2062results about "Wings" patented technology

A method of and apparatus for providing tunable gas injection in a plasma processing system (10, 10′). The apparatus includes a gas injection manifold (50) having a pressurizable plenum (150) and an array of adjustable nozzle units (250), or an array of non-adjustable nozzles (502, 602), through which gas from the plenum can flow into the interior region (40) of a plasma reactor chamber (14) capable of containing a plasma (41). The adjustable nozzle units include a nozzle plug (160) arranged within a nozzle bore (166). A variety of different nozzle units are disclosed. The nozzle plugs are axially translatable to adjust the flow of gas therethrough. In one embodiment, the nozzle plugs are attached to a plug plate (154), which is displacable relative to an injection plate (124) via displacement actuators (170) connecting the two plates. The displacement actuators are controlled by a displacement actuator control unit (180), which is in electronic communication with a plasma processing system control unit (80). The gas flow into the chamber interior region is preferably controlled by monitoring the pressure in the plenum and in the chamber and adjusting the nozzle units accordingly. Where the nozzle units are not adjustable, a portion of the nozzles are sized to a first flow condition, and another portion of the nozzles are sized to a second flow condition.

An advanced internal combustion-electric hybrid airplane, having at least double the flight range and flight duration than a conventional equivalent airplane, while using the same amount of any desirable fuel. This is achieved by using 2-3× smaller and ultra-lightweight engine for cruising, and ultra-lightweight electric motor powered by lithium batteries during take-off and climbing. The electric motor becomes a generator during cruising and descent, recharging said batteries. The airplane has also temporary silent electric stealth capability and added safety by the electric back-up power. Due to its high efficiency, the operational cost is substantially reduced. Additional features include highly advanced, minimum drag and weight airframe.

This patent provides for a method and apparatus for mitigating the formation of concentrated wake vortex structures generated from lifting or thrust-generating bodies and maneuvering control surfaces wherein the use of contour surface geometries promotes vortex-mixing of high and low flow fluids. The method and apparatus can be combined with various drag reduction techniques, such as the use of riblets of various types and / or compliant surfaces (passive and active). Such combinations form unique structures for various fluid dynamic control applications to suppress transiently growing forms of boundary layer disturbances in a manner that significantly improves performance and has improved control dynamics.

A single dielectric barrier aerodynamic plasma actuator apparatus based on the dielectric barrier discharge phenomenon is disclosed and suggested for application to aerodynamic uses for drag reduction, stall elimination and airfoil efficiency improvement. In the plasma actuator apparatus non-uniform in time and space, partially ionized gasses are generated by one or more electrode pairs each having one electrically encapsulated electrode and one air stream exposed electrode and energization by a high-voltage alternating current waveform. The influence of electrical waveform variation, electrode polarity, electrode size and electrode shape on the achieved plasma are considered along with theoretical verification of achieved results. Light output, generated thrust, ionizing current waveform and magnitude and other variables are considered. Misconceptions prevailing in the present day plasma generation art are addressed and are believed-to-be corrected. The influence of electrostatic shielding effects of the developed plasma on the applied electric field are also considered.

Systems and methods are provided for swapping the battery on an unmanned aerial vehicle (UAV) while providing continuous power to at least one system on the UAV. The UAV may be able to identify and land on an energy provision station autonomously. The UAV may take off and / or land on the energy provision station. The UAV may communicate with the energy provision station. The energy provision station may store and charge batteries for use on a UAV. The UAV and / or the energy provision station may have a backup energy source to provide continuous power to the UAV.

A vertical takeoff and landing (VTOL) air vehicle disclosed. The air vehicle can be manned or unmanned. In one embodiment, the air vehicle includes two shrouded propellers, a fuselage and a gyroscopic stabilization disk installed in the fuselage. The gyroscopic stabilization disk can be configured to provide sufficient angular momentum, by sufficient mass and / or sufficient angular velocity, such that the air vehicle is gyroscopically stabilized during various phases of flight. In one embodiment the fuselage is fixedly attached to the shrouded propellers. In another embodiment, the shrouded propellers are pivotably mounted to the fuselage.

A vertical takeoff and landing aircraft, using for vertical lift and lateral thrust a redundant plurality of essentially similar electrically-powered and electronically-controlled thrust units mounted in a mechanically static or fixed fashion relative to one another in a substantially horizontal plane. The thrust units are situated in this planar array in aerodynamically approximate pairs, such that a complete failure of a single thrust unit would not substantially compromise the ability of the aircraft to maintain flight.

A stiffened skin panel of an aircraft that includes an aircraft skin panel of composite material and a composite stringer consolidated with the aircraft skin panel. The stringer has a pair of stringer sides, each extending from a stringer top to a stringer leg. The stringer sides may each extend through a wide radius, smooth continuous curve to a stringer leg. The wide radius, smooth continuous curve may be proximate the base of the stringer side in a region where the stringer side transitions to the stringer leg.

An integrated system for automatic formation flight control of multiple vehicles such as aircraft, helicopters, or space platforms. The system is used to control multiple aircraft in a pre-determined flight formation and provide positive identification, control and discrete communications between any number of aircraft in order to prevent mid-air collisions between aircraft in formation flight. The system includes a processor located on the aircraft to enable communications, position computations and control messaging between any number of aircraft in formation flight, a communications transceiver located on the aircraft that provides discrete communication links to any number of aircraft or aircraft in formation flight, an autopilot, an autothrottle and a display. On aircraft, the system will display the current formation and aircraft relative positions under control by the system. The system may also be used to circumvent control of a hijacked aircraft through control system intervention provided through this invention.

According to a non-limiting embodiment of the present invention, an appliqué coating is provided for a substrate. The appliqué coating includes a metal foil and a first polymer film underlying the metal foil. A topcoat overlying the metal foil may be provided as a layer of paint or as a second polymer film. An adhesive, such as a pressure sensitive adhesive, underlying the first polymer film may be provided to affix the appliqué coating to the substrate. If desired, fibers may be dispersed throughout the second polymer film to provide anti-static properties. Also, an ink layer may be provided between the metal foil and the second polymer film, if desired, for aesthetic and / or anti-static purposes.

A method of forming a structural airframe component for an aircraft and an airframe structural component The method includes placing at least two components (1,2) in abutting relationship with each other and joining them together by friction stir butt welding (3), and the structural airframe component comprises a component manufactured according to the method of the invention.

A hybrid propulsion aircraft is described having a distributed electric propulsion system. The distributed electric propulsion system includes a turbo shaft engine that drives one or more generators through a gearbox. The generator provides AC power to a plurality of ducted fans (each being driven by an electric motor). The ducted fans may be integrated with the hybrid propulsion aircraft's wings. The wings can be pivotally attached to the fuselage, thereby allowing for vertical take-off and landing. The design of the hybrid propulsion aircraft mitigates undesirable transient behavior traditionally encountered during a transition from vertical flight to horizontal flight. Moreover, the hybrid propulsion aircraft offers a fast, constant-altitude transition, without requiring a climb or dive to transition. It also offers increased efficiency in both hover and forward flight versus other VTOL aircraft and a higher forward max speed than traditional rotorcraft.

An overlay of patterned metal foil above a substrate to be protected supports development of localized coronas after a lightning strike. The localized coronas transport energy of a lightning strike above the substrate's surface with limited removal of metal foil from the lightning strike attachment point. A first polymer film underlies patterned metal foil. A topcoat overlies the patterned metal foil. An adhesive, underlying the first polymer film affixes the lightning diverter overlay to the substrate. If desired, semiconductor particulates may be dispersed throughout the second polymer film to contribute to instantaneous generation of localized coronas. Also, an ink layer may be provided between the patterned metal foil and the topcoat or on the exterior, if desired, for aesthetic and / or anti-static purposes.

A wing load alleviation system and method for alleviating the lift-inducing structural-bending force (i.e., moment) experienced by each of the wings of an aircraft. The apparatus includes a deployable panel and an actuator mounted in each wing. The actuators are responsive to a command generator. The actuator is mounted inside the wing and the panel is mounted flush with an outer surface of its respective wing. Each panel can be moved between a retracted position, where it has no affect on airflow moving over the wing, to a deployed position in which it deflects air off of the wing. Each panel is preferably located at a span-wise location at least about halfway along the length of the wing toward the wing tip, and more preferably at least in part outboardly of the outboard-most trailing edge device in the wing. The apparatus effectively shifts the lift-inducing structural-bending forces experienced by the wing more inboard towards the fuselage.

Integrally stiffened and formed, load carrying structures comprising a plurality of elongated thin-walled tubes placed co-extensively in a complementary side-by-side fashion which together form a hollow structure having a desired external contour. Integral skins forming the external and internal surfaces of the structure cooperatively therewith. The structure can be formed with an underlying internal support member spanning the interior of the load carrying structure, thereby connecting opposite sides of the structure together. Also, each of the tubes are wound with fibers in controlled orientations generally paralleling the direction of the loads applied to the tubes to optimize the strength to weight ratio of the tubes.

The present invention relates to a method of controlling an aircraft, missile, munition or ground vehicle with plasma actuators, and more particularly to controlling fluid flow across their surfaces or other surfaces, which would benefit from such a method. The method includes the design of an aerodynamic plasma actuator for the purpose of controlling airflow separation over a control surface of a aircraft, missile, or a ground vehicle, and more particularly to the method of determining a modulation frequency for the plasma actuator for the purpose of fluid flow control over these vehicles. The various embodiments provide the steps to increase the efficiency of aircraft, missiles, munitions and ground vehicles. The method of flow control provides a means for reducing aircraft, missile's, munition's and ground vehicle's power requirements. These methods also provide alternate means for aerodynamic control using low-power hingeless plasma actuator devices.

Systems and methods for providing variable geometry winglets to an aircraft are disclosed. In one embodiment, a winglet includes a base portion configured to attach to a wing. The winglet further includes a body portion. In turn, the body portion includes at least one of a deflectable control surface, a shape memory alloy (SMA) bending plate, and a SMA torque tube. The base portion is configured to attach to the wing such that the body portion projects at an upward angle from the wing.

A method and apparatus for manufacturing wings includes a fixture that holds wing panels for drilling and edge trimming by accurate numerically controlled machine tools using original numerical part definition records, utilizing spatial relationships between key features of detail parts or subassemblies as represented by coordination features machined into the parts and subassemblies, thereby making the parts and subassemblies intrinsically determinant of the dimensions and contour of the wing. Spars are attached to the wing panel using the coordination holes to locate the spars accurately on the panel in accordance with the original engineering design, and in-spar ribs are attached to rib posts on the spar using accurately drilled coordination holes in the ends of the rib and in the rib post. The wing contour is determined by the configuration of the spars and ribs rather than by any conventional hard tooling which determines the wing contour in conventional processes.

The present invention is directed to the manufacture of and the use of an aerodynamic flow control device having a compact array of a plurality of fluidic actuators in planar, curved, circular and annular configurations. The compact array of fluidic actuators of the invention may be designed to produce oscillating or pulsed jets at the exit ports with frequencies in the range of 1-22 kHz. They may be integrally manufactured along with the wing sections, flaps, tail and rudder of airplane, the inlet or exit geometries of a jet engine. When supplied with a source of fluid such as air, these arrays of actuators produce a set of fluid jets of random phase of high velocity and influence the main stream of air over the subject surface. The beneficial effects of modifying flow using the present invention include increased lift, reduced drag, improved performance and noise reduction in jet engines.

A compound aircraft features variable incidence wings, a thruster, a rotor and a clutch connecting the engines to the rotor. Upon reaching a suitable forward speed, the control system of the compound aircraft unloads the rotor and disengages the clutch, disconnecting the rotor from the engines. The control system configures the cyclic and collective pitch effectors to cause the rotor to autorotate. The control system selects an autorotation rotational speed that is adequately rapid to prevent excessive deformation of the rotor blades due to aerodynamic forces acting on the blade and that is adequately slow to prevent excessive compressibility effects.

An unmanned aerial vehicle, comprising: a fuselage having a first side board and a second side board spaced apart and connected by at least one transverse board; the first side board, the second side board, and the at least one transverse board being printed circuit boards; at least one of the first side board, the second side board, and the at least one transverse board having formed and mounted thereon conductive traces and at least one component, respectively, for controlling and monitoring the unmanned aerial vehicle; first and second wings mounted to the fuselage; and, a tail mounted to the fuselage.

An aerial vehicle capable of convertible flight from hover to linear flight includes a body having a longitudinal body axis, a plurality of forward wings, a plurality of aft wings, at least one motor, and at least three aerodynamic propulsors driven by the at least one motor. Each forward wing extends a forward wing plane. Each aft wing extends from an aft wing plane. The aerodynamic propulsors are mounted longitudinally between the plurality of forward wings and plurality of aft wings.

A system and method for controlling boundary layer flow over an aircraft wing are provided. The system includes at least one wing element, and a plurality of ports defined in the wing element and in fluid communication with one another. The system also includes at least one fluidic device operable to continuously ingest the fluid through at least one of the ports and eject the fluid out of at least one other port to control boundary layer flow of the fluid over the wing element.

A sandwich beam including in one example first and second spaced walls, a core configured to maintain a predetermined spacing between the walls when the core is filled with pressurized gas and to resist shear when the beam is loaded in bending and a port for filling the core with gas biasing both walls in tension. The tension tends to increase in the second wall and decrease and cause a compression load in the first wall in response to a sufficiently large applied bending load. A compression element is fixed only with respect to the first wall and is configured (a) to support the compression load so that the beam is stronger at a given gas pressure and (b) to flex sufficiently to allow the beam to be rolled up when the gas is emptied from the core via the port.

An aircraft comprises a wing tip device, for example a winglet, a raked-tip device, a wing tip fence or a planar wing extension, mounted in the region of the tip of a wing on the aircraft. The wing tip device is rotatably moveable between a first position and a second position, in which the upward lift produced by the wing or the wing tip device is reduced. During flight, the bending moment at the root of the aircraft wing therefore changes in dependence on the position of the wing tip device. The maximum bending moment in the aircraft wing sustained during high-load conditions is thereby reduced, allowing the structural mass of the aircraft to be reduced.

The present invention relates to a device for locking the pitch of the blades of a main rotor of a rotary-wing aircraft in which each blade is firstly rotated about an axis of rotation of the rotor by a rotor mast and is secondly constrained to pivot about a longitudinal pitch axis of the blade together with a pitch lever which is controlled by a pitch link connected to a rotary plate rotating with the rotor mast and belonging to a cyclic swash plate mechanism in which the rotary plate is rotatably mounted on a non-rotary plate capable of sliding axially along said rotor mast and of tilting in any direction relative to the rotor mast under the drive of at least three servo-controls each comprising a body fixed on a support secured to the aircraft and a rod having a free end secured to the non-rotary plate. The device includes immobilization means for holding the rod of each of the servo-controls relative to the corresponding body so as to lock the blades in a predetermined pitch position.