1225results 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.

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 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.

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.

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 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.

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.

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.

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.

A tape structure for use in automated fiber placement method to manufacture a composite structure includes a first layer and a second layer connected to the first layer with the first layer being configured as a support layer and the second layer being configured as a pre-impregnated layer, wherein the first layer is removable from the second layer during or after placement of the second layer on the composite structure to be manufactured, an automated fiber placement device an automated fiber placement method and a composite structure manufactured in the method.

The invention provides a wave rider capable of vertically ascending and descending, and belongs to the technical field of wide-speed-range aircrafts, the wave rider specifically comprises a front edge shock wave generating surface, a central fuselage and sweepback wings located on the two sides of the central fuselage, the two sweepback wings are provided with mounting holes penetrating through the upper surface and the lower surface of the wave rider, and lift fans are mounted in the mounting holes; the lift fans are embedded into the mounting holes and used for providing vertical lift force for the wave rider, the axes of the lift fans are parallel to the vertical axis of the wave rider, and the lift fans on the two sweepback wings are symmetrically distributed on the two sides of the central fuselage. According to the treatment scheme, the vertical lifting capacity is provided for the wave rider, and the high-speed flight aerodynamic performance is considered.

The present invention addresses the problem of providing a hollow molded article that can be suitably applied to propeller blades for UAMs and drones, is lightweight, and has exceptional impact resistance. In order to solve the aforementioned problem, a hollow molded article comprises: a skin part (skin part (A)) composed of a composite material that contains a resin and reinforcing fibers, the skin part forming the outer surface of the hollow molded article; and a hollow part and a porous body in the internal space of the skin part (A), the hollow molded article being characterized in that, in regard to three planes (plane a, plane b, and plane c, in order from the plane with the greatest area) forming a rectangular parallelepiped that circumscribes the hollow molded article and has the smallest volume, the ratio (α / β) is 1.5-150, where α (cm) is the length of the long side of the plane a and β (cm) is the arithmetic average value of the lengths of the long and short sides of the plane c, and in that conditions 1 and 2 are satisfied.　Condition 1: The porous body has a specific gravity of 0.3 or lower, and at least one porous body is in contact with all or part of the inner surface of the skin part (A) forming a ridge line from either or both of a point and a surface in contact with the plane b to the plane c in top view seen from the plane a.　Condition 2: In all cross-sections perpendicular to the longitudinal direction at positions where the hollow molded article is equally divided into four sections in the longitudinal direction, the area occupied by the hollow part is 50-95%, where 100% is the sum of the area occupied by the hollow part and the area occupied by the porous body.

An aircraft stair system is provided with flexible bumpers configured to prevent abrupt impact of the first structure against the second. Each of the bumpers has a flexible body, a base portion, and a lever arm extending out over the base to create a gap and an open end.

A monument assembly for a vehicle and method for attaching a monument assembly to an aircraft. The monument assembly is capable of being a heavily loaded interior monument enclosure for an aircraft that includes a load bearing structural frame to be attached to a floor of the aircraft. The monument assembly includes a plurality of support members connected together and a set of floor fittings connected to the plurality of support members. The set of floor fittings configured to be connected to the vehicle. A plurality of non-loadbearing panels is connected to the plurality of support members. A load path of the monument assembly is carried by the plurality of support members and not the plurality of panels.

An emergency backup power source for an electrically controlled and motorized window shade unit includes an input power switch, and charge and discharge control circuit, and a supercapacitor bank. During normal operation a nominal voltage level from a main power source of the aircraft is provided directly to the window shade unit through the input power switch, and the supercapacitor bank is charged. The input voltage is monitored by the emergency backup power source, which controls the input power switch. When the input voltage drops below a threshold level, the input power switch is opened and the charge and discharge circuit operates to provide an emergency output voltage to the window shade unit from the supercapacitor bank so that the window shade can be fully opened or closed, as is required.

The present disclosure provides a method of inspection of a surface of an aerodynamic structure. The method includes acquiring image(s) of the surface using image sensor(s) disposed adjacent to the surface, and determining, using the image(s) applied to a model, predicted defect(s) of the surface and corresponding location information. The method further includes controlling, using the location information, the position of tactile sensor(s) disposed adjacent to the surface to acquire dimensioning information for at least a first predicted defect of the predicted defect(s). The method further includes characterizing the first predicted defect using the dimensioning information.

The invention discloses a wide-speed-range variant high-speed aircraft aerodynamic layout, and belongs to the technical field of aircrafts, the wide-speed-range variant high-speed aircraft aerodynamic layout comprises a fuselage, the front end of the fuselage is provided with a drag reduction rod, the rear end is provided with a thrust vectoring nozzle, the left and right sides are symmetrically provided with triangular sweepback wings, and the top is provided with a supporting structure; a high-pressure capturing wing is arranged on the top of the supporting structure. The shock wave separation effect of the drag reduction rod is combined with the high-pressure capture and reflection shock wave induction mechanism of the high-pressure capture wings, so that the aircraft can achieve effective drag reduction and lift augmentation effects in different flight states, and the lift-drag ratio and adaptability of the high-speed blunt cone aircraft are remarkably improved.

Systems, devices, and methods including: an assembly configured to maintain a wing on an aircraft, the assembly comprising: a connector configured to removably connect to a first portion of the wing; and an actuator having a mechanism configured to removably engage a second portion of the wing, the second portion of the wing is configured to be moveable; wherein the actuator is configured to control the second portion of the wing.

Disclosed is a wing assembly for an aircraft. The wing assembly comprises a fixed wing, a wing tip device, a wiring harness and a harness guide. The wing tip device is moveably mounted at a joint at an end of the fixed wing. The wing tip device is moveable about the joint between: (i) a flight configuration for use during flight, and (ii) a ground configuration for use during ground-based operations, in which ground configuration the wing tip device is moved relative to the fixed wing such that the span of the wing is reduced. The wiring harness extends between the fixed wing and the wing tip device. The wiring harness comprises a plurality of conductors and is arranged to transmit electrical power and / or data to the wing tip device. The harness guide is located between the fixed wing and the wing tip device. The harness guide comprises a helical channel configured to receive the wiring harness and guide the wiring harness between the fixed wing and the wing tip device.

Disclosed are devices, systems and methods for mission-adaptable aerial vehicle. In some aspects, a mission-adaptable aerial vehicle includes a configuration having swappable, manipulatable, and interchangeable sections and components connectable by a connection and fastening system able to be modified by an end-user in the field. In some embodiments, a mission-adaptable aerial vehicle can be configured to include a main center body extending along a longitudinal direction, a wing with a lateral cross-sectional airfoil shape, and / or stabilizer and control surface structures with corresponding cross-sectional airfoil shapes.

Disclosed here is an embodiment of a system that can be mounted to an outer panel of a vehicle, such as an aircraft. The disclosed system includes: an antenna; a carrier housing coupled to the antenna, the carrier housing having an upper section configured to receive and hold the antenna, and having a lower section opposite the upper section; and a mounting pedestal. The mounting pedestal includes: a rim having a support platform configured and arranged for coupling to the lower section of the carrier housing; a bottom section opposite the rim, the bottom section configured and arranged for coupling to the outer panel of the vehicle; and a passageway having an opening in the bottom section. The passageway extends from the bottom section toward the rim, and is configured and arranged to accommodate a structural feature of the vehicle that extends from the outer panel of the vehicle.

A system includes a first camera located within a cockpit of an aircraft to track the eye movement of at least one pilot, and a second camera to track an ambient light source. The system includes an electrochromic panel having an array of dimmable cells. The system further includes a controller, equipped with one or more processors responsible for modifying the array of dimmable cells. The controller receives eye tracking data from the first camera associated with a gaze target of the pilot, and light tracking data from the second camera, which includes parameters associated with the ambient light source. Based on the eye tracking and light tracking data, the controller determines a glare location on a surface of the electrochromic panel where the gaze target intersects. The controller may then selectively modify a tint level of specific dimmable cells in the panel to address the determined glare location.

The invention relates to the technical field of aerospace, and particularly provides a liquid metal closed-loop thermal management system and method for a hypersonic aircraft. The system comprises a liquid metal circulation loop which comprises a liquid metal pump used for driving a liquid metal working medium to circularly flow; the heat absorption device is integrated in the liquid metal circulation loop and used for receiving operation heat of the hypersonic flight vehicle and transmitting the operation heat to the liquid metal working medium flowing through the hypersonic flight vehicle. The thermo-acoustic power generation device at least comprises a hot end heat exchanger connected to the heat absorption downstream of the liquid metal circulation loop in series and used for converting part of heat energy of the liquid metal working medium heated after heat absorption into electric energy so as to drive a liquid metal pump. The liquid metal circulation loop further comprises a first diverter valve and a second diverter valve, the first diverter valve is connected into the heat absorption device, and the second diverter valve is connected into the thermo-acoustic power generation device. The dynamic cooperative controller is used for controlling the dynamic cooperative controller based on the flight state parameters of the hypersonic aircraft, the engine thrust requirement and the temperature of the key thermal protection part. The total flow of the liquid metal working medium and flow distribution in the heat absorption device and the thermo-acoustic power generation device are dynamically and cooperatively adjusted by controlling the liquid metal pump, the first diverter valve and the second diverter valve. According to the invention, the compactness of the system can be obviously improved, the heat sink consumption is greatly reduced, and a reliable and efficient heat management solution is provided for long-endurance hypersonic flight.