321 results about "Composite wing" patented technology

A single piece co-cured composite wing is disclosed. The wing has a flying surface and structural members. In one embodiment the structural members may be a plurality of spars. The spars may have various shapes to increase the buckling strength. The spars may be wave shaped, such as a sinusoidal shape. The flying surface and the structural members are co-cured in order to form a single piece, integral wing structure. A process for manufacturing a single piece co-cured wing is also disclosed. The process may include laying out composite sheets for the flying surface of the wing. Then, the composite material of the spars is arranged around a plurality of pressurizable forms. Finally, the composite material is cured in a clamshell frame. A single piece co-cured composite structure is also disclosed that includes an outer skin, and inner skin, a frame, and a stringer.

A single piece co-cured composite wing is disclosed. The wing has a flying surface and structural members. In one embodiment the structural members may be a plurality of spars. The spars may have various shapes to increase the buckling strength. The spars may be wave shaped, such as a sinusoidal shape. The flying surface and the structural members are co-cured in order to form a single piece, integral wing structure. A process for manufacturing a single piece co-cured wing is also disclosed. The process may include laying out composite sheets for the flying surface of the wing. Then, the composite material of the spars is arranged around a plurality of pressurizable forms. Finally, the composite material is cured in a clamshell frame.

An aircraft fuel system architecture which reduces the fleet-wide flammability exposure of the fuel tanks. In one embodiment, the aircraft center fuel tank fuel is cooled at certain times in a flight to reduce its flammability exposure to be similar to that of an unheated conventional metal wing fuel tank. Aircraft fuel tanks that have adjacent heat sources are also insulated to minimize heat flow into the fuel. Fuel tanks that have lower cooling properties, such as composite wing tanks are cooled at certain times during flight such that their temperatures are reduced to be similar to metal wing tanks when the fuel is flammable. A fuel tank that is pressurized relative to outside pressure at altitude having a lower flammability exposure than unpressurized tanks is combined with the cooling of fuel in the tank to reduce the fleet-wide flammability exposure of the fuel tank to be similar to that of an unpressurized metal wing tank. Fuel is cooled by recirculating flow from a tank, passing through a heat exchanger and returning to the tank to be cooled. The heat exchanger is optionally cooled by flow of air from the outside of the aircraft, or by conditioned air from the aircraft environmental control system. The system is controlled by start and stop of fuel flow to the various tanks by means of a system controller. The controller uses sensors in the fuel tanks to command flow only when required to reduce the flammability exposure of the fuel tanks.

A novel design and construction method for an inflatable, rigidizable wing for a terrestrial or planetary flying vehicle. The wing is caused to deploy from an initially packed condition and to assume its functional shape by means of an inflation gas. After inflation, the wing is rigidized by any of several means, such that the inflation gas is no longer required. The composite wing is fabricated from a base reinforcement material, often a fabric, which is coated with a polymer resin that hardens when exposed to a curing mechanism. Several activation mechanisms exist by which to initiate rigidization of such a structure, including elevated temperature, ultraviolet light, and chemical constituents of the inflation gas. The resultant wing has fundamental advantages compared to existing inflatable wings, including improved stiffness, and reduced susceptibility to structural failure in response to puncture.

Embodiments of integral composite panels and joints for composite structures are described In one implementation, an integrated panel spanning substantially the entire wingspan of an aircraft, includes at least a center portion and a pair of outwardly projecting wing portions. The portions may include a skin formed from successive layers or plies of composite material which overlap and offset at the joint between respective sections creating a pad-up area to carry loads between the portions. In a particular implementation, the skin is laid over one or more structural stringers which are transitioned into the joints between sections such as by tapering of the thickness and/or stiffness of the stringer.

Aircraft fuel system reduces flammability exposure of fuel tanks. In one embodiment, aircraft center fuel tank fuel is cooled in flight to reduce flammability exposure to be similar to that of an unheated metal wing tank. Fuel tanks having adjacent heat sources are insulated to minimize heat flow into the fuel. Fuel tanks having lower cooling properties, e.g., composite wing tanks, are cooled during flight so temperatures are reduced to be similar to metal wing tanks. A fuel tank, pressurized relative to outside pressure at altitude, having lower flammability exposure than unpressurized tanks and cooling fuel in the tank reduces flammability exposure of the fuel tank to be similar to that of an unpressurized metal wing tank. Recirculating flow from a tank, passing through heat exchanger, and returning to the tank cools fuel. A controller starts/stops fuel flow to tanks and uses sensors to command flow to reduce flammability.

A wing may include a wing skin, a laminate composite first stringer, a rib, and at least one fastener. A majority of the first stringer may be characterized by a stacked plurality of generally planar plies of reinforcement material structurally joined as a stack to an interior surface of the wing skin and extending generally parallel to the interior surface and a span-wise direction of the wing along a substantial portion of the interior surface. The first stringer may have a generally solid trapezoidal cross section when viewed in a plane that is generally perpendicular to the span-wise direction. The rib may be positioned adjacent the interior surface, and may extend generally perpendicular to the span-wise direction. The trapezoidal cross section may be interfaced with the rib flange. The fastener may extend through the wing skin, the trapezoidal cross section, and the rib flange.

The invention relates to the technical field of aircrafts, in particular to a composite wing unmanned aerial vehicle. The composite wing unmanned aerial vehicle comprises a wing and body combination body, a fixed wing part, a multi-rotor part, an undercarriage, a propeller and a power mechanism. The fixed wing part comprises two fixed wings arranged on the two sides of the wing and body combination body. The multi-rotor part comprises multiple rotor mechanisms symmetrically arranged on the two sides of the wing and body combination body, each rotor mechanism comprises a connecting rod and a rotor, each connecting rod is connected with the corresponding fixed wing, each rotor is horizontally arranged on the corresponding connecting rod, each connecting rod is provided with an installation shaft arranged perpendicular to the axial direction of the connecting rod, and the end portion of each installation shaft is provided with a first slope; and each fixed wing is provided with a connecting hole matched with the corresponding installation shaft, clamping grooves are formed in the hole wall of each connecting hole, telescopic buckles are arranged in the clamping grooves, and each installation shaft is provided with grooves matched with the telescopic buckles. By the adoption of the composite wing unmanned aerial vehicle, the fixed wing part and the multi-rotor part of the unmanned aerial vehicle can be quickly installed and detached without a tool.

The invention discloses a composite wing unmanned plane automatic pilot and a control method employed by the composite wing unmanned plane automatic pilot. Through the automatic pilot and the control guidance algorithm, a multi-shaft and fixed wing coordination control problem is solved, and full-state full-autonomous course flight of a composite wing unmanned plane is realized.

The invention provides a composite wing optimization design method based on an aerodynamic reduced-order model, and belongs to the technical field of aircraft design. Firstly, a structural model and an aerodynamic model of a reference model are established, and modalities of a finite-element model of the reference model are adopted to obtain the corresponding aerodynamic reduced-order model; modalities of the reference model are used as hypothetical modalities to obtain a reduced-order model of the parametric finite-element model; then aeroelastic analysis of the parametric model is carried out; thickness and an angle of composite ply are used as design variables for experiment design, and the design variables obtained by experiment design are used as input and aeroelastic analysis is usedas output to obtain an agent model of the parametric model; and optimization design of a composite wing is carried out on the basis of the agent model. The method improves accuracy of aeroelastic analysis of the wing, improves efficiency of analysis, and is suitable for use in optimization calculation on composite structure design to facilitate obtaining of a composite wing structure with betterperformance.

A method of manufacturing a tiltrotor wing structure including providing a spar mold having a plurality of bores extending from an exterior surface of the mold to an interior surface of the mold, the plurality of bores corresponding to a plurality of primary coordination holes in a spar member, the spar mold having a periphery defined by a top edge, a bottom edge and outboard ends; selecting a plurality of resin impregnated plies to ensure that the plies continuously extend beyond the periphery of the spar mold; laying the plies in the spar mold; curing the plies in the mold to form a cured spar member that extends beyond the periphery of the spar mold; and accurately drilling a plurality of primary coordination holes in the cured spar member in the spar mold using a tool positioned in the plurality of bores.

The invention relates to a sensing and driving integrated composite wing cover, and a test system and application thereof, and belongs to the field of pneumatic design and structure design of airplanes. The cover is formed by splicing N-layer semi-honeycomb structure composite material sensing layers in sequence from the top down, and the longitudinal cross section of the cover is of a honeycomb structure after the sensing layers are spliced; a shape memory alloy drive sandwich layer (3) is arranged between the adjacent semi-honeycomb composite sensing layers; each semi-honeycomb composite sensing layer comprises upper and lower fiber-enhanced composite material layers (1) and an optical fiber sensing sandwich layer (2) arranged between the upper and the lower fiber-enhanced composite layers; the hollow part of the honeycomb structure is filled with silicon rubber (4); an optical fiber grating sensor (8) is arranged in each optical fiber sensing sandwich layer (2); each shape memory alloy drive sandwich layer (3) is provided with a shape memory alloy wire (10). According to the wing cover, the self-sensing and efficient active control requirements of a deformation state of a morphing wing can be met.

The invention provides a method for aeroelastic tailoring of a composite wing. The method for aeroelastic tailoring of the composite wing comprises the steps that aeroelastic tailoring of the composite wing is achieved by means of aeroelasticity optimization, and solution of aeroelasticity optimization is conducted according to a genetic/sensitivity-based hybrid optimization method. The genetic/sensitivity-based hybrid optimization method comprises the steps that (A) a group of individuals are generated randomly according to a genetic algorithm; (B) performance evaluation is conducted on the individuals in the group one by one; (C) according to a user definition, whether sensitivity-based optimization is applied or not is judged; (D) whether the end condition is met or not is judged; (E) if the end condition is not met, a new generation of excellent individuals are generated; if the end condition is met, operation is ended. By the adoption of the method for aeroelastic tailoring of the composite wing and the genetic/sensitivity-based hybrid optimization method, under the condition that multiple requirements for flutter, diffusion, inherent frequency, deformation, aileron efficiency, strength, the flight load, weight and the like are met, tailoring design of the structure is achieved and optimized.

Apparatus and method for joining composite structures of aircrafts are disclosed. An example apparatus disclosed herein includes an inboard composite wing panel having a first end portion configured to be coupled to upper and lower inboard fittings composed of a metallic material and an outboard composite wing panel having a first end portion configured to be coupled to upper and lower outboard fittings composed of a metallic material. A side of body rib has a plate shaped first end portion composed of a metallic material. A fuselage frame is configured to be joined at one end of the upper inboard fitting, and the upper and lower inboard fittings and the upper and lower outboard fittings are coupled to the side of body rib at the plate shaped first end portion.

The invention relates to a method for manufacturing a composite wing box of a small and medium-sized unmanned aerial vehicle to solve the problem that the application of the integral forming of the wing box is limited and to improve the operability and demoulding of the product in the forming process, and improve the complexity of the integral forming tooling and the scope of application of the integral forming of the wing box. The invention solves the problem that the application of the wing box of existing conventional integral molding machine is limited, improves the operability and the demoulding property in the molding process of the product, and effectively improves the complexity of the integral molding tool and the applicable range of the integral molding of the wing box. The invention effectively reduces the quality hidden trouble in the existing integral molding, makes the quality of the integral molding wing box more stable, and effectively improves the reliability of the product. The integral molding tooling is simplified effectively, the molding process is simplified, the product performance of the integral molding is guaranteed to be reliable, the dimension is stable,the design/manufacture integration is realized, the process difficulty is reduced, and the production cycle is shortened.

According to the automatic wind facing method for hovering of the composite wing unmanned aerial vehicle, when the unmanned aerial vehicle vertically takes off and lands in a rotor wing mode, the headdirection of the unmanned aerial vehicle is controlled to be automatically aligned with the incoming flow direction of wind in a fixed-point hovering stage and a heading direction adjusting stage, sothe head faces the wind, and wind resistance of the unmanned aerial vehicle is improved. When the unmanned aerial vehicle just takes off or is ready to vertically land, the unmanned aerial vehicle firstly enters a spot hovering stage, immediately enters a course adjusting stage after the spot hovering time t of the unmanned aerial vehicle, and is judged according to a roll angle instruction, andthe control capability of the yaw rate of the unmanned aerial vehicle is improved by combining the horizontal component force of the rotor tension, the rotor rotation reaction torque difference and afixed wing rudder; and the heading of the nose is adjusted to realize automatic wind alignment.