784 results about "Aileron" patented technology

A method of reducing the bending moment effect of wind gust loads acting on the wing of an aircraft involves adjusting the aerodynamic configuration of the wing so as to alter the distribution of lift generated by the wing during phases of flight in which critical wind gusts are expected to occur. Particularly, during climb and descent phases of flight below cruise altitude, the lift generated by outboard portions of the wings is reduced while the lift generated by inboard portions of the wings is increased. Thereby, the 1 g basis load acting on the outboard portions of the wings is reduced, and consequently the total load applied to the outboard portions of the wings, resulting from the 1 g basis load plus the additional wind gust load, is correspondingly reduced. This leads to a reduction of the bending moments effective on the wings, and of any rolling moment effective on the aircraft. The required adjustment of the lift distribution is preferably achieved by deflecting the ailerons of both wings symmetrically upward and/or deflecting the flaps of both wings symmetrically downward during climb and descent. The adjustment of the wing configuration is carried out dependent only on flight parameters such as the altitude, speed and gross weight, and does not require rapid sensing of the occurrence of a wind gust and rapid actuation of control surfaces to try to instantaneously counteract a wind gust as it occurs.

A vertical axis wind turbine includes an upstanding support structure, a plurality of generators disposed on the support structure, a central shaft in rotatable communication with the generators and positioned along a central axis of the vertical-axis wind turbine, a plurality of struts extending from the central shaft, and a plurality of blades, each blade positioned at an end of a corresponding strut and oriented substantially vertically. The vertical axis wind turbine optionally includes strut ailerons, blade extension elements, or blade ailerons to increase the efficiency and duty cycle of the wind turbine.

The composite rotating fixed-wing aircraft consists of the coaxial counter-paddle, reverse gear, engine output shaft, engine with a starter-generator, wings, tail blades, landing device, culvert, steering gear, fairing, fuselage, motor, motor drive shaft, wing control unit of small angle of attack, aileron rudder surface, and tailplane rudder surface. Co-axial counter-blade is located in the upper part of the aircraft, which connects with the engine output shaft; the reverse gear is installed between the co-axial counter-propellers; the motor connects with motor drive shaft powered by the starter-generator of the engine; the wings are located on both sides of aircraft and connect with the fuselage; the tail blades are located at the aircraft tail, which are installed behinde motor drive shaft; the landing device is located under the lower part of the fuselage and fixed with it; the culvert connects with the landing device; the fairing is installed in the culvert and connects with the steering gear; the wing control unit of small angle of attack is installed on the wing. The aircraft design method has six steps with strict scientific idea; this invention has a wide range of practical value and application prospect.

The present invention provides a flight control system and method for various models of aircraft, including thin-winged, long range, transonic and low supersonic aircraft, and may be beneficial on supersonic and hypersonic aircraft as well. One embodiment of the flight control system comprises a plurality of lifters for generating lift or drag, the plurality of lifters mechanically associated with a lower surface of each wing in a pair of wings. During operation, the lifters may be selectively deployed alone or in combination with other flight control devices, such as ailerons and spoilers, to provide various operative benefits such as improved roll control, speed braking functionality, and maneuver load alleviation.

The invention provides a tiltrotor controlled by a double-propeller vertical duct. The tiltrotor, using the design of parallel twin-rotor and conventional aerodynamic configuration, consists of a fuselage, straight wings, rotors, a nacelle, an aileron, a vertical empennage, a rudder, an elevator, a horizontal empennage, a duct end cover, a vertical duct, an undercarriage, a power and decelerationsystem, a propeller power input shaft, a lower propeller, an upper propeller, louver-type slipstream sheets, a propeller gearing and collective-pitch controlling device, a propeller gear box supporting structure, a duct end cover driving device, a duct end cover motion slide rail and a slipstream sheet controlling device. Compared with the cyclic pitch control mode, the tiltrotor of the inventionsimplifies the operations during the vertical flight and the flight conversion and improves the reliability by using the structure of the double-propeller vertical duct to control the vertical flightand the conversion of the flight mode, and the tiltrotor has the advantages of long control force arm, high control efficiency and simple structure. Therefore, the tiltrotor constitutes a novel aircraft with significant development potential and promising future.

The invention relates to an aileron rotary folding and unfolding type flapping wing device, which includes aileron rotary folding and unfolding type flapping wings and a related drive device, wherein the wings are divided into main wings and ailerons along the wingspan direction, exterior edges of the main wings and exterior edges of the ailerons which are connected via butt hinge can relatively rotate on the wings plane with the butt hinge as an axis, and the aileron plane can be folded under the main wings plane, the drive device comprises a main wings drive device and an aileron drive device which are driven to be connected via a main drive shaft, the main wings drive device is positioned in the head section, and composed of two symmetrical crank rocker mechanisms which respectively drive the left wing and the right wing, and the aileron drive device is positioned in the middle-rear portion of the wings chordwise direction. The aileron rotary folding and unfolding type flapping wing device employs a bevel gear pair to realize rotation shaft changes, thereby multi-degree-of-freedom motion can be driven by single power, approximately saving one main drive shaft and related motors, and greatly reducing weight.

Micro-electro-mechanical (MEM) translational tabs are introduced for enhancing and controlling aerodynamic loading of lifting surfaces. These microtabs are mounted at or near the trailing edge of lifting surfaces, deploy approximately normal to the surface, and have a maximum deployment height on the order of the boundary layer thickness. Deployment of this type of device effectively changes the camber, thereby affecting the lift generated by the surface. The effect of these microtabs on lift is as powerful as conventional control surfaces such as ailerons. Application of this simple yet innovative lift enhancement and control device will permit the elimination of some of the bulky conventional high-lift and control systems and result in an overall reduction in system weight, complexity and cost.

The invention relates to a convertible aircraft operating method. According to the invention, the aircraft comprises: a fuselage, standard fixed wings which are equipped with ailerons, a tail unit with flight-control surfaces, engines, a rotor with blades, a transmission which is placed between the engines and the rotor and which is equipped with rotor clutch and braking means, a landing gear, means for transition from helicopter mode to gyroplane mode and vice versa, and means for direct or reverse transition from gyroplane/helicopter mode to aeroplane mode. The lift for a range of low speeds is produced by means of the rotor, while the lift for a range of high speeds is produced by means of the wings. In addition, the lift for a range of intermediate speeds can be produced using the wings and the rotor in gyroplane mode simultaneously, and take-off and landing can be performed in gyroplane mode or in helicopter mode with the engines coupled to the rotor. The aircraft comprises a hybrid helicopter/gyroplane/aeroplane aircraft and, as such, can perform the direct or reverse transition to aeroplane mode from both helicopter mode and gyroplane mode.

Improved mechanisms for folding and locking an aircraft wing and controlling the wing's aileron through a wing fold have been motivated by development of a roadable aircraft. Applicable to a broader class of aircraft, these mechanisms allow for a safer, lighter-weight solution to the wing-folding challenge than currently available. The cable control system allows an outer wing section to be moved in concert with an actuated inner wing section. The locking mechanism allows for automated operation and visual inspection. The aileron control mechanism provides centering and locking in addition to flight control through the hinge axis.

The invention discloses a method for testing ground load calibration of an airplane wing and a special device for ground load calibration of the airplane wing. The method comprises the following steps: 1, calibrating a coordinate system, setting an intersection point of a horizontal line of an airframe structure of the airplane and a zero bit line as the original point of the coordinate system toensure that an X axis is coincided with the horizontal line of the airframe structure of the airplane and is positive afterwards, an Y axis is in a symmetric surface of the airplane and is positive upwards, and a Z axis points at the left wing of the airplane and forms a right-handed system with the X axis and the Y axis; 2, supporting a main undercarriage of the airplane on the ground, supportinga front undercarriage of the airplane on a platform, maintaining the wing in a horizontal state, and leading a fore flap, a tailing flap and a flap to be in the horizontal state; 3, applying restraint load to left and right horizontal stabilizer shafts of the airplane; and 4, selecting load points of the wing, applying calibration load to the selected load point of the wing, and acquiring the value of the calibration load. The method realizes field ground load calibration of the airplane wing.

The invention discloses a fault detection, diagnosis and performance evaluation method for a redundant aileron actuator. According to the method, fault detection, diagnosis, evaluation and real-time detection of the actuator are performed by means of an input order signal, an output displacement signal, a force motor current signal and aerodynamic loading data of the actuator; the fault detection is realized by a two-stage neural network, a first neural network is used as a system observer and is matched with actual output to acquire a residual error, and a second neural network outputs a self-adaptive threshold value synchronously; the fault detection is realized by the system observer and a force motor current observer; a time domain feature is extracted from a residual error signal and output to a self-organizing mapping neural network, and a minimum quantization error is acquired and normalized to a health degree, so that the actuator performance is evaluated; and on the basis of fault detection, the aerodynamic loading data is introduced, by means of a specific input order spectrum, the system observer and the neural network with the self-adaptive threshold value are trained, and the real-time fault detection is realized.

Wing tip vortices are evident from airliner vapor trails, and helicopter blade slap. Elliptically loaded high aspect ratio tapered wings have minimum induced drag but cannot eliminate it. Different methods are disclosed herein, for upper and lower surface boundary layers to cancel their opposing vorticity upon shedding from the trailing edge, thereby eliminating wake vorticty, induced drag and associated noise. This requires wing-rotor-propeller or fan blades with a platform designed for uniform bound circulation and with boundary layer control near the tip. In addition this requires special techniques to counter span-wise pressure gradients, such as tip circulation control blowing or an upwind small propeller or wind turbine on each tip. These techniques can eliminate wake vorticity with its induced drag, noise, flying on the backside of the power curve and the option for asymmetric loading by pneumatic means to eliminate need for cyclic pitch control or conventional ailerons.

The present application relates to the compressor stator of an axial turbomachine. The stator comprises an annular row of main stator blades and auxiliary blades each of which are associated with a main blade. The auxiliary blades are located at the trailing edges of the main blades and are in the vicinity of the pressure faces of the main blades. The auxiliary blades are aligned to generate a low-pressure area at the trailing edges of the main blades. Thus, a flow bypassing a main blade by its suction face is sucked in by the low-pressure area when it approaches the trailing edge of the main blade. Stalling is thus avoided and the efficiency of the machine is improved.

The invention provides a single aircraft with vertical take-off and landing capacity. The single aircraft comprises a body (1), wings (2), rotor wings (5 and 13) which are arranged on the wings (2) on two sides, and elevons (7) which are arranged at the rear edges of the two wings (2), wherein rolling and pitching can be stabilized and controlled through deflection of the elevons (7); in a vertical take-off and landing period, the head of the aircraft is upward; the aircraft takes off and/or lands in a tail seat manner; the gravity is overcome through elevating force generated by rotation of the rotor wing (5); vertical take-off and landing can be realized. The single aircraft has the advantages that (1) through a tailless layout, the air resistance is effectively reduced, and the flight speed is increased; (2) the advantages of a helicopter and a fixed wing aircraft are taken into consideration, and high efficiency can be maintained in both high-speed level flight and low-speed hovering periods; (3) a rotary mechanism is eliminated, and the single aircraft is simple in structure, high in reliability and low in cost; (4) the aircraft can be stabilized and controlled under low-speed and large-incidence conditions, and is applied to city airspace; (5) an emergency is handled in an engine connecting shaft and full-aircraft parachute-opening manner, and the aircraft is high in safety.

The invention relates to an automatic-piloting simulator of an aeroplane, and belongs to the field of simulation technique. The automatic-piloting simulator of the aeroplane is characterized by comprising a flight trainer, a model airplane rudder control mechanism and a flight vision system which are connected into a whole by a computer network, wherein the flight trainer is provided with a server, a central control platform, a side pole and a display mechanism; the model airplane rudder control mechanism is provided with a semi-physical model, ailerons, a rudder, an elevator and a rudder loop control system of the aeroplane; and the flight vision system is driven by aeroplane motion data generated by the flight trainer, and displays a flight attitude, flight data and a flight trajectory image of the aeroplane. The automatic-piloting simulator of the invention has the advantages of simple structure, compact assembly, economy, practicality, accurate data, high simulation degree, good automatic and manual piloting dynamic performance of the simulated airplane and the like.

An aircraft defining an upright orientation and an inverted orientation, a ground station; and a control system for remotely controlling the flight of the aircraft. The ground station has an auto-land function that causes the aircraft to invert, stall, and controllably land in the inverted orientation to protect a payload and a rudder extending down from the aircraft. In the upright orientation, the ground station depicts the view from a first aircraft camera. When switching to the inverted orientation: (1) the ground station depicts the view from a second aircraft camera, (2) the aircraft switches the colors of red and green wing lights, extends the ailerons to act as inverted flaps, and (3) the control system adapts a ground station controller for the inverted orientation. The aircraft landing gear is an expanded polypropylene pad located above the wing when the aircraft is in the upright orientation.

An aircraft control surface failsafe hinge configuration is disclosed. The arrangement comprises a drop link configuration which is adapted to attach a control surface to an aircraft structure where the drop link additionally incorporates a failsafe hinge means which provides a backup hinge connection between the control surface and the aircraft structure to which it is attached. The improved failsafe hinge construction may be applied to any aircraft control surface such as an aileron, elevator, vertical tail plane and the like.

The invention discloses a GPS-guided unmanned aerial vehicle automatic carrier-landing adaptive control system and method. The system includes: a GPS guidance reference trajectory generation and trajectory error calculation module which is used for inputting signals detected by a GPS, establishing a reference trajectory in a ground coordinate system with an ideal carrier-landing point being an original point, and finally outputting the signals; a longitudinal guidance law module which uses a pitch attitude as an internal loop, and realizes control of flight height by suppression of a height error; a lateral guidance law module used for subtracting an actual lateral position signal from a designated lateral position signal to obtain an error signal and eliminating the error signal; and a flight control loop includes control law modules of four channels of accelerator, elevator, aileron and rudder. The GPS-guided unmanned aerial vehicle automatic carrier-landing adaptive control system provided by the invention realizes conversion of a trajectory tracking error signal into an attitude tracking command signal, solves the problem of attitude tracking through adaptive control, and an unmanned aerial vehicle automatic carrier-landing guidance and control system is formed.

Provided is a cross-medium aircraft with a changeable shape like a flying fish. A pair of machine wings (2) are movably installed on the two sides of the middle of a machine body (1) respectively. A pair of duck wings (3) are installed below the middle rear portion of the machine body (1). The machine wings (2) and the duck wings (3) can rotate around rotating shafts to change the tapping angle. A flap and an aileron are distributed on the inner side and the outer side of each machine wing (2) respectively. An air-propulsion foldable spiral paddle (4) is installed at the head portion of the machine body (1). A V-shaped tail wing (5) is arranged at the middle rear section of the machine body (1). An underwater-propulsion spiral paddle (6) is installed at the tail portion of the machine body (1). A rectifying cover (7) is arranged at the lower middle portion of the head of the machine body (1). The cross-medium aircraft has the advantages that the air-water dynamic layout of two different media including air and water and the light pressure-resistant shape-changeable structure are taken into consideration, multi-mode control and water sealing can be easily achieved, and the various military and civilian requirements can be met.

A supersonic aircraft having unique structural, geometrical, electromagnetic, mechanical, thermal and aerodynamic configuration is described. Its design characteristics maximize aerodynamic performance, speed, efficiency, comfort and range in the operational affect of carrying passengers over long distances at very high flight speeds and Mach numbers, generally above Mach 4.0+. It has fully integrated-mated fuselage twin engine nacelle structures which reduce and stabilize wave drag, wherein engines sit forward of the sonic boom electrode swept aerospace but are integrated right through the sub-wing, the aero-spike and inlet just above the top surface of the central wing, and above and the major wing mating join, attached via a massive composite titanium keel structure, through the central wing box and fuselage.

The invention relates to a simulation device for a steering wheel of an airplane flight simulator. The device mainly consists of a steering wheel surface, a transmission mechanism, angular displacement sensors, a mechanical structure, a serial bus (USB), an HID (Human Interface Device), and the like. By controlling the steering wheel to make longitudinal displacement and rotary displacement around a vertical shaft, the device controls the state of elevators and the state of ailerons of the flight simulator, thus controlling the flight attitude; and the device simulates the operating force of an airplane through a spring and the structural frictional force. The device has a simple structure and high simulation degree, and directly measures the displacement value of the steering wheel by adopting an aileron angular displacement sensor and an elevator angular displacement sensor, thus avoiding the error of mechanical transmission of indirect measurement and achieving high measurement accuracy. The device adopts a universal interface so that a driver does not need to be installed on a computer of a host machine, and the device can be applied to different types of flight simulators.