1589results about "Fuselages" patented technology

An aircraft includes a flight deck including a flight deck door configured to be moved between an open position and a closed position, a first monument, second monument, and a secondary barrier door moveably secured between the first monument and the second monument. The secondary barrier door is configured to fold into a stowed position in which a path to the flight deck is open, and unfold into a deployed position in which the path to the flight deck is closed.

An aircraft comprising an aircraft transceiver to transfer data to and from the aircraft by free-space-optical communication when parked. The aircraft transceiver comprises a transmitter to generate outgoing light carrying aircraft-to-ground data; a lens to collimate the outgoing light to generate a collimated beam; a beam steering device to transform the collimated beam to generate a steered beam; a control system to operate the beam steering device to adjust an angle of the steered beam; and a receiver to sense incoming light carrying ground-to-aircraft data. The receiver receives the incoming light via the beam steering device and the lens. Also a vehicle with a similar transceiver to transfer data to and from a parked aircraft by free-space-optical communication.

A door assembly configured to be mounted in an opening of a cabin area of an aircraft. The door assembly includes a door configured to be movable between a deployed position across the opening and a stowed position out of the opening. A lock is mounted to the door and movable between a locked position and an unlocked position. An engagement assembly is mounted to the door and movable between an engaged position and a disengaged position. The engagement assembly includes a plunger, a lock mechanism, and an elongated member that extends between the lock and the lock mechanism. In the engaged position, the lock mechanism and the elongated member position the plunger in an extended position outward from the door. In the disengaged position, the lock mechanism and the elongated member position the plunger in a retracted position relative to the door.

A method for automated spacecraft design includes receiving mission parameters provided by a client for spacecraft design, generating one or more concept designs based on the received mission parameters, a concept design including one or more components selected for each subsystem included in a designed spacecraft, for the concept design, automating a generation of a frame design that includes a plurality of thermal management components to be formed through a single metal-additive manufacturing process, performing a series of analyses or simulations to evaluate a performance of the frame design, and generating a digital twin representing a virtual representation of the designed spacecraft based on the generated frame design and the series of analyses and simulations.

A modular vertical take-off and landing aircraft for transporting persons and / or loads, comprising a flight module with multiple drive units arranged on a supporting framework structure, each drive unit having an electric motor and a propeller operatively connected to the electric motor, a transport module having a conveying pod and a connection device for connecting the conveying pod to the flight module. The connection device has an elongated shaft, one end of which is attached to the conveying pod, and a coupling device for connecting the flight module to another end of the elongate shaft of the transport module.

An embodiment herein provides a re-configurable unmanned aerial vehicle that re-configures its shape based on the shape, size, weight of a payload, and efficiently performs payload delivery in real-time. The re-configurable unmanned aerial vehicle includes one or more rotor units placed at corners and is connected by one or more scissor units. The re-configurable unmanned aerial vehicle approaches the payload in a first location, and analyses the position and dimension of the payload with a camera, that enables the one or more scissor units to adjust its length by at least one elongation or compression following size and shape of the payload and fit the payload within the re-configurable unmanned aerial vehicle. The re-configurable unmanned aerial vehicle takes off carrying the payload from the first location and lands at a second location.

A wing fuselage joint for an aircraft having joining members which include a first configuration and a second configuration. The joining members of the first configuration include a platform and bracket configured to pivot in a lateral plane and translate in a longitudinal direction. The joining members of the second configuration include a platform and bracket configured to pivot in a longitudinal plane and translate in a lateral direction. The joining members of either configuration join the fuselage structure and the spar of an aircraft together.

The present invention relates to a transparent colorless spinel microcrystalline glass, and preparation and a use thereof. The microcrystalline glass contains (Zn, Mg)Al2O4 as a principal crystalline phase and ZrO2 as a secondary tetragonal crystalline phase, and does not contain TiO2. In mole percentage on an oxide basis, an oxide contained therein satisfies the following relational expressions: Al2O3 / (MgO+ZnO)=1.40~2.00, and K2O+Li2O+Na2O+B2O3) / (SiO2+Al2O3)=0.10~0.20. The microcrystalline glass is obtained by performing heat treatment at a temperature of being not greater than 800°C. The microcrystalline glass has a transmittance of 85.0% or above for light having a wavelength of 550 nm. Anti-abrasive paper drop testing is performed on the microcrystalline glass by using 80-mesh abrasive paper, the average anti-abrasive paper drop height of the microcrystalline glass is greater than or equal to 1.00 m; the Vickers hardness is 800kgf / mm2 or above, and the fracture toughness of the microcrystalline glass is 1.200MPa·m1 / 2 or above. According to the present invention, the colorless transparent spinel microcrystalline glass having high transparency and excellent mass productivity can be obtained, and the microcrystalline glass can be subjected to chemical strengthening to obtain a chemically strengthened microcrystalline glass having excellent deep stress characteristics.

An unmanned aerial vehicle (UAV) having wings stowed against a fuselage of the UAV in a first arrangement is disclosed. Methods and systems for deploying the wings into a second arrangement are disclosed. For example, after a launch of the UAV, the UAV monitors for at least one precondition. The at least one pre-condition being a pre-condition associated with deploying wings of the UAV into the second arrangement. Upon detecting the at least one precondition, the wings of the UAV are deployed into a second arrangement. Deploying the wings comprises activating, in response to detecting the at least one precondition associated with the UAV, a gearbox configured to transition the wings from the first arrangement to the second arrangement. Roll control may be maintained throughout launch and deployment.

The invention provides a bionic hollow petal-shaped SiC fiber porous thermal insulation material as well as a preparation method and application thereof, and belongs to the technical field of thermal insulation materials. The preparation method comprises the following steps: taking a viscose-based carbon fiber felt as an initial structure template, dipping the carbon fiber felt in siloxane sol, taking out the carbon fiber felt, and carrying out suction filtration to remove redundant sol on the surfaces of carbon fibers; curing the impregnated carbon fiber felt, and carrying out heat treatment to obtain the carbon fiber felt of which the carbon fiber surface is covered with a uniform SiC coating; and carrying out thermal oxidation treatment on the carbon fiber felt with the SiC coating to obtain the bionic hollow petal-shaped SiC fiber porous thermal insulation material. Compared with a traditional sol-gel method, the method has the advantages of being short in preparation period, low in cost, small in volume shrinkage rate and the like, and the obtained bionic hollow petal-shaped SiC fiber porous heat insulation material has good rebound resilience and excellent high-temperature heat insulation performance and has wide application prospects in a high-temperature oxidation environment.

The invention discloses an Al-Mg-Zn aluminum alloy plate with excellent strong plasticity and light weight characteristics as well as a preparation method and application of the Al-Mg-Zn aluminum alloy plate. The alloy comprises the following components in percentage by mass: 5.0%-5.5% of Mg, 1.5%-2.0% of Zn, 0.4%-0.5% of Mn, 0.1%-0.3% of Cu, 0.15%-0.25% of Sc, 0.1%-0.15% of Zr, 0.05%-0.15% of Ag, less than or equal to 0.1% of impurities and the balance of Al. The density is not more than 2.63 g / cm < 3 >, the tensile strength is not less than 470 MPa, and the ductility is not less than 23%. The preparation method comprises the steps of high-purity raw material smelting and casting, two-stage homogenization, hot rolling, intermediate annealing, cold rolling, solid solution treatment and single-stage aging. Wherein the total pressing amount of hot rolling is 80-85%, the total pressing amount of cold rolling is 20-35%, and the single-stage aging time is 120-140 DEG C * 16-24 h. The microstructure of the alloy is characterized in that the average size of dispersed AlSc particles is smaller than or equal to 100 nm, the area fraction of deformed grains is larger than or equal to 70%, the average size of nanometer beta-AlMg and T-Mg (Al, Zn) composite precipitated phases is smaller than or equal to 20 nm, the number density is larger than or equal to 1 * 10 < m >, the area fraction of micron-order second phases is smaller than or equal to 0.5%, and the alloy is only formed by impurity Fe. According to the method, the traditional two-stage aging process is simplified, low density, high strength and high plasticity are synchronously improved, and the method is suitable for large-scale industrial application.

The present invention discloses a porous material transpiration cooling leading edge structure, which includes multiple layers of porous material layers and a cooling cavity provided on one side of the multiple layers of porous material layers. A coolant supply channel is provided in the cooling cavity, and the coolant supply channel is used to supply coolant to the cooling cavity. The multiple layers of porous material layers include an outer layer, an intermediate layer, and an inner layer, and the material of the inner layer is a shape memory alloy porous material. The present invention can adaptively adjust the porosity of the porous material, automatically adjust the usage amount of the coolant, effectively improve the usage efficiency of the coolant, and further enhance the thermal protection effect in the high-temperature tip region. Moreover, the shape memory alloy material can be reused. The porous material transpiration cooling leading edge structure can be used for structures such as the nose cone of an aircraft, the leading edge of a rudder, and the leading edge of a wing that are subjected to more high-temperature heat fluxes.

A preparation method of porous hafnium carbide superhigh-temperature ceramic is characterized by comprising the following steps: dispersing graphene oxide in a solvent to form a graphene oxide dispersion liquid, mixing a monomer hafnium source solution with the graphene oxide dispersion liquid, and carrying out hydrothermal treatment on a hafnium oxide carbon precursor network structure to obtain the porous hafnium carbide superhigh-temperature ceramic. The organic-inorganic hybrid precursor is subjected to two-stage heat treatment, the first stage is inorganic treatment, the second stage is carbon thermal reduction reaction, and the porous hafnium carbide ultrahigh-temperature ceramic material is obtained. Graphene oxide is used as a structure-directing agent and an in-situ carbon source, a monomer hafnium source is uniformly anchored on the surface of a GO sheet layer on the molecular scale, an organic-inorganic hybrid precursor is formed, after the precursor is subjected to hydrothermal crosslinking, Hf species and a carbon skeleton are highly mixed on the nanoscale, and the Hf / GO composite material is obtained. And the subsequent carbon thermal reduction reaction can be efficiently completed in a low-temperature interval of 1300-1450 DEG C.

The invention provides an aircraft passenger accommodation unit (1000) comprising a seat (100), a shell (110) defining an opening (120), and a door (130). The door (130) is translatable along a first path across the opening (120) between a closed position in which passenger egress through the opening (120) is at least partially blocked, and an open position which allows passenger access between a seat (100) and an aisle (300). The door (130) is translatable in a second direction that is transverse to the first direction, between a first configuration and a second configuration. In the first configuration, the door (130) is translatable across the opening (120) along the first path, and when the door (130) is in the second configuration, the door (130) is translatable across the opening (120) along a second path. The invention also provides a door arrangement and a method of using a door.

An aircraft wing having a fixed wing and a moveably mounted wing tip device is disclosed. A first cable harness and a second cable harness which are received in a protective sleeve extend between the fixed wing and the wing tip device. The protective sleeve is in a secured configuration in which at least two opposing parts of the protective sleeve are secured together between the first cable harness and the second cable harness to provide the sleeve with a first channel in which the first cable harness is received a second channel in which the second cable harness is received, and a webbing region extending between the first channel and the second channel such that the first and second cable harnesses are held apart by the webbing region.

An aircraft section, between a cockpit and a cabin, comprises an aisle to access to the cockpit from the cabin, and a monument extending in a longitudinal direction and a transverse direction. The monument includes a first lavatory, a second lavatory adjacent thereto in the transverse direction and comprising the aisle, a wall separating the first lavatory from the second lavatory, a boundary wall delimiting the first lavatory and a portion of a rear wall of the cockpit, and a second boundary wall delimiting the second lavatory and a second portion of the rear wall of the cockpit. The boundary walls are parallel or continuous. The monument further includes an outer door movable between an open position, where the cockpit aisle is accessible from the cabin area, and a closed position, where the cockpit aisle is blocked with respect to the cabin area.

The invention provides a transparent bulletproof ceramic composite structure with an electric heating function and a preparation method of the transparent bulletproof ceramic composite structure. The composite structure comprises an anti-splashing layer (1), a first middle layer film (2), a multi-layer stacked structure (3) of glass sheets and films, a first transition layer (4), an electric heating film layer (5), a second transition layer (6), a ceramic layer (7), a first antireflection film layer (8), a second antireflection film layer (9), a third antireflection film layer (10), a second middle layer film (11), an inorganic glass layer (12), a third transition layer (13) and an anti-static layer (14) which are arranged in sequence. The composite structure provided by the invention has excellent impact resistance and bulletproof grade, can effectively resist the damage of high-speed elastic sheets or impact objects, and meets the requirements of high-safety application scenes such as weapon armors, armed airplanes and high-speed trains; and the anti-static film has clear optical observation capability, high transmissivity and anti-static property.

The provided is a connector and a method for connecting a motor and an arm of multi-rotor aircraft. The connector includes a base, a locking disc component, and a guiding positioning component; one side of the base is connected to the arm, the locking disc component includes a forward locking disc structure and a reverse locking disc structure; the guiding positioning component includes a forward guiding positioning structure and a reverse guiding positioning structure; the forward guiding positioning structure and the reverse guiding positioning structure are respectively embedded in the forward locking disc structure and the reverse locking disc structure, and one end of the forward locking disc structure is connected to a forward propeller motor, another end of the forward locking disc structure is inserted into a forward end of the base through the forward guiding positioning structure and connected to the forward end of the base.

An interchangeable fairing system has a removeable fairing with a common mounting pattern for mounting onto an aircraft and aerodynamically coupling a fixed fairing structure with a fuselage structure. Different types of fairings having the common mounting pattern are interchangeable, allowing for swapping of the different fairings mounted to the aircraft.

The invention relates to the technical field of aerospace, in particular to a ceramic 3D printing A sandwich structure with a broadband wave-transparent characteristic. According to the ceramic 3D printing A sandwich structure with the broadband wave transmission characteristic, good compatibility of broadband wave transmission, strength-structure accuracy and reliability is achieved, an obvious effect on broadband wave transmission of 1 GHz to 18 GHz is achieved, the dielectric constant and the tangent loss are in a reduction trend under the same frequency band, excellent stability is shown, and the dielectric constant and the tangent loss are reduced under 10 GHz. The dielectric constant of the ceramic 3D printing A sandwich structure with the broadband wave transmission characteristic is reduced by 40% or above compared with that of a full solid structure, the wave transmission rate is increased by 50% or above, the ceramic 3D printing A sandwich structure has excellent electromagnetic performance and mechanical performance, and support is provided for implementation of reliable service application targets such as diversion, heat protection, wave transmission, bearing and rain erosion resistance of hypersonic aircrafts.

A wing-fuselage truss joint assembly for an aircraft includes joining members secured to the aircraft's wing and fuselage structures. The joining members each include forward and aft wing joints and forward and aft fuselage joints connected by an upright member. The forward joining members are pivotally connected using interconnecting members and the aft joining members are connected using aft truss members. The forward-right and aft-right joining members and the forward-left and aft-left joining members are connected by truss link members and diagonal truss link members. The wing-fuselage truss joint assembly is configured to reduce forces experienced by the wing from being transferred to the aircraft fuselage.

A thermal gradient management assembly comprising a functional component. The functional component includes an exterior facing segment, where the exterior facing segment is configured to experience aerodynamic heating, and an interior segment with the exterior facing segment, where the interior segment is configured for isolation from the aerodynamic heating. The functional component includes a thermal gradient extending between the exterior facing segment and the interior segment. The thermal gradient management assembly also comprises a gradient mitigating heating system includes a power source and a gradient heating element in communication with the power source, wherein the gradient heating element is coupled with the interior segment. The gradient heating element is configured to diminish the thermal gradient between the exterior facing and the interior segments.

There is disclosed in one example an unmanned vertical lift aircraft, comprising: an airframe; a drive system configured for vertical takeoff and landing (VTOL); a power plant to power the drive system; and an internal cargo bay comprising a fore aperture and an aft aperture, the fore and aft apertures having dimensions substantially conforming to a cross section of the internal cargo bay, and the internal cargo bay comprising a substantially linear and open volume between the fore and aft apertures.

The invention relates to the technical field of aircraft windshields, in particular to an organic laminated windshield transparent part with asymmetrically connected edges. The organic laminated windshield transparent part comprises a composite transparent part provided with a mounting hole; the first pressing plate is arranged on the first surface of the composite transparent part; the second pressing plate is arranged on the second surface of the composite transparent part; the fastening piece is used for connecting the first pressing plate, the composite transparent piece and the second pressing plate; the composite lining is contained in the mounting hole in the axial direction of the composite lining partially or completely, the fastener penetrates through the composite lining, the composite lining comprises an inner lining body and an outer lining body provided with a plurality of branch lining bodies, the branch lining bodies correspond to at least one layer of the composite transparent part, the inner lining body is selected according to the function of the fastener, and the outer lining body is selected according to the function of the fastener. And each sub-bushing is selected according to the function of at least one layer of the corresponding composite transparent part. According to the windshield transparent part, the effects of load transmission, sealing, protection and the like can be improved, the stability and reliability of the windshield transparent part are improved, the using effect of the windshield transparent part is enhanced, and the service life of the windshield transparent part is prolonged.

A wing assembly for an aircraft, the wing assembly having a fixed wing and a wing tip device moveably mounted at the end of the fixed wing. A wiring harness extends between the fixed wing and the wing tip device, a length of the wiring harness being encased in a length of protective sleeve. The protective sleeve is pre-shaped, for example pre-molded, to follow a path along a helical curve.

Metal to composite joints and methods of forming are presented. A metal to composite joint for a platform comprises two metal structural components forming a portion of a first surface and a portion of a second surface of the metal to composite joint; and a dividing set of composite plies between the two metal structural components providing an escape path for volatiles between the two metal structural components.

A method for automated spacecraft design includes receiving mission parameters provided by a client for spacecraft design, generating one or more concept designs based on the received mission parameters, a concept design including one or more components selected for each subsystem included in a designed spacecraft, for the concept design, automating a generation of a frame design that accommodates the one or more components selected for each subsystem included in the designed spacecraft, the frame design being a Computer Aided Design (CAD) model that includes individual frame pieces generated, sized and fitted automatically via automated scripts compatible with a CAD system, including one or more structural design software applications; performing a series of analyses or simulations to evaluate a performance of the frame design; and generating a digital twin representing a virtual representation of the designed spacecraft based on the generated frame design and the series of analyses and simulations.

Systems, devices, and methods for a ground support system for an unmanned aerial vehicle (UAV) including: at least one handling fixture, where each handling fixture is configured to support at least one wing panel of the UAV; and at least one dolly, where each dolly is configured to receive at least one landing pod of the UAV, and where each landing pod supports at least one wing panel of the UAV; where the at least one handling fixture and the at least one dolly are configured to move and rotate two or more wing panels to align the two or more wing panels with each other for assembly of the UAV; and where the at least one dolly further allows for transportation of the UAV over uneven terrain.

An aircraft wing assembly includes a wing box including an upper surface, a lower surface, and a leading edge structure connecting the upper surface to the lower surface and secured to the upper surface at a connection point. The aircraft wing assembly includes a drive mechanism secured to the wing box and a high lift device mounted to the drive mechanism. The drive mechanism is configured to move the high lift device between a retracted position and an extended position in use, and wherein the high lift device is configured to overlap the connection point when in the retracted position.