310results about "Sandwich constructions" 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 system and method for forming structural assemblies with 3-D woven joint pre-forms. The method of the present invention forms complex structural assemblies with pre-formed structures. Adhesive is applied between the preformed structures and uncured 3-D woven textile pre-forms. Then together the preformed structures and uncured resin impregnated 3-D woven textile are cured with heat and/or pressure to form the larger complex structural assemblies.

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 de-icing/anti-icing system including at least two conductive structures embedded in an article that includes an outer surface designed as an aerodynamic surface. At least one of the conductive structures is arranged adjacent the outer surface. A control unit adapted to control the energy supply to the conductive structures for generating heat to the outer surface. A first of the conductive structures includes a first conductive nano structure and a second of the conductive structures includes a second conductive nano structure. A conductive property of the first of the conductive structures differs from a conductive property of the second of the conductive structures. The first conductive nano structure serves as a heating conductor and the second conductive nano structure serves as a heating element. The first and second conductive nano structures are embedded in a common plane of a resin layer forming the outer surface.

An aircraft wing has a plurality of wing segments mounted on a wing spar by joints that allow relative movement between the spar and the wing segments. Tuning of coefficients of thermal expansion is employed to reduce induced stresses from changes in temperature.

The invention provides a vertical take-off and landing fixed-wing unmanned aerial vehicle, which belongs to the technical field of aircrafts. The purpose is to solving problems that the take-off and landing condition of a fixed-wing aircraft is harsh, the loitering speed of a rotor craft is low and the endurance of the rotor craft is short in the prior art are solved. The vertical take-off and landing fixed-wing unmanned aerial vehicle comprises a fixed-wing structure system, a fixed-wing power system, a rotor structure and power system, a rotor structure supporting piece as well as a flight control system and an avionic system, wherein the fixed-wing power system comprises a single cylinder two-stroke engine and a propeller A, which are fixed at the tail end of an aerial vehicle body; a power output end of the engine is connected with the propeller A; the rotor structure and power system is arranged on the fixed-wing structure system; the rotor structure and power system comprises a motor, an electronic speed controller and propellers B; the motor is electrically connected with the electronic speed controller; an output end of the motor is connected with the propellers B; four propellers B are respectively arranged on the upper end surfaces of two supporting rods of the fixed-wing structure system; the propeller A is perpendicular to the propellers B; the rotor structure supporting piece is fixed at the upper end of a wing of the fixed-wing structure system and used for mounting a rotor power system; and the flight control system and the avionic system are used for controlling the flight path and posture of the unmanned aerial vehicle.

An airfoil includes at least one precured composite spar having a web and at least one flange integrated with an end of the web. A precured composite skin is attached to the spar by adhesive bonding the skin to the flange.

A composite material made of multiple filament bands is disclosed. The filament bands are wound to create a composite material. The filament bands may be impregnated with a solid or semi-solid resin. The filament bands are wound in multiple orientations to form a sheet of composite material. Additionally, the composite material may be made to have a varying number of layer and fiber orientations throughout the sheet of composite material. In another embodiment, a composite component requiring multiple layers and fiber orientations may be substantially manufactured during a filament winding process.

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 a co-curing process for the joining of composite structures. A method of fabricating a composite assembly (200) may include providing a first laminate (206) and a second laminate (216) respectively formed of first and second composite plies (204, 214), and having a respective first and second cured section (208, 218) and a respective first and second uncured section (210, 220). The method (200) may further include interleaving the first composite plies (204) in the first uncured section (210) with the second composite plies (214) in the second uncured section (220) to form an interfacial region (256). The method may additionally include curing the interfacial region (256) to join the first laminate (206) to the second laminate (216) and form a unitized composite assembly (200).

The invention discloses a trailing edge (2) of an aircraft aerodynamic surface, which comprises an upper skin (6) and a lower skin (7), being the upper and lower skins (6, 7) made of composite material. In the trailing edge (2) a lightweight part (1) is assembled at an end section (13) of the upper and lower skins (6, 7). The lightweight part (1) is covered by a metallic sheet (8), and has a “V”-shaped cross section with arms and peak, where the peak of the “V”-shaped cross section has a rounded shape and is located opposite the end section (13) of the upper and lower skins (6, 7).

A dual-skin structure comprises a first skin, a second skin and an intermediate structure. The intermediate structure comprises a plurality of first contact portions connected to an interior surface of the first skin, a plurality of second contact portions connected to an interior surface of the second skin and a plurality of interconnecting web portions integral to the first and second contact portions and extending between ones of the first and second contact portions to form an internal supporting structure alternating between the interior surfaces of the first and second skins.

A composite laminate has a primary axis of loading and comprises a plurality resin plies each reinforced with unidirectional fibers. The laminate includes cross-plies with fiber orientations optimized to resist bending and torsional loads along the primary axis of loading.

An aircraft wing cover (10) comprising an integrally stiffened sandwich panel including first and second skins (11, 12) sandwiching a core layer (13), wherein at least one of the skins has a plurality of spanwise extending integral regions of increased thickness (18) accommodated by corresponding regions (19) of decreased thickness in the core. Also, a method of fabricating an aircraft wing cover, and a method of designing an aircraft wing.

The invention relates to a deformable aerofoil, which solves the problems of complex structure, and large mass of the aerofoil of the existing deformable aircraft and the problems of low flying efficiency and poor air-tightness of the aircraft. The invention consists of a honeycomb sandwich matrix (1) and a cover (2); the honeycomb sandwich matrix (1) is an integral honeycomb structure of a prefabricated aerofoil shape; the cover (2) is coated on the outer surface of the honeycomb sandwich matrix (1). The aerofoil of the invention employs the honeycomb sandwich structure which has higher specific strength and specific rigidity, corrosion and fatigue resistance, excellent electric insulating performance and electromagnetic wave transmission characteristic. The honeycomb sandwich structure greatly reduces the mass of the aerofoil, and improves the flying efficiency of the aircraft. The integral honeycomb structure of a prefabricated aerofoil shape of the honeycomb sandwich matrix leads the aircraft to have better air-tightness and is a simple. The design of the honeycomb sandwich structure leads the aerofoil to be suitable for various shape changes.