50 results about "Wing configuration" patented technology

The invention belongs to the technical field of glide control, and relates to a direct lift active disturbance rejection control method suitable for flying wing layout. According to the method, a trailing edge flap is introduced to serve as a main lift control surface, meanwhile, a constant attack angle is kept through a lifting aileron so as to balance part of pitching moment generated by the trailing edge flap, and therefore the attitude of the aircraft is stabilized, and the constant speed is maintained. In the actual flight process, external disturbance is inevitable, so that an original glide control strategy is improved and optimized in combination with an active disturbance rejection control method. Compared with a traditional glide control method, the flying wing layout direct lift active-disturbance-rejection control method provided by the invention has remarkable anti-interference capability and excellent glide control performance, and shows a wide application prospect.

Provided are computer-implemented methods for autonomously controlling an aircraft with vertical take-off and landing capabilities and folding wings that includes controlling a plurality of thrust producing components of an aircraft to cause the aircraft to rise vertically when wings of the aircraft are in a first folded configuration, where when the wings of the aircraft are in the first folded configuration, a leading edge of each wing is oriented in a vertical direction setting motor controller gains based on the wings of the aircraft being in the first folded configuration, and causing the aircraft to align with a direction of airflow when the wings of the aircraft are in the first folded configuration, and controlling thrust producing components and control surfaces and internal articulation mechanisms of the aircraft to cause the aircraft to transition from folded wing configuration to unfolded wing configuration. Systems and computer program products are also provided.

A system of a multi-rotor aircraft that capitalizes on the advantages of fixed wing elements combined with rotary wing structures. The fixed wing elements can help to generate lift once the aircraft is airborne and can thus reduce the need for larger lifting rotors which can allow for longer flight times and distances. Additionally, the systems disclosed herein take advantage of a partial in-wing configuration with a number of rotors to reduce the overall footprint of the vehicle while maintaining the flight efficiency that comes with combining features of fixed and rotary wing elements, and increasing operator safety by shrouding rotating parts. The unique configurations allow for a decoupling of the pitch, yaw and roll authority to reduce the complexity in control systems and improve the flight efficiency of the aircraft. Additional configurations implement the use of smaller thrust rotors that can be used to generate thrust as well as control yaw and thus counteract any remaining unbalanced torque.

The invention provides a design method of a variable-size variable-density lattice structure wing with optimized heat dissipation load, and relates to the technical field of aerospace engineering, and the design method comprises the following steps: S10: carrying out variable cross-section design on a D-type unit to obtain a variable cross-section D-type unit, taking the variable cross-section D-type unit as a lattice unit configuration, and filling the wing to form a skin lattice structure; the skin lattice structure of the wing comprises a skin and a lattice filling area; s20, cell element variable-size design is adopted for the dot matrix units, and variable-size dot matrix units are constructed; s30, according to the stress field of the wing, cell element variable-density design is adopted for the variable-size dot matrix units, and variable-size variable-density dot matrix units are constructed; and S40, establishing a wing configuration. The variable-cross-section D-type unit is used as a dot matrix unit configuration, deep coupling of integrated design of heat dissipation and bearing performance is achieved through variable-size and variable-density design, application scenes are flexibly adjusted, and complex working condition requirements are met.

The invention discloses a dynamic modeling and analysis method for a gull-wing morphing aircraft in the technical field of flight system control. The dynamic modeling and analysis method comprises the following steps: S1, calculating aerodynamic force and torque borne by the aircraft according to a three-dimensional lift line theory; s2, establishing a multi-body dynamic model through a quasi-Lagrange equation; s3, analyzing the aerodynamic characteristics of the aircraft according to the calculation result of the experiment; s4, analyzing the kinematics characteristics of the aircraft according to the calculation result of the experiment; and S5, adjusting the wing configuration according to the analysis result, so that the aircraft keeps stable flight. The method solves the problem that a traditional variant aircraft is not matched in performance at different flight stages, especially is large in lift force loss under the low-speed condition.

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 application discloses a deformable wing, an aircraft and a deformation control method, and relates to the field of aircraft design.

The application provides a fixed-wing aircraft airspeed calibration flight test method based on a standard machine. The method comprises the following steps: determining the weight of the fixed-wing aircraft as a first weight, setting the aircraft to a specified take-off configuration, and taking off according to a take-off procedure; the standard machine flies to a specified airspace according to a plan; at an initial specified flight height, a tow cone is released to a specified length, and the tow cone can obtain an accurate static pressure value in the current area without being affected by a spoiler; the flap and slot wing configuration of the fixed-wing aircraft is set to a first flap and slot wing configuration; the aircraft is trimmed at a given speed; the speed in the above step is set to other speeds within the speed envelope of the first flap and slot wing configuration, and the steps are repeated; the standard machine obtains airspeed errors under the first flap and slot wing configuration and different speeds; and the airspeed errors under the first flap and slot wing configuration and different speeds are fitted to obtain an airspeed correction curve under the first weight, the first flap and slot wing configuration. More accurate airspeed errors can be obtained.

Toy aircraft comprise a fuselage and a main wing coupled to the fuselage. The main wing is configured to be bent and / or twisted at any position along a substantial span-wise length of the main wing amongst a plurality of distinct wing configurations and the main wing is configured to be retained in a selected one of the plurality of distinct wing configurations. In some examples, the main wing comprises an external body and a malleable frame that is supported by the external body. The malleable frame is configured to be selectively bent and / or twisted at any position span-wise along the malleable frame amongst a plurality of distinct wing configurations and the malleable frame is configured to retain the main wing in a selected one of the plurality of distinct wing configurations. The external body is configured to be pliable to conform to the selected one of the distinct wing configurations.

The invention discloses a forward-swept-wing variable-configuration unmanned aerial vehicle and a configuration changing method thereof, and relates to the technical field of aircrafts, the forward-swept-wing variable-configuration unmanned aerial vehicle comprises a fuselage, two wings are symmetrically arranged on two sides of the fuselage, and the two wings and the fuselage form a forward-swept-wing configuration; the wings can horizontally rotate around the connecting positions of the wings and the fuselage so as to change the size of a sweepforward angle; the rotor wing is arranged at the wingtip position of the wing, and the rotor wing can rotate around the axis of the wing where the rotor wing is located; and the two vertical empennages are symmetrically arranged above the tail part of the fuselage. According to the forward-swept wing variable-configuration unmanned aerial vehicle and the configuration changing method thereof, it is guaranteed that the unmanned aerial vehicle can have good stability and aerodynamic performance under different flight working conditions.

The present disclosure relates to a variable aircraft. One of the first fixed wing and the second fixed wing is fixedly connected to the left side of the fuselage, and the other is fixedly connected to the right side of the fuselage; one end of the first connecting rod is hinged to the first fixed wing, and the other end is hinged to the first movable wing; one end of the second connecting rod is hinged to the second fixed wing, and the other end is hinged to the second movable wing; the directions of the hinged shafts are consistent with the longitudinal axis of the fuselage; the first movable wing and the second movable wing can be switched to the cross-wing configuration under the action of aerodynamic force and roll torque, in which the first movable wing is connected to the fuselage at one end and suspended at the other end, and the second movable wing is connected to the fuselage at one end and suspended at the other end; the first fixed wing, the first movable wing, the second fixed wing and the second movable wing are arranged in sequence around the longitudinal axis of the fuselage, the first fixed wing and the second fixed wing are opposite to each other, and the first movable wing and the second movable wing are opposite to each other.

The present application relates to the technical field of aircraft design, and particularly relates to a wide-speed-range aircraft based on a variable high-pressure capture wing shape and a control method thereof, which comprises a fuselage, an upper wing surface, a lower wing surface, a foldable canard, a variable high-pressure capture wing and a control system. The variable high-pressure capture wing is arranged above the fuselage and the upper wing surface, and is used for switching between a subsonic forward-swept configuration and a supersonic backward-swept configuration according to a flight speed range; the foldable canard is arranged at a leading edge of the wing surface, and is used for unfolding in a subsonic stage and folding to be attached to the lower wing surface in a supersonic stage; and the control system determines a current stage according to flight state parameters, and cooperatively drives the high-pressure capture wing and the canard to perform a configuration switching action. The present application realizes adaptive matching of an aerodynamic shape and a flight speed range through variable design.

This invention discloses a method for improving the flutter velocity of a high aspect ratio telescopic wing unmanned aerial vehicle (UAV) with a body degree of freedom. The method is characterized by using the UAV's telescopic wing to perform asymmetrical extension and retraction to obtain an asymmetrical wing configuration, thereby suppressing body degree of freedom flutter in the asymmetrical configuration of the high aspect ratio wing and thus improving the flutter velocity of the UAV. This invention achieves an asymmetrical configuration of the UAV by controlling the asymmetrical extension and retraction of the high aspect ratio telescopic wing, thereby increasing the flutter velocity of the high aspect ratio UAV.

This invention discloses a stable flight control method for heavy-load unmanned aerial vehicles (UAVs) with compound wing configurations, relating to the field of UAV flight control technology. This invention utilizes a third-order nonlinear ESO to unify all internal and external disturbances, such as rotor-fixed-wing coupling disturbances, wind field disturbances, and model uncertainties caused by load changes, into a total disturbance for real-time estimation and feedforward compensation. This eliminates the need for precise modeling of complex coupled aerodynamic models and load change models, significantly reducing modeling difficulty. Simultaneously, it achieves full compensation for both matched and mismatched disturbances, significantly improving the anti-disturbance capability and robustness of heavy-load compound wing UAVs under strong disturbance scenarios, and solving the problems of attitude instability under sudden load changes, coupling disturbances, and complex wind fields.

This invention discloses a vernier-type precision internal resistance measurement device for low aspect ratio flying wing configurations, belonging to the field of wind tunnel testing. It includes a model to be tested on a support rod, two symmetrical support blades, one end of which is clamped to the support rod, and the other end of each support blade has a support cylinder. The axis of the support cylinder is aligned with the axis of the model to be tested. The rear section of the pressure-measuring rake is a columnar structure, which can be inserted into the support cylinder and moved along its axis. The front end of the pressure-measuring rake has several pressure-measuring holes, through which a pressure-measuring tube extends into the model to be tested. The structural improvements of this invention significantly enhance the overall performance of the support system, solving the long-standing technical bottleneck of low accuracy and poor stability in internal resistance measurement. It has significant engineering application value for promoting the development of low aspect ratio flying wing configuration aircraft.

The application relates to the technical field of unmanned aerial vehicles, and particularly discloses an unmanned aerial vehicle with a near-distance coupling tandem wing layout, which comprises a holder, a front normal lift wing, a rear normal lift wing, a flight controller, a propeller and an energy management system; the front normal lift wing is fixedly arranged on the holder; the rear normal lift wing is fixedly arranged on the holder, is located below and behind the front normal lift wing, and forms a tandem wing layout with the front normal lift wing; the propeller is located in front of the front normal lift wing; the rear normal lift wing is provided with ailerons and elevators, and the ailerons and the elevators are electrically connected with the flight controller; the unmanned aerial vehicle with the near-distance coupling tandem wing layout is designed with two wings which can both provide normal lift and are connected in series, and the aerodynamic performance of the whole machine is improved by utilizing the advantageous aerodynamic coupling effect.

The application belongs to the field of aircraft design, and particularly relates to a variable angle tail wing configuration for a high Mach number aircraft, comprising a left wing, a right wing, a left tail wing and a right tail wing; the left tail wing and the right tail wing are arranged in mirror symmetry along a fuselage symmetry plane, the root of the left tail wing and the root of the right tail wing are connected to a tail wing rotating shaft, and the tail wing rotating shaft is rotationally connected to a wing trailing edge; the rotation angle of the left tail wing and the right tail wing is 0° to ±90°; a power unit is connected to a flight control system of the aircraft, and the flight control system of the aircraft can control the left tail wing and the right tail wing to rotate by 0° to ±90° along the tail wing rotating shaft. The tail wing serves as a lifting surface at a low speed stage according to the aerodynamic characteristic requirement at different flight speeds, and increases the lift of the aircraft. The tail wing serves as a heading stability and control surface at a high speed stage, and solves the problem of insufficient heading stability at a high Mach number. A dihedral angle of the tail wing is reasonably selected, and the peak value of wing area caused by the requirement of lift and heading stability in traditional design is avoided.

The application discloses a design method of a wing arm tilting mechanism of a tilt-rotor aircraft, and steps are as follows: combining a connecting rod mechanism principle, a tilting transition working condition rotor six force element and a tilting rudder torque threshold, transmission ratio design is completed; a distributed method and a topological optimization technology are adopted to obtain the best force transmission route and the optimal configuration of a motor support and an end frame support; a tilting transition and a peak torque working condition are taken as loads, and digital simulation and theoretical mechanics are combined to design static strength; a dynamics model is constructed based on vibration characteristics, a multi-rotor and a fixed-wing configuration wing arm vibration spectrum is taken as a load, and dynamic stress is superposed on static stress to design dynamic strength as a checking threshold. The application combines theoretical mechanics, vibration mechanics and digital simulation technology, provides a systematic design route, can realize high-G value resistance, lightweight, high stability and high reliability of the tilting mechanism design, guarantees reliable work of tilting, saves cost and shortens a development period.

The invention relates to the technical field of aircraft design, and discloses a low-noise wing structure based on an acoustic soft material and an aircraft, a porous acoustic soft material layer is arranged between a main wing and a high lift device (slat and flap), and the arrangement of the porous acoustic soft material combines the outer contours of the main wing and the high lift device. The outer contours of the main wing and the high-lift device are the same as those of the original main wing and the high-lift device, the aerodynamic shapes and aerodynamic performance of the original main wing and the high-lift device are not affected, the aerodynamic performance of the wing is not affected, meanwhile, the strength of an aerodynamic noise source of the wing can be further weakened, noise is absorbed in the sound transmission process, and the sound transmission efficiency is improved. Radiation noise generated when the high lift device is unfolded is remarkably reduced, and the flight comfort degree is improved. In addition, the main wing and the high lift device are simple in structure, easy to manufacture and apply and suitable for various aircrafts with fixed wings.

The application discloses a wing robust optimization design method based on Hessian eigenvector dimension reduction, and belongs to the technical field of aircraft design optimization, and comprises the following steps: determining initial design variables corresponding to an initial geometric shape of a wing; updating a wing geometric shape, and then adjusting a fluid calculation grid by using a grid deformation module; performing dimension reduction on an uncertainty space based on a Hessian eigenvector, and calculating aerodynamic performance data and a first-order gradient of sample points; inputting the aerodynamic performance data and the first-order gradient of the sample points into a UQ module, and calculating gradient information of the initial design variables; and judging, by an optimizer, whether a wing configuration determined by the initial design variables is convergent. The application reduces the calculation cost and improves the calculation efficiency while ensuring the accuracy, and provides an efficient technical path for engineering-level aerodynamic robust design under high-dimensional uncertainty conditions.

The present invention relates to an aerodynamic heat recovery and variant coupling device for a high-speed aircraft, comprising: an aerodynamic heat recovery structural unit (1), a speed sensor, a deformable wing frame (2), and a wing configuration control unit (3) connected to the aerodynamic heat recovery structural unit (1), the speed sensor, and the deformable wing frame (2); the aerodynamic heat recovery structural unit (1) is a thermoelectric conversion unit and is arranged in a heat concentration area at the front end of the aircraft, for converting thermal energy into electrical energy; the wing configuration control unit (3) controls the deformable wing frame (2) to change the wing configuration of the aircraft based on the electrical energy generated by the aerodynamic heat recovery structural unit (1) and the aircraft speed input by the speed sensor. This solution has the advantages of simple structure, excellent coupling, and strong practicality.

The invention discloses a flying wing layout course control method based on a distributed passive fluid thrust vector, and belongs to the technical field of aircraft attitude control. According to the method, the distributed jet flow yaw control system is arranged in the spanwise direction of a wing according to the layout form of the flying wing aircraft and yaw control requirements, and jet flow is deflected to an upper / lower wing surface jet flow channel through injection and cut-off of passive secondary flow energy in a secondary flow channel and Coanda wall surfaces on the two sides in an expansion flow channel; according to the wing yaw control system, left and right staggered jet flows are formed in the spanwise direction, protrusions similar to virtual bulge structures are generated, the blocking effect is formed on incoming flow, different virtual bulge structures are formed on wings by setting different yaw control schemes, different blocking effects are formed, and then different course control effects are achieved. The method does not need any mechanical control surface, does not need any additional secondary flow active air source, and has the advantages of fast control response and small torque cross-axis coupling.

The invention provides a distributed multi-propeller tilt rotor aircraft and a transient transition aeroelastic analysis method. The method comprises the following steps: step 1, establishing a coordinate system of a distributed multi-propeller / tilt wing configuration coupling system; 2, establishing a coupling system structure model according to the coordinate system; 3, establishing a coupled system aerodynamic model according to the coordinate system; 4, wake bending correction is carried out on the aerodynamic model of the coupling system; 5, according to the structure model and the aerodynamic model of the distributed multi-propeller / tilt wing configuration coupling system, forming a dynamic overall array of the coupling system; 6, the tilt transition corridor parameters are substituted into the dynamic overall array, and a transient dynamic equation of the coupling system is obtained; and 7, carrying out numerical solution on the transient kinetic equation by adopting a Newmark time finite element method, and solving the aeroelastic response of the transient transition process of the distributed multi-paddle / tilt wing configuration coupling system under the condition of gradually updating the tilt transition parameters and modal parameters.

The invention belongs to the technical field of aircraft design, and particularly relates to a method for determining the takeoff weight of an aircraft after wing area adjustment, which comprises the following steps: step 1, based on an original wing configuration, calculating the wing structure weight Ww0, the task fuel weight Wful0 and the used empty aircraft weight Wkj0 under the original wing configuration to obtain the used empty aircraft weight Wnw without wings; step 2, based on the new wing configuration after wing area adjustment, calculating the weight Ww of the wing structure under the new wing configuration by using the takeoff weight Wto of the airplane in the previous round; 3, based on the new wing configuration obtained after wing area adjustment, the task fuel oil weight Wful under the new wing configuration is calculated through the airplane take-off weight Wto of the last round; 4, calculating the takeoff weight Wto'of the aircraft under the new wing configuration based on the wing structure weight Ww and the task fuel weight Wful under the new wing configuration, the weight Wnw of the used empty aircraft without the wings and the aircraft commercial load Wsz, and 5, obtaining the final takeoff weight of the aircraft under the new wing configuration.

A system of a multi-rotor aircraft that capitalizes on the advantages of fixed wing elements combined with rotary wing structures. The fixed wing elements can help to generate lift once the aircraft is airborne and can thus reduce the need for larger lifting rotors which can allow for longer flight times and distances. Additionally, the systems disclosed herein take advantage of a partial in-wing configuration with a number of rotors to reduce the overall footprint of the vehicle while maintaining the flight efficiency that comes with combining features of fixed and rotary wing elements, and increasing operator safety by shrouding rotating parts. The unique configurations allow for a decoupling of the pitch, yaw and roll authority to reduce the complexity in control systems and improve the flight efficiency of the aircraft. Additional configurations implement the use of smaller thrust rotors that can be used to generate thrust as well as control yaw and thus counteract any remaining unbalanced torque.

The invention discloses a ground-effect wing triphibian carrying platform, and relates to the technical field of triphibian carriers, and the platform comprises a configurable switchable multi-vehicle body module, a variable configuration driving mechanism, a multi-mode propulsion system, a new energy hybrid energy supply system and a comprehensive control unit. The multi-vehicle-body module is formed by connecting a main vehicle body and symmetrical side vehicle bodies through hinges, and the main vehicle body and the symmetrical side vehicle bodies have independent floating capacity. The variable-configuration driving mechanism drives the side vehicle body to turn over by 0-90 degrees through the turning oil cylinder, and the stretching oil cylinder controls the air wing to be folded and unfolded. The multi-mode propulsion system comprises ground, water and flight assemblies and is adaptive to different scenes; the comprehensive control unit coordinates linkage of all the components. The platform can be switched to a land closed configuration, a water surface catamaran configuration and a flight wing unfolding configuration, the air wing unfolding width overtaking length is two times, the performance is improved in cooperation with the ground effect, ground cross-country, water surface high-speed sailing and low-altitude water surface sweeping flight are achieved, the contradiction of triphibian performance is solved, and stability, safety and high efficiency are achieved.

The application discloses a method for winged layout wing multi-objective aerodynamic optimization based on DOE hierarchical sampling, and relates to the technical field of aircraft optimization. Firstly, multi-objective aerodynamic optimization design is carried out based on the straight and level flight state of the winged layout wing system. Secondly, the optimization target is evaluated based on the overall lift-drag ratio of the system, and a multi-objective optimization objective function and constraint condition are constructed. Then, the geometric model of the system is parameterized, the representation parameters of the cruise flight state are decomposed, and optimization variables are selected, so as to construct a multi-dimensional optimization space. Next, according to aerodynamic design experience, the optimization variables are layered and processed according to the significance of the influence on the objective function. Finally, the optimized result is obtained through cyclic iteration and evaluation. The method disclosed by the application solves the problems of high time cost and economic cost caused by too many optimization variables and large optimization space in the optimization design.

The present invention discloses a ground-effect transport aircraft using distributed propulsion, belonging to the field of transportation equipment technology. It features a front-to-rear tandem twin-wing configuration with retractable flaps on the trailing edges of both wings. When extended, the front and rear wings are connected as one, with the rear wing closer to the water surface, enhancing ground effect. Furthermore, the present invention utilizes a hybrid power system, a retractable and foldable flotation device, and a distributed propulsion system based on embedded ducted fans. This allows for highly economical flight while achieving a high payload, efficient cargo storage and handling, and the potential for conversion to passenger transport. This invention boasts a high payload, high controllability, good economy, and adaptability, providing a complementary solution for my country's coastal logistics system.

The invention relates to a parallel loading wing type large-span cantilever truss structure system and a mounting method thereof, the structure system is formed by connecting a plurality of loading wing type trusses which are arranged along the outer edge of a main body structure in parallel at intervals and unfolding wing type trusses on the outermost side in parallel, and the trusses are connected into a space whole through horizontal connecting components. The wing carrying type truss comprises a vertical bearing part and a horizontal overhanging part which are inclined from inside to outside, a wing carrying structure is formed, and overhanging loads are efficiently and obliquely transmitted to a main body structure. Reinforcing components such as three-fork web members are arranged at the ends of the wing-unfolding type trusses, and the rigidity of the ends is enhanced. The mounting method adopts sequential construction of pushing truss by truss from one end to the other end and synchronously connecting horizontal members. The problems that a traditional solid-web cantilever beam is huge in dead weight, insufficient in rigidity and large in building space occupation are solved, and the large-span solid-web cantilever beam has the advantages of being reasonable in stress, small in steel consumption, large in spanning capacity and wide in building space and is suitable for large-span cantilever buildings such as stadiums and terminal buildings.

The application discloses a thrust vector control mechanism suitable for flying wing layout and belongs to the field of aviation technology. The thrust vector control mechanism comprises a machine body, jet outlets, jet flow guide vanes, vector adjusting vanes and a rudder. The machine body is a flying wing layout aircraft with a wing-body fusion. The jet outlets are located on both sides of the trailing edge of the machine body. The jet flow guide vanes are fixed on both sides of the machine body, and the symmetry plane of the jet flow guide vanes is the upper and lower symmetry plane of the jet outlets. The vector adjusting vanes are hinged to the jet flow guide vanes. The output shaft of the rudder is coaxial with the hinge shaft of the vector adjusting vanes, and is fixedly connected with the jet flow guide vanes. The application controls the jet flow and adjusts the deflection of the vector adjusting vanes, so that greater three-degree-of-freedom control moments can be provided for the aircraft, the system design of the aircraft is simplified, and the damage of the rudder to the stealth performance is reduced.