721results about "Fuselages" patented technology

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 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.

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.

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.

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.

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 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.

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.

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.

A backscatter x-ray system has an x-ray device, a movable platform, and a processor. The x-ray device includes an x-ray emitter and an x-ray detector. The movable platform positions the x-ray device relative to a frontside of a structure. The processor is communicatively couplable to the x-ray device. The x-ray emitter emits x-rays that penetrate the structure from the frontside. The x-ray detector detects a backscatter of the x-rays reflected from the structure. The processor generates x-ray images of the structure based on the backscatter, and analyzes the x-ray images in a manner facilitating at least one of manufacturing and inspection of the structure.

An unmanned aerial vehicle (UAV) may detect a risk of collision with one or more objects in an airspace serviced by a mobile edge computing (MEC) node. The MEC node may provide an edge detect and avoid (edge-DAA) function for use in the airspace. The UAV may determine a first resolution advisory (RA) to be acted on in order to avoid the collision with the one or more objects based on a local DAA function within the UAV. The UAV may receive, from the MEC node, a second RA to be acted on in order to avoid the collision with the one or more objects based on the edge-DAA function. If the second RA can be acted onto avoid the collision with the one or more objects, the UAV may act on the second RA and may send a message to the MEC node with an acknowledgement.

An airframe for a projectile having a cylinder having a bottom spring and a piston within the cylinder is provided. The cylinder has a locking assembly having a housing. A leaf spring is in the housing and coupled thereto. The housing has a limiter where the leaf spring operatively couples with the limiter and the limiter biases the leaf spring. The piston is moveable relative to the cylinder between a retracted configuration and expanded configurations. The piston has a housing where the piston moves between the retracted configuration and the expanded configurations when the piston housing receives a fluid. The piston has a groove that receives the leaf spring. The groove can align with the leaf spring in one expanded configuration where the piston groove receives the leaf spring thereby locking the piston with cylinder. In another expanded configuration, the leaf spring is outside the piston groove.

A compliant mount for an aircraft, for attaching equipment (such as a latching unit, for example) to a structural member of the aircraft, the compliant mount comprising: a mounting member for attachment to the structural member of the aircraft; a receptacle for supporting the equipment, the receptacle being coupled to the mounting member along a longitudinal axis; and one or more spring elements arranged to resiliently bias the receptacle into alignment with the longitudinal axis and to allow angular movement of the receptacle away from the longitudinal axis. Also provided is an aircraft equipment assembly comprising such a compliant mount, and aircraft equipment supported by the receptacle, and an aircraft having one or more such equipment assemblies. The aircraft may be a folding wingtip aircraft and the equipment assembly may be part of a folding wingtip actuation mechanism.

Disclosed are specialized airframe structures for mission-adaptable unmanned aerial vehicles (UAVs) that are fabricated by rapid, low-cost additive manufacturing techniques. In some embodiments, an integral fastening system for a mission-adaptable aerial vehicle includes a first airframe section, a second airframe section, and a bayonet mount system that comprises one or more passageways produced on a first end region of the first airframe section to include a first channel and second channel, and one or more protrusions produced on a second end region of the second airframe section, such that, when the first and second airframe sections being assembled, the one or more protrusions are first aligned with an entry region of the first channel to be translated and then aligned with the second channel of the one or more passageways to be rotated to securely connect the first and second airframe sections together.

A fuselage of an aircraft includes a front portion and a rear portion, with at least one opening in the rear portion, with the opening being closeable by a structural panel fully removeable from the opening, wherein a locking structure for the structural panel is in the opening and configured to transfer loads between a primary structure of the fuselage and the structural panel. A method is disclosed for closing an opening of a fuselage, and an aircraft is disclosed having such a fuselage.

The invention discloses an aircraft radome lightning protection structure and a processing method thereof. The lightning protection structure comprises a radome body, and a first metal shunting strip, a second metal shunting strip and a medium type shunting strip are arranged on the outer side face of the radome body. Lightning protection shunting strips are reasonably arranged on the outer surface of the radome, a medium type shunting structure and a metal shunting structure are designed in a mixed mode, the distance and section parameters of the shunting strips are accurately designed, the lightning protection structure of the aircraft radome is machined, lightning current can be fully and efficiently introduced into a main lightning path of an aircraft body, and when the aircraft body is struck by lightning, the lightning protection structure can be used for lightning protection. The radome has the advantages that lightning can be effectively attracted, lightning current is attached to the shunt strips, the radome body cannot be broken down, the shunt strips can safely bear 200KA lightning current and cannot be damaged, the lightning protection design requirement is met, the loss of the wave transmission rate of the radome body is low, the lightning protection effect is remarkable, and the safety and reliability of an aircraft are ensured; the method can be widely applied to various aircraft radomes.

An aircraft propulsion unit comprising a motor, a system for supplying fuel to the motor, a fairing inside which the motor and at least part of the fuel supplying system are positioned, and at least one pressure relief device configured to establish communication between the inner and outer zones of the fairing when the pressure in the inner zone is greater than or equal to a given threshold or when a leak is detected. Also an aircraft comprising at least one such propulsion unit.

The present application relates to an electron emission cooling thermal protection material and a preparation method and application thereof, and belongs to the technical field of composite materials, and solves the problem that the existing electron emission cooling thermal protection material cannot consider low emission work and high conductivity. The present application provides a preparation method of an electron emission cooling thermal protection material, ceramic matrix powder and conductive reinforcing phase material are homogenized and mixed, then densification forming and high reduction heat treatment are performed to obtain a composite material with a continuous conductive network. The electron emission cooling thermal protection material prepared by the preparation method has excellent comprehensive performance, the emission work is not higher than 3.5 eV, the conductivity can be up to 4.72*10 ‑2 S / m, and is not prone to performance degradation in a high-temperature oxidation environment. After heating and oxidation treatment of the material in air at 1600 DEG C, the emission work decreases by not more than 15% and the conductivity decreases by not more than 10% after 5 minutes of oxidation treatment, and the electron emission cooling can be realized efficiently and sustainably in a severe environment.

A snap-on support assembly has a snap-on base and at least one ring post. The snap-on base includes a base bottom portion and a base top portion. The base bottom portion has a spaced-apart pair of bottom portion attachment arms configured for snap-on attachment to a bar member. The base top portion comprises a monument co-molded with the base bottom portion and having a monument top end having a monument threaded bore. The ring post has a ring for securing one or more elongate elements to the ring post. The ring post is configured to be attached to the monument top end via a threaded element engaging the monument threaded bore.

This invention, relates to a body (2), a door (3) on the body (2), which allows the user access to the body (2) or prevents access to the body (2), at least one hinge (4) on the door (3), allowing the door (3) to be supported on the body (2) and allowing the door (3) to be opened by pivoting around the axis on which it is supported on the body (2), at least one trigger (5) to trigger the door (3), a first hinge channel (6) located on the hinge (4) and having a through opening, at least one hinge pin (7) located on the hinge (4) and extending along its length, a closed position (I) in which the door (3) encloses the body (2) in a form-fitting manner, the door (3) being closed on the body (2), an open position (II) in which the hinge pin (7) contacts the first hinge channel (6) within the first hinge channel (6), the door (3) being moved outwardly from the body (2) by means of rotation of the hinge (4) from the closed position (I) to allow the user to access the body (2), an emergency trigger position (III) in which the hinge pin (7) protrudes from the first hinge channel (6) and there is a distance between the first hinge channel (6) and the hinge pin (7), in which the trigger (5) is triggered in the event that the user needs to be evacuated from the body (2), an evacuation position (IV) in which the door (3) is brought from the emergency trigger position (III) so that the hinge (4) and the body (2) move independently of each other, and the door (3) and hinge (4) are pushed out of the body (2), allowing the user to evacuate from the body (2), at least one latch (8) located on the door (3), which allows the door (3) to be locked, at least one latch holder (9) located on the body (2), which enables the door (3) to be locked in the closed position (I) by being placed in the latch (8), and which enables the door (3) to be opened in the open position (II) when it is brought out of the latch (8) so as to keep a distance between it and the latch (8).

A system for monitoring a pressure change within at least one compartment of an aircraft and operating a pyrotechnic device associated with a door latch, includes a securing mechanism arranged in an aircraft door that includes a pyrotechnic actuator and a latch bolt. The system includes the pyrotechnic actuator configured to move the latch bolt when the pyrotechnic actuator is actuated. The system also includes a pressure sensor configured to provide a pressure signal. The system also includes a controller configured to determine a pressure drop representing a decompression event. The system also includes an output driver configured to generate an output signal to actuate the pyrotechnic actuator when the controller determines the occurrence of a decompression event.

The application provides a kind of co-woven three-dimensional interpenetrating network reinforced aerogel composite material and its preparation method and application.The composite material is composed of reinforced phase and matrix phase, the reinforced phase is three-dimensional interpenetrating network fiber felt prepared by integrated weaving or needling process, containing hollow fiber and solid quartz fiber, which are interpenetrated and entangled in three-dimensional space to form a continuous network skeleton;The matrix phase is silica-based aerogel filled and coated network pores.The preparation method includes the steps of forming, sol impregnation, gelation and aging, supercritical drying and other steps of co-woven preform, and can also include hydrophobic post-processing.The material has the characteristics of light weight, high strength, ultra-low thermal conductivity and excellent wave-transparent performance, and is suitable for hypersonic aircraft radome, antenna window, radar fairing and satellite communication wave-transparent components, which can meet the requirements of load bearing, heat insulation and wave-transparent integration.

This invention discloses a high-modulus, high-toughness bismaleimide resin composition, a carbon fiber prepreg, and a preparation method thereof, belonging to the field of composite material technology. The bismaleimide resin composition of this invention comprises the following raw materials in parts by weight: 40-80 parts bismaleimide resin, 10-20 parts o-diallyl bisphenol A, and 5-10 parts a powder toughening agent containing phenolphthalein side chains. The carbon fiber prepreg of this invention is made from the above resin composition and carbon fibers, with thermoplastic polyimide felt interlayered to form an integrated structure. This allows the composite material prepared from the carbon fiber prepreg to maintain high modulus and high temperature resistance while significantly improving its toughness, impact damage tolerance, and interlayer delamination resistance. Furthermore, the prepreg of this invention has good processability and is suitable for manufacturing high-performance complex components.

The invention provides a trailing edge variable camber airfoil composed of a bistable composite material structure and a design and control method of the trailing edge variable camber airfoil, and relates to the technical field of aircraft design and aerodynamics. Wherein the trailing edge variable camber airfoil sequentially comprises a leading edge section, an airfoil middle section and a trailing edge section in the chordwise direction, a front beam is arranged between the leading edge section and the airfoil middle section, the trailing edge section is fixed to an airfoil rear beam, and the trailing edge section is a core function section and is formed by arranging multiple layers of glass / carbon hybrid bistable composite material structures from top to bottom in the wingspan direction; each glass / carbon hybrid bistable composite material structure has a first stable state and a second stable state which correspond to a first flight state and a second flight state of the airfoil profile respectively, one end of each glass / carbon hybrid bistable composite material structure is fixedly connected to the airfoil profile rear beam, and the other end of each glass / carbon hybrid bistable composite material structure is fixed into a whole through glue joint. According to the invention, the steady-state characteristic of the bistable composite material is utilized, so that non-driven conversion between two stable bending states is realized.

A handling system for imparting a desired longitudinal conformation, for example a twist or skew, into a longitudinal component, for example a composite stringer for an aircraft wing, comprises a rigi