21 results about "Wing configuration" patented technology

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.

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 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 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 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 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 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 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.

The invention discloses a vertical take-off and landing aircraft with tandem tilting wings and a hybrid power control method of the vertical take-off and landing aircraft, and belongs to the technical field of advanced air traffic. The aircraft comprises a tandem tilting wing structure, wherein a plurality of power units are distributed on a front wing and a rear wing and integrally tilt together with the units; the hybrid power system is integrated in the fuselage and comprises a hydrogen fuel cell system matched with cruise steady-state power and a lithium battery system matched with take-off and landing transient peak power; and the energy management controller is in signal connection with the hybrid power system and the tilt angle sensor, and is configured to dynamically adjust the power output ratio of the hydrogen fuel cell to the lithium battery in response to the wing tilt angle change. According to the invention, the power system and the tandem tilting wing configuration are deeply cooperated, so that the breakthrough improvement of the flight range is realized, the millisecond-level power response and safety redundancy in the transition flight stage are ensured, and the service life of the lithium battery is obviously prolonged through shallow charging and shallow discharging management of the lithium battery.

This invention discloses a control method for an adaptive deformable wing based on Hessian vector product, belonging to the field of aircraft design technology. The method includes: S1, defining the wing's design variables; S2, updating the design variables using an optimizer based on gradient information; S3, updating the wing's geometry using the FFD geometric parameterization method and adjusting the fluid computation grid; S4, calculating the aerodynamic performance data and first-order gradient of sample points using the fluid computation grid after perturbation of the design variables; S5, calculating the robust objective function and perturbation vector, and calculating the total gradient using the Hessian vector product module; S6, determining whether the wing configuration determined by the design variables has converged based on the robust objective function and the total gradient. If converged, the wing configuration is taken as the final robust optimal wing design; otherwise, the total gradient is used as gradient information, and the process returns to S2. This achieves a balance between efficiency and accuracy, ensuring the reliability and economy of the aircraft in actual operation.

A tail sitting VTOL aircraft with nonplanar tandem rotor blown wing configuration, capable of traveling in an airplane mode with its fuselage oriented horizontally, and a hover mode during take-off and landing with its fuselage oriented vertically, with capability to have precise controlled hover, and capability of making controlled and safe assisted transition between two modes during a horizontal movement and without need for much headroom and overhead clearance. Transition from hover mode to airplane mode is performed by moving forward in hover mode and rotating the fuselage around the pitch axis by the assist of the moment created by differential thrust of the propellers, or the moment created by differential lift created by the rotor blown tandem airfoil-shaped cross section bodies or a combination of both.

A vertical landing and take-off aircraft VTOL transitions from a vertical takeoff state to a cruise state where the vertical takeoff state uses propellers to generate lift and the cruise state uses wings to generate lift. The aircraft has an M-wing configuration with propellers located on the wingtip nacelles, wing booms, and tail boom. The wing boom and / or the tail boom can include boom control effectors. Hinged control surfaces on the wings, tail boom, and tail tilt during takeoff and landing to yaw the vehicle. The boom control effectors, cruise propellers, stacked propellers, and control surfaces can have different positions during different modes of operation in order to control aircraft movement and mitigate noise generated by the aircraft.

The invention discloses a blade tip jet rotor wing driven by a combined type ion wind propeller. The blade tip jet rotor wing comprises a rotating main shaft, rotor wing blades installed on the main shaft and a power supply system. The improvement is that two types of ionic wind propellers are integrated on the rotor blade, namely a needle-circular truncated cone sleeve configuration propeller which is mounted at the tip end of the blade and is provided with a nozzle facing the rotating tangential direction, and a line-wing configuration propeller which is arranged along the wingspan direction of the blade. The power supply system transmits electric energy to the high-voltage generation module rotating along with the main shaft through the collector ring, and then high-voltage excitation is provided for the two types of propellers. After electrification, the propellers at the propeller tips generate tangential ion wind jet flow to provide main driving torque, and the propellers in the wingspan direction generate ion wind flowing along the airfoils to provide auxiliary thrust and additional lift force and implement active flow control. Mechanical transmission parts and pneumatic / gas pipelines are omitted, a rotor wing driving structure is simplified, and the propelling efficiency and pneumatic performance of the system are improved through cooperation of the double-configuration propellers.

The invention relates to an adjusting method and mechanism for stable flight during variable sweep-forward of wings of an aircraft, and the method comprises the steps: obtaining the rotation angle and the flight speed of the wings during the flight of the aircraft, and obtaining a focus position interpolation table according to the rotation angle and the flight speed; calculating the real-time position of the focus of the aircraft by adopting a linear interpolation algorithm; calculating the target position of the center of gravity of the whole aircraft through the real-time position of the focus of the aircraft and the expected static stability margin value, and calculating the target position of a center-of-gravity adjusting slide block in combination with the mass of the whole aircraft; and the real-time position of the gravity center adjusting sliding block is obtained, the gravity center adjusting sliding block is made to move to the target position of the gravity center adjusting sliding block according to the difference value between the target position of the gravity center adjusting sliding block and the real-time position of the gravity center adjusting sliding block, and stable flight of the aircraft wing during forward sweep changing is achieved. The problem of adjustment deviation easily occurring in traditional control is avoided, the gravity center adjustment precision and response timeliness are improved, and the gravity center of the aircraft can dynamically adapt to wing configuration changes and flight speed changes.

The invention provides a double-shaft variable-sweep upper single-wing layout of a speed-domain-crossing patrolling missile, which comprises a double-shaft variable-sweep wing mechanism arranged on the back of a fuselage and can form an M-shaped integral structure; each double-shaft variable-swept-wing mechanism comprises an inner wing section and an outer wing section, and the wing root of the inner wing section is pivotally connected with the fuselage, so that the sweep-forward angle of the inner wing section is adjusted; the wing root of the outer wing section is pivotally connected with the wing tip of the inner wing section, so that the sweepback angle of the outer wing section is adjusted. In the embodiment, the angles of the inner wing section and the outer wing section can be independently changed, the left side and the right side have asymmetric deformation capacity, and the overall aerodynamic configuration is changed to adapt to different flight speeds from low subsonic velocity to supersonic velocity. The upper single-wing layout is adopted, and space can be released for placement of an air inlet channel and task loads. The high-aspect-ratio wings bend and deform upwards to deviate from the fuselage under the aerodynamic load, and the problem of structural interference of the wings and the fuselage caused by aerobombs is solved. Most airfoils and the whole variable-sweep rotating mechanism are located in the leeward shadow area of the fuselage, so that the stealth performance is improved, and the aerodynamic heating problem of the deformed airfoils during hypersonic flight is relieved.

1. Name of the product in this design: Aircraft (Compound Wing Configuration Manned Electric Vertical Take-Off and Landing). 2. Purpose of this design: For use in manned electric vertical takeoff and landing aircraft. 3. The key design features of this product are its overall shape. 4. The image or photograph that best illustrates the design's key points: a 3D model.

This application discloses a method and mechanism for stabilizing flight of an aircraft during wing forward sweep. The steps include: acquiring the aircraft's rotation angle and flight speed during wing forward sweep; calculating the real-time position of the aircraft's focus using a linear interpolation algorithm based on the rotation angle, flight speed, and a focus position interpolation table; calculating the target position of the aircraft's center of gravity using the real-time position of the focus and the desired static stability margin value; calculating the target position of the center of gravity adjustment slider based on the aircraft's mass; acquiring the real-time position of the center of gravity adjustment slider; and moving the slider to the target position based on the difference between the target position and the real-time position, thereby achieving stable flight during wing forward sweep. This method avoids the adjustment deviation problems common in traditional control systems, improves the accuracy and responsiveness of center of gravity adjustment, and enables the aircraft's center of gravity to dynamically adapt to changes in wing configuration and flight speed.