735results about "Fuselage frames" patented technology

The invention provides an aircraft framework and an aircraft, the aircraft framework is used for the aircraft flying in the atmosphere, the aircraft framework comprises a cage serving as a framework main body, the cage bears the self weight and the flight load of the whole aircraft, at least two first turbojet engines can be contained in the cage, and the first turbojet engines are arranged in the cage; the cage is open at least in the lower portion and used for allowing the jet flow of the first turbojet engine to pass through. Through the arrangement of the cage which is a skeleton main body, a power source device of the aircraft is collected and gathered, so that the miniaturization of the whole aircraft is realized.

The invention relates to a self-adaptive energy-absorbing conformal head cap of an air-water cross-medium aircraft, which is mainly applied to aircrafts capable of switching motion between two media of air and water. The head cap comprises a top cover, a skin combination, an energy absorption mechanism, a shape preserving mechanism and a modular adaptive interface, wherein the skin combination comprises an outer-layer skin and an inner-layer skin; the energy absorption mechanism is arranged in the inner-layer skin; the conformal structure is arranged between the inner-layer skin and the outer-layer skin; the modular adaptive interface comprises a base and a switching section, the base is connected with one end of the inner-layer skin, and the switching section is connected with the aircraft body; and the top cover is connected with the other end of the inner-layer skin. Stable hydrodynamic performance and low underwater resistance can be provided under different water entry speeds and angles, head appearance damage caused in the deformation process of a traditional energy absorption structure is effectively avoided, and the key technical problems in design of the air-water cross-medium aircraft are solved.

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 belongs to the technical field of flexible structure design, and particularly relates to a variable-arc-length arched beam structure which comprises a sliding groove section, a sliding key section, a corrugated beam, a variable-height supporting column and a variable-length supporting rod. 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; when the shape and the arc length of the arched beam are changed, the variable-height supporting columns and the variable-length supporting rods can automatically generate height changes, then the sliding key sections and the sliding groove sections are driven to relatively slide, and meanwhile the corrugated beams evenly distribute the arc length changes to the positions between the adjacent wave crests, so that external loads are evenly distributed to the arched beam. Therefore, the size of the fuselage is correspondingly reduced along with fuel consumption.

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.

An aircraft that is certifiable for freighter transport is disclosed herein. The aircraft can include a hybrid fuselage including a plurality of modular sections, wherein the plurality of modular sections includes a first section, a second section, and an intermediate section, and wherein the intermediate section is disposed intermediate the first section and the second section. The aircraft can further include a plurality of systems traversing the plurality of modular sections and a plurality of structural components configured to mechanically couple the intermediate section to the first section and the second section of the plurality of modular sections. The first section and the second section are configured for use with a fuselage of a passenger aircraft, and wherein the intermediate section is configured for use with a fuselage of a freighter aircraft.

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.

Disclosed is a fuselage 100, 100', 100" of or for an aircraft. The fuselage comprises a shell 10, at least one hydrogen tank 20, an energy absorbing structure 30 arranged between the shell 10 and the at least one hydrogen tank 20, and a support structure 40, 40' supporting at least a portion of the at least one hydrogen tank 20 within the shell 10. The support structure 40, 40' comprises an interstitial section 52, 54i located between the at least one hydrogen tank 20 and the energy absorbing structure 30. Furthermore, the support structure 40, 40' comprises a strut system 60 supporting the at least said portion of the one hydrogen tank 20 against the interstitial section 52. The support structure 40' may further brace at least said portion or at least a further portion of the at least one hydrogen tank 20, 20b in a roof section 100R of the fuselage, e.g. by means of one or various support elements 70'. Further disclosed is an aircraft comprising such fuselage 100, 100', 100".

A blended wing body (BWB) aircraft including a blended wing body having a main body and wings with no clear demarcation between the wings and the main body along leading edges of the BWB aircraft, a passenger cabin located within the main body and more than one passenger bays located within the passenger cabin, wherein a perimeter of at least one passenger bay is demarcated by at least a divider and wherein the at least one passenger bay is substantially parallel to at least one leading edge of the BWB aircraft.

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.

Proposed is a collision load distribution structure of a fuselage, wherein the structure includes a support unit positioned between a front unit and a rear unit of the fuselage and configured to allow a collision load of the fuselage to be transferred by being connected to the front unit and the rear unit, a wing unit positioned at the inside of the support unit and connected to the support unit to transfer the collision load of the fuselage, and a seat unit coupled to a connection frame mounted in a second-row passenger space of a floor frame constituting the front unit and fastened to a first rear frame extending in a height direction from the floor frame toward the support unit.

The invention belongs to the technical field of morphing aircraft structure design, and particularly relates to a deformable fuselage structure for a hypersonic speed environment, which comprises a deformable fuselage frame and a deformable fuselage skin, the deformable machine body frame comprises a sliding groove section, a sliding key section, a corrugated beam, a height-variable supporting column and a length-variable supporting rod. 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 deformable fuselage skin comprises a deformable supporting framework, sliding scales and an S-shaped heat insulation plate. The shape and the section size of the fuselage can be greatly changed by controlling the shape and the section size of the fuselage, so that the large-size fuselage can provide enough space to carry more fuel, the small-size fuselage section reduces flight resistance and improves the fuel utilization rate, the thermal insulation skin prevents pneumatic heat from being transmitted into the fuselage, and the fuselage is easy to implement based on existing materials and mature processes.

Proposed is a structure. The structure includes a front body frame configured to surround a front surface of an airframe, a crash unit disposed at a front end of the body front frame, and a dash reinforcement assembly fastened to the front body frame and disposed on the rear surface of the crash unit.

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.

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 present invention belongs to the technical field of aircraft. Particularly, the present invention provides an aircraft fuselage section (1) with improved crashworthiness. The aircraft fuselage section (1) comprises a fuselage structure, a floor (3) connected to the fuselage structure, a cushioning element (4) arranged on the floor (3) and crash absorption means (5) arranged between the fuselage structure and the floor (3).

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.

The present application discloses a method of assembling or disassembling an aircraft. The assembly method includes a mounting a first engine nacelle to a first wing, mounting a second engine nacelle to a second wing, causing the first wing to be extended relative to a center wing, and causing the second wing to be extended relative to the center wing.

An example carriage is configured for mounting to an unmanned aerial vehicle. The carriage generally includes a housing assembly configured for mounting to the unmanned aerial vehicle, a movable grip mounted to the housing assembly for movement between a capturing position and a releasing position, a latch device, and a driver. The latch device has a latching state and an unlatching state, is configured to retain the movable grip in the capturing position when the latch device is in the latching state, and is configured to permit movement of the movable grip from the capturing position to the releasing position when in the unlatching state. The driver is operable to transition the latch device from the latching state to the unlatching state.

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.

An embodiment aircraft boom (66, 76) includes a boom member (34, 42), a vertical stabilizer (40, 50), a rear landing gear (114, 116), and a front landing gear (110, 112). The boom member (34, 42) may extend in a first direction and have a front internal structural member (154) disposed therein, where a portion of the front internal structural member (154) is a first attachment point for a first wing. The vertical stabilizer (40, 50) may extend from the boom member (34, 42), where the vertical stabilizer (40, 50) has a rear internal structural member (120) disposed therein and extends from the boom member (34, 42) and through an interior of the vertical stabilizer (40, 50), and where a portion of the rear internal structural member (120) is a second attachment point for a second wing. The rear landing gear (114, 116) may be coupled to the rear internal structural member (120). The front landing gear (110, 112) may be coupled to the front internal structural member (154) of the boom member (34, 42).

The invention provides an air-breathing aircraft for long-term flight in an upper atmosphere, and relates to the technical field of aircrafts in the upper atmosphere, the air-breathing aircraft comprises an aircraft body, the aircraft body is provided with an air-breathing electric propulsion system, the outer contour of the aircraft body is cylindrical, and an air inlet channel is arranged in the center of the aircraft body; the air inlet channel comprises a gradually-shrinking conical inlet pipeline, an inwards-concave gradually-shrinking compression pipeline, a cylindrical ionization acceleration pipeline and a gradually-expanding conical outlet pipeline which are sequentially arranged from front to back. The solar cell wings are arranged on the outer side of the aircraft body, the solar cell wings are forward-swept trapezoidal wings, and the solar cell wings can supply power to the air suction type electric propulsion system; the problem that in the prior art, technical defects exist in the aerodynamic configuration design of how to achieve thrust and resistance balance of an aircraft flying for a long time in the upper atmosphere is solved.

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.