238 results about "Wingspan" patented technology

The invention provides a morphing wing integrating telescopic variable-elongation and rotary variable-sweepback functions and a morphing method. The morphing wing comprises an inner wing section, an outer wing section and an elastic skin, the inner wing section and the outer wing section are positioned in the same spanwise plane; the inner wing section stretches and retracts in the spanwise plane along the spanwise direction of the wing, so that the span length of the inner wing section is adjusted; the outer wing sections rotate in the chordwise direction relative to the inner wing sections in the spanwise plane, and then the sweepback angles of the outer wing sections are adjusted. By adjusting the extension length and the sweepback angle of the wings, the lift force of the aircraft during low-speed flight is improved, and the resistance of the aircraft during high-speed flight is reduced. Specifically, the extension length of the wings is increased during low-speed flight to improve lift force, and the extension length is reduced and the sweepback angle is increased during high-speed flight to reduce resistance. The optimal aerodynamic performance can be achieved no matter in a low-speed flight state or a high-speed flight state, and the adaptability and the flight performance of the aircraft are remarkably improved.

The invention discloses a deformable wing with the characteristics of large unfolding-folding ratio and high bearing capacity. The deformable wing comprises a bearing structure, a driving mechanism, a wing surface dimensional structure and an accessory structure, the bearing structure comprises a plurality of groups of telescopic mechanisms, wing ribs and corner boxes; the accessory structure comprises a mounting plate, and the mounting plate is connected with the machine body; each group of telescopic mechanisms comprises a plurality of stages of telescopic rods which are arranged along the chord length direction and are fixed on the mounting plate; the wing ribs are arranged in the wingspan direction and are connected with the multi-stage telescopic rods through the angle boxes to form a wing framework; the driving mechanism penetrates through the hollow parts of the wing ribs, and two ends of the driving mechanism are respectively connected with the mounting plate and the wing rib on the outermost side of the wing; the wing skeleton is externally coated with the wing surface dimensional structure to form a wing shape.

The invention discloses a regional unmanned aerial vehicle safety risk assessment method, which comprises the following steps: acquiring parameters such as unmanned aerial vehicle body length, wingspan width and body height, and preset task scene data such as regional tall building area ratio, unmanned aerial vehicle density, specific parameters and quantity of aircrafts; based on the data, regional ground and airspace safety risk values are calculated. And determining visibility, rainfall and wind speed as meteorological risk assessment indexes, performing weighting by using an analytic hierarchy process, and calculating a meteorological risk assessment coefficient. And constructing a regional unmanned aerial vehicle safety risk assessment model by integrating assessment data obtained under three environmental conditions of weather, ground and airspace. And comprehensively evaluating the safety risk of the unmanned aerial vehicle in the region by using the model. According to the technical scheme of the invention, three environmental factors of meteorological conditions, ground conditions and airspace conditions are integrated, and various environmental risk types possibly experienced by the unmanned aerial vehicle in the operation process can be comprehensively evaluated.

According to various embodiments, there may be provided a monocopter. The monocopter may include a body chassis. The monocopter may further include a wing structure extending from the body chassis, the body chassis being at a root of the wing structure. The wing structure may include a plurality of rigid wing segments distributed along a spanwise direction extending between the root and a tip of the wing structure. The wing structure may further include a plurality of flexible wing segments, each flexible wing segment adjoining a pair of adjacent rigid wing segments of the plurality of rigid wing segments. The monocopter may further include a propulsion unit coupled to a rigid wing segment of the plurality of rigid wing segments between a midspan of the wing structure and the tip of the wing structure.

The invention relates to the field of aircraft wing design, in particular to a multi-stage telescopic wing for improving the flight performance of an aircraft, the extension length and the area of the wing are changed through multi-stage stretching of the wing, and different requirements of the aircraft for the wing in low-speed and high-speed flight states are met. When the aircraft takes off, lands and flies at a low speed, the telescopic wings stretch out to increase the area of the wings and enable the aircraft to obtain enough lift force, and when the aircraft flies at a high speed, the telescopic wings retract to reduce the extension length and area of the wings and enable the aircraft to obtain enough lift force and reduce the resistance of the wings, so that the extension length and area of the wings are changed through extension of the wings; and the aircraft can simultaneously meet different requirements on the extension length and the area of the wings in low-speed and high-speed flight states.

Towed aerodynamic platforms approach the behavior of flat plate airfoils to achieve high lift-to-drag (L:D) performance when operated at low pitch angles. A lead aircraft may tow multiple platforms to form train units. Low form drag and high L:D enable solar aircraft that are more robust and faster than approaches extending wingspan rather than extending the length of the aerial train. Applications extend beyond solar aircraft and include vehicles towed along an overhead monorail system and aerial drones.

The invention relates to the technical field of unmanned aerial vehicles, in particular to an inflatable wing type variable wingspan unmanned aerial vehicle. The unmanned aerial vehicle comprises a vehicle body, variable inflatable wings are arranged on the two sides of the vehicle body, the variable inflatable wings are connected with an inflation and deflation mechanism arranged in the vehicle body, each variable inflatable wing comprises a hollow fixed inflatable wing body and an extended inflatable wing body, and an inflatable part of each extended inflatable wing body is arranged outside a hollow cavity of the corresponding hollow fixed wing body; the uninflated part of the extended inflatable wing body is arranged in the hollow cavity, and the extended inflatable wing body is connected with the telescopic part of the extended driving mechanism. According to the inflatable wing type variable wingspan unmanned aerial vehicle, the variable inflatable wings are adopted, the wingspans of the whole wings are adjusted by lengthening the inflatable wing bodies, the unmanned aerial vehicle adapts to different flight states, and the unmanned aerial vehicle is of a relative folding structure, so that multi-stage adjustment is achieved.

The invention discloses a safety interval calculation method for an unmanned aerial vehicle and a manned aircraft, and relates to the technical field of aviation safety. The safe interval calculation method comprises the following steps: step 1, firstly obtaining the length, wingspan and height of a first target, then obtaining the length, wingspan and height of a second target, marking the first target as a mass point at the moment, and making up for an error caused by volume loss when the first target is regarded as the mass point. A new collision model is used for carrying out quantitative analysis on the collision probability of a manned aerial vehicle and an unmanned aerial vehicle so as to find the relation between the flight interval and the collision probability and calibrate a reasonable safety interval, and the requirement of the safety interval with a fixed value for the airspace cost is high. Therefore, the dynamic safety interval is found by adopting interval division iterative calculation, and a more flexible dynamic safety interval between the unmanned aerial vehicle and the manned vehicle is given on the basis, so that the construction of a fusion flight system is accelerated.

The invention belongs to the technical field of aircraft flight load design, and particularly relates to an aircraft pneumatic rolling torque correction method, which comprises the steps of initializing calculation parameters, calculating aircraft vacuum speed, equivalent speed and aircraft speed pressure based on the calculation parameters, then calculating motion parameters based on maneuvering simulation dynamic response, and calculating full-aircraft target pneumatic rolling torque based on the motion parameters. Giving unbalanced torque tolerance; solving an unbalanced moment based on the whole-aircraft target rolling moment; and comparing and analyzing the difference value between the unbalanced moment and the unbalanced moment tolerance, and if the absolute value of the unbalanced moment is smaller than the unbalanced moment tolerance, ending the calculation. By adjusting the local attack angle of each section in the spanwise direction of the wing part and reconstructing the pressure distribution characteristic in the spanwise direction of the wing, the unbalanced pneumatic rolling moment of the aircraft caused by the pressure distribution characteristic is eliminated, the purpose of load balance is achieved, and the design requirements in engineering are met.

An aircraft wing tip that is movable by an actuator between a flight configuration, and a ground configuration which has a shorter wingspan for use in ground-based operations. Lugs that connect to arms of the actuator are integrally formed as part of the primary structure of either the fixed wing or outboard wing section, such as a folding tip section.

The present invention provides an extreme STOL airplane comprising: a fuselage and a wing having port and starboard wing sections. Each wing section has a body with top and bottom surfaces, leading and trailing edges, and a span, i.e., the distance between a wing section tip and root connected to the fuselage, and flap coves. Aerodynamic elements integrated into the wing sections include Fowler flaps, having leading and trailing edges and a span, flap tracks, wherein the flap tracks are external to the wing section body, extend aftward beyond the trailing edge of the wing section, and are configured to enable the Fowler flaps to rotate or deflect and to translate or extend and retract, Frise ailerons, wherein the Frise ailerons are located outboard of the Fowler flaps, and spoilerons located over the leading edge of the Fowler flaps when the flaps are in the fully extended position.

The invention provides a horizontal measurement method of a large-wingspan movable wing aircraft, which comprises the following steps of: establishing a coordinate system by taking a virtual coordinate on a tool as a reference, adjusting a basic platform and a supporting arc to be within an allowable deviation range, positioning a middle section of a product by using the supporting arc, measuring the middle section, judging whether the levelness of a mounting surface of the middle section reaches the standard or not, and adjusting. And sequentially assembling the rear section and the front section by taking the middle section as a standard of complete machine assembly, completing assembly and measurement of the fuselage, and adjusting each section of the fuselage in real time according to a measurement result. And then the wing is assembled, and measurement of the wingtip coordinate and the deflection angle is completed. The tool has the advantages that assembly and measurement are carried out on one tool, alternate station change of products between an assembly table and a measurement table is avoided, a machine body does not need to be rotated during measurement, measurement is completed in the assembly state, the assembly efficiency is guaranteed, and the product safety can be guaranteed.

The utility model relates to a gliding type precipitation enhancement bomb which comprises a bomb body, a telescopic wingspan frame, a gravity center seat, an endurance motor and flying wings, the telescopic wingspan frame is arranged on the bomb body in a foldable mode, the gravity center seat is arranged between the bomb body and the telescopic wingspan frame, and the endurance motor is arranged on the telescopic wingspan frame. And the flying wings cover the telescopic wingspan frames and are unfolded along with the telescopic wingspan frames. The rain enhancement bomb is simple in structure, convenient to produce and manufacture and low in manufacturing cost. In addition, besides the advantage that rainfall is sufficient due to long-time hang, the gliding type rain enhancement bomb is low in speed and safer when landing.

The invention belongs to the technical field of aircraft design, and relates to an aircraft wing rib section mass discretization method, which comprises the following steps of: (1) setting a virtual trapezoidal shape covering a wing, and equidistantly setting a plurality of chordwise grid points on each spanwise tangent plane on the virtual trapezoidal shape along the spanwise direction of the wing; (2) determining the position of the mass gravity center of the rib section on the spanwise tangent plane; (3) dispersing the mass of the rib sections to adjacent spanwise tangent planes along the spanwise direction according to a lever ratio; (4) dispersing the mass of the section to chordwise grid points along the chordwise direction according to trapezoidal distribution; (5) segmental integration is carried out to calculate the quality of chordwise grid points; according to the method, a trapezoidal shape is arranged along the shape of a wing, the mass of a wing rib section is dispersed to adjacent spanwise sections along the spanwise direction of the wing according to a lever ratio, and the mass of the sections is dispersed to chordwise grid points along the chordwise direction of the wing according to trapezoidal distribution. The mass is reasonably dispersed under the condition that the mass gravity center of the rib section is not changed, so that the wing mass distribution is closer to the real condition, the wing load calculation precision is improved, and the structural design is safer and more reliable.

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.

A method for reducing aerodynamic loads on an aircraft located on the ground includes moving a wing tip relative to a fixed wing to an on-ground configuration in which the span of the wing is reduced. The wing tip includes an airflow channel, which is in an open configuration in which airflow through the channel between upper and lower surfaces of the wing tip is enabled, when the wing tip is in the on-ground configuration. This has been found to reduce aerodynamic loads on the wing tip and / or parts of the aircraft connected thereto during this phase of the aircraft's use.

Closed wing aircraft design and configuration with three wings on either side of the fuselage wherein at least one closed frame is established between the 1st wing and the 2nd wing, by using at least one bracing entity. A separate closed frame is established between the 2nd wing and the 3rd wing, by using at least one bracing entity. Each of said closed frames defines its own aerodynamic channel. The framework is strong and stiff, the mutually supported parts level out the stress. The fuselage is lifted in three points. The Aspect Ratio is high and the wingspan is relatively short but there are embodiments wherein the 3rd wing has an extended folding tip section.

The invention provides a patrol flight unmanned aerial vehicle with W-shaped folding wings. The patrol flight unmanned aerial vehicle comprises a fuselage, inner-section wings, inner-section wing limiting structures, inner-side folding and unfolding mechanisms, outer-section wings, ailerons, outer-side folding and unfolding mechanisms, empennages, empennage folding and unfolding mechanisms and empennage protection structures. The fuselage is designed by adopting a flat lifting body configuration, the inner and outer sections of wings are designed by adopting low-speed friendly plane shapes such as rectangles and trapezoids, and the empennage is designed by adopting a V-shaped empennage / inverted V-shaped empennage. The inner section wing and the outer section wing of the unmanned aerial vehicle are integrally in a W shape in the unfolding process, and a folding wing layout with the larger aspect ratio and the higher aerodynamic efficiency can be derived based on the W-shaped layout. According to the layout, the extension length of the wings is slightly restrained by the length of the fuselage, the contradiction between the aerodynamic characteristic design of the wings and the storage space of the launch canister is well balanced and solved under the condition that the length of the fuselage is fixed, and the layout wings are large in aspect ratio, high in aerodynamic efficiency and simple in structural design.

A wing structure is disclosed having a first span portion including a depth that reduces in a spanwise outward direction, and a second span portion outboard of the first span portion and having a substantially uniform spanwise depth D3. The second span portion may be non-tapered, and may include parallel fore and aft spar portions. In some embodiments the non-tapered portion supports a movable wing tip device. A wing assembly and an aircraft are also disclosed.

The invention discloses a flying wing and buoyancy lift integrated layout mooring aerostat type high-altitude wind power generation system which comprises an air aerostat wind power generation system, an air-ground electric energy transmission and mooring cable and a ground anchoring system. The air aerostat wind power generation system comprises a flying wing configuration thin-wall truss force bearing structure, a plurality of buoyancy air bags, two empennages, a skin component and three wind power generator sets, and the flying wing configuration thin-wall truss force bearing structure is provided with a plurality of constraint spaces distributed in the wingspan direction; the plurality of buoyancy airbags are respectively arranged in the plurality of constraint spaces; the two empennages are arranged at the two ends of the rear portion of the flying wing type thin-wall truss force bearing structure respectively. The skin component is arranged on the flying wing configuration thin-wall truss force bearing structure and the two empennages in a sleeving manner; the three wind generating sets are installed at the bottom of the flying wing type thin-wall truss force bearing structure through a supporting structure. The device is stable in structure, can stably generate power in a long-time air-staying mode, is low in cost, and is autonomous and controllable.

The utility model discloses a telescopic variable-pitch cross-medium propeller mechanism which comprises a vane seat shell, a supporting frame is arranged in the vane seat shell, an electric telescopic rod assembly is arranged on the supporting frame, vane assemblies are connected to the left end and the right end of the electric telescopic rod assembly in a threaded mode, each vane assembly is formed by sleeving a plurality of multi-section vanes, and the vanes are arranged on the vane seat shell. The front end of the electric telescopic rod assembly is connected with a wing piece at the foremost end of the wing piece assembly, and the wing piece assembly is driven by the electric telescopic rod assembly to stretch and retract. Proper lift force is provided for the cross-medium aircraft or the aircraft, modular design is adopted, the wing panel serialization design can be carried out according to the load requirement of the cross-medium aircraft or the aircraft, the wing panel extends into a large wingspan during air flight, the wing panel contracts into a small wingspan during underwater navigation, and the screw pitch of the propeller is increased.

A wing component for an aircraft, comprising a leading edge part and a trailing edge part pivotably mounted to the leading edge part so as to pivot about a pivot axis, an actuator unit configured for moving the trailing edge part relative to the leading edge part, and a damper device configured for damping uncontrolled movement between the leading edge part and the trailing edge part. The trailing edge part comprises a first part portion and a second part portion arranged next to each other in a span direction and configured to pivot about the pivot axis individually, the damper device comprises a damping element coupled between the first part portion and the second part portion to damp relative movement of the first part portion and the second part portion.

A vertical take-off and landing aircraft includes a fuselage, at least one propulsion system, at least four lift propellers and at least two fixed wing planes. The fixed wing planes which are located behind the most forward lift propellers are located above the lift propellers. The lift propellers are distributed on either side of at least one of the fixed wing planes and on either side of the fuselage such that the lift propellers are longitudinally separated by at least the magnitude of the chord of the fixed wing plane located therebetween, and are laterally separated by at least the width of the fuselage located therebetween. At least one of the fixed wing planes has a variable span and the aircraft includes a control system for varying this span in flight.

The present invention provides an extreme STOL airplane comprising: a fuselage and a wing having port and starboard wing sections. Each wing section has a body with top and bottom surfaces, leading and trailing edges, and a span, i.e., the distance between a wing section tip and root connected to the fuselage, and flap coves. Aerodynamic elements integrated into the wing sections include Fowler flaps, having leading and trailing edges and a span, flap tracks, wherein the flap tracks are external to the wing section body, extend aftward beyond the trailing edge of the wing section, and are configured to enable the Fowler flaps to rotate or deflect and to translate or extend and retract, Frise ailerons, wherein the Frise ailerons are located outboard of the Fowler flaps, and spoilerons located over the leading edge of the Fowler flaps when the flaps are in the fully extended position.

The invention discloses a multi-mode self-rotating tailstock type vertical take-off and landing aircraft. Multi-mode motion comprises a multi-axis mode, a self-rotating mode and a fixed wing mode. The unmanned aerial vehicle comprises a fuselage, a variable collective pitch structure and four symmetrical wings on the two sides; the wing type of each wing is a symmetrical wing type, the tail end of each wing is a full-moving wing tip, and each wing is provided with a driver; the driver is a propeller driven by a motor; each wing, the driver and each full-motion wingtip can rotate by + / -90 degrees around a rotating shaft in the spanwise direction of the wing, and switching among multiple modes and attitude control are achieved. The aircraft disclosed by the invention has more flexible maneuvering capability, has vertical take-off and landing performance of a multi-axis mode, efficient hovering performance of a spinning mode and high-speed cruising performance of a fixed wing mode, can meet different flight requirements, and has better environment adaptability.

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.

An aircraft wing including a flap system mounted at a leading or trailing edge of the wing and includes at least two movable flap devices spaced apart in a spanwise direction Y from each other, each movable flap device supported by at least one support station and driven about a hinge axis by at least one drive system. The flap system includes a synchronizing torque element connected to the movable flap devices, wherein the synchronizing torque element is configured to transmit a torque between the movable flap devices to actuate a rotational movement of the movable flap devices, and wherein the synchronizing torque element is positioned between the movable flap devices and around an adjacent system or structural component that is also positioned between the movable flap devices.

The application provides a morphing aircraft fault detection model generation method and a fault detection method. The morphing aircraft fault detection model generation method comprises the following steps: obtaining a state control feedback matrix corresponding to each wingspan change rate based on an initial value of each wingspan change rate, an initial value of an aircraft state variable, and a preset flight linear change parameter model of the morphing aircraft; generating flight simulation data based on to-be-simulated data, wherein the to-be-simulated data comprises the initial value of the aircraft state variable, the wingspan change rate, and the state control feedback matrix, and each set of flight simulation data comprises flight state parameters at multiple time points; generating sample data based on the flight simulation data, wherein the sample data comprises the flight simulation data and a fault label corresponding to the flight simulation data; and obtaining the morphing aircraft fault detection model by training based on multiple sets of sample data. The application can effectively detect faults of the morphing aircraft.

The utility model relates to the technical field of wingspan type folding boxes, in particular to a wingspan type folding box which comprises a bottom plate and a first vertical plate, four corners of the outer wall of the bottom plate are fixedly connected with the bottom of the inner side of the first vertical plate respectively, a first straight plate is rotatably connected to the upper portion of the inner side of the first vertical plate, and a hinge is fixedly connected to the top of the outer side of the first straight plate. Second straight plates are fixedly connected to the outer walls of the hinges, control structures are connected to the front portions and the rear portions of the left side and the right side of the bottom plate correspondingly, and opening and closing structures are connected to the front sides and the rear sides of the bottom plate correspondingly; the inner side of the outer wall of a first screw rod is in threaded connection with a round opening of the outer wall of a first straight plate, at the moment, the first screw rod limits pulling of the first straight plate, the first straight plate is prevented from rotating through a pin shaft on the inner side of a first vertical plate, the situation that an object is pressed by inner side rotation is avoided, and the protection performance of the wingspan type folding box is improved; and normal use of the wingspan type folding box is ensured.

The invention discloses an aircraft takeoff stress measuring device and an aircraft takeoff safe taxiing lift envelope obtaining method, and relates to the technical field of aircraft test flight tests.A target aircraft is carried and moved through a T-shaped body support provided with wheels, the T-shaped body support is matched with the approximate distribution of an aircraft landing gear, and the aircraft takeoff safety taxiing lift envelope is obtained. The undercarriage and the movable flat plates are connected through the connecting mechanism after the aircraft is supported and adjusted, accidental take-off caused by gust wind or unstable control over the speed of a sports car is prevented, and the vehicle traction rod is used for being connected with the sports car; the target aircraft is dragged to slide, the force borne by the aircraft in the off-frame and on-frame sliding process is fed back through the force measuring sensor, the force bearing condition of the aircraft is accurately measured, and the take-off state of the aircraft can be confirmed by combining the rudder deflection angle parameter and the sliding speed set in the testing process. And finally, determining a lift force safe taxiing envelope of the aircraft, so as to realize the flight test of the ultra-light large-wingspan aircraft.