), an inner surface (132) of the stiffener body (130) that forms a hollow channel (134) through the stiffener body (130), a cathode assembly (144), and an anode assembly (164). The battery assembly (116) further includes a pump (101) in fluid communication with the hollow channel (134) of the stiffener bodies (130), where the pump (101) is configured to circulate an electrolyte through the hollow channels (134).

The application relates to a locking structure of a dismounting structure of a unmanned aerial vehicle. The locking structure of the dismounting structure of the unmanned aerial vehicle comprises a connecting assembly and a locking assembly, the connecting assembly comprises a connecting base and a rotating piece for outputting power; the rotating piece is hinged to the connecting base; the locking assembly comprises a locking piece hinged to the connecting base, the locking piece is provided with a locking portion, and the locking portion is connected with the rotating piece when the connecting assembly is in a connecting state to prevent the included angle between the rotating piece and the connecting base from increasing. The application can ensure that the connecting state is relatively stable, prevent the occurrence of unhooking, facilitate the operation of dismounting the parts of the unmanned aerial vehicle by holding the rotating piece with hands, ensure that the jacking protrusion does not interfere with the rotation of the rotating piece to be opened, enable the tongue to automatically move away from the supporting body, and facilitate the dismounting of the parts of the unmanned aerial vehicle.

An aircraft component including a structural foam having a predetermined shape, an outer flange located near a first end of the structural foam, an inner flange located near a second end of the structural foam, a web skin encapsulating the inner flange and at least a portion of the structural foam wherein the structural foam is affixed to the web skin, wherein the aircraft component comprises a portion of a blended wing body aircraft (BWB), the BWB having no demarcation between a wing and a main body of the BWB.

A support box apparatus includes longitudinal members that span forward to aft and transverse members that span inboard to outboard between the longitudinal members. The longitudinal members and the transverse members are formed of composite honeycomb sandwich panels. The support box apparatus also includes attachment components configured to couple the longitudinal members or the transverse members to structural components of an aircraft, electrical wiring attached to the longitudinal members or the transverse members, environmental control system (ECS) ducting attached to the longitudinal members or the transverse members, and life support lines attached to the longitudinal members or the transverse members.

A fuselage structure of an aircraft includes a first fuselage section and a second fuselage section connected to the first fuselage section, the second fuselage section including at least one first fuselage vessel extending the first fuselage section in a longitudinal direction, with the first fuselage section and the first fuselage vessels comprising circumferential flanges arranged at longitudinal ends of the first fuselage section and the first fuselage vessels extending radially in the first fuselage vessels and the first fuselage section wherein fasteners extend through abutting flanges of the first fuselage section and the fuselage vessel to establish a force-fit connection between the first fuselage section and the first fuselage vessels, when assembled, and wherein the first fuselage vessel provides a storage space for at least one tank structure. Also disclosed are a method for assembling the fuselage structure and an aircraft including the fuselage structure.

The invention is related to a structural panel for use in aeronautical applications. The structural panel includes a composite skin (200) that comprises composite material and a stiffening framework that is rigidly attached to the composite skin (200). The composite skin (200) is configured as a load bearing monolithic structure (370) with a structure thickness (310) of less than 1 mm. The composite skin (200) comprises a composite laminate (480) with a plurality (400) of carbon composite plies (405). Each one of the plurality (400) of carbon composite plies (405) comprises a ply thickness (410) in a range from 0.02 mm to 0.05 mm.

An interior aircraft section having a cover for a gap between a lining and a monument. The cover comprises a basic body, at least one mounting end configured to be fixedly mounted to the monument, and a connector configured to fixedly mount the basic body to the lining. The connector is arranged in an area of the basic body overlapping the lining. Also, an aircraft having such interior aircraft section.

A method for manufacturing a fuselage structural component (200) for a triangular section of an aircraft is disclosed, comprising the following steps: providing a one-piece sheet (211, 212, 213) of frame material, the sheet having an opening (250), and deforming at least one inner edge of the sheet (211, 212, 213) forming the opening (250) into an inner flange (215, 216, 217), and deforming at least one outer edge of the sheet (211, 212, 213) into an outer flange (214). The one-piece and deformed sheet (211, 212, 213) forms, adjacent to the opening (250), a frame section (210), a bottom support (212), and a vertical support (213).

Assembly comprising a wall and at least one beam provided with at least one attachment passing through the wall, central wing box of aircraft comprising said assembly. The invention relates to an assembly comprising: at least one wall (54) having an inner face (Fint) and an outer face (Fext) opposite the inner face (Fint), at least one beam (52) positioned against the outer face (Fext) of the wall (54) and connected to the latter by connecting elements (56), at least one attachment (62) projecting from the inner face (Fint) of the wall (54) integral with the beam (52) and forming with the latter a single and the same piece made in one piece, the wall (54) comprising, for each attachment (62), a through hole (66) to allow the attachment (62) to pass through the wall (54).The invention also relates to an aircraft wing center box comprising at least one such assembly and to an aircraft comprising said wing center box. Figure 11.

The novel unmanned aerial vehicle arm comprises a supporting assembly and a limiting assembly, the limiting assembly is installed in the supporting assembly and used for further limiting the supporting assembly, the supporting assembly comprises a vehicle body, the vehicle body extends outwards to be provided with a first vehicle arm, and the first vehicle arm is provided with a second vehicle arm. The first vehicle arm is arranged at one end of the vehicle body, a second vehicle arm corresponding to the first vehicle arm is arranged at one end of the vehicle body, and the second vehicle arm is arranged at one end of the vehicle body. The novel unmanned aerial vehicle arm can show higher flexibility and adaptability in a complex and changeable task environment. The unmanned aerial vehicle can be rapidly adjusted according to different task requirements through the adjustable structures of the first vehicle arm and the second vehicle arm and the ingenious layout of the arc-shaped plate, the connecting support and other components, and therefore the operation efficiency and accuracy are effectively improved.