13 results about "Aeroelasticity" patented technology

This invention relates to a method for aircraft angle-of-attack trim considering nonlinear aeroelastic effects, belonging to the fields of aerodynamics and aeroelasticity. Addressing the need for angle-of-attack trim in aircraft load calculation, this invention compares the trim results based on high-precision fluid dynamics (CFD) and solid dynamics (CSD) nonlinear aeroelastic calculations, employs a bisection method to adjust the trim angle of attack, iterates repeatedly until the trim converges, and ultimately achieves angle-of-attack trim considering nonlinear aeroelastic effects. This enables accurate calculation of the aircraft trim angle of attack, meeting the high-precision load calculation requirements in modern aircraft calculations.

The invention discloses a synchronous dynamic measurement method for a surface boundary layer and an interstage flow field of an aero-engine rotor blade, and belongs to the technical field of mechanical testing of aero-engine / gas turbine impellers. According to the method, fluid-solid coupling numerical simulation is carried out on a to-be-tested rotor system under a given rotating speed and a given working condition, and arrangement positions and installation postures of a surface hot film, a hot wire probe or a dynamic probe and a strain gauge are determined; a synchronous measurement system composed of a surface hot film measurement channel, a hot wire / dynamic probe measurement channel, a strain gauge measurement channel, a synchronous phase locking device, a high-fidelity slip ring and a multi-channel data acquisition system is constructed; and synchronous dynamic measurement and conjoint analysis of a blade surface boundary layer, an interstage flow field and a blade vibration signal are realized. According to the method, high-precision three-field coupling data can be obtained, and a reliable basis is provided for fluid-structure interaction mechanism research and aeroelasticity and fatigue life design of a high-load aero-engine.

The invention discloses a rigid-elastic coupling aeroelasticity control flight test platform, and relates to the technical fields of unmanned aerial vehicles, mechanics, aeroelasticity, control and the like. The rigid-elastic coupling aeroelasticity control flight test platform comprises a double-triangle high-aspect-ratio sweepback tailless flying wing pneumatic structure with a rigid-elastic coupling effect, a lightweight aircraft structure design and aeroelasticity control airborne equipment. The lightweight aircraft structure design comprises fuselage structure design, wing structure design and fuselage-wing mounting mechanism design; the aeroelastic control airborne equipment comprises control equipment, communication equipment and power equipment. The rigid-elastic coupling aeroelasticity phenomenon can be obviously captured and effectively inhibited in a flight test, a reliable test platform is provided for rigid-elastic coupling aeroelasticity mechanism analysis and inhibition technology verification, and a foundation is laid for deep research of an aeroelasticity control technology.

The invention relates to a propeller wing structure aeroelasticity experimental simulation device with an adjustable propeller wing spacing. The device comprises an engine bracket, an engine, a pitching bearing seat, a yawing bearing seat, a first nacelle beam, a yawing fixing seat, a pitching spring piece, a yawing spring piece, a pitching spring support, a balancing weight, a third nacelle beam, a hanging frame, a propeller, a deep groove ball bearing and a universal ball bolt. The device has the advantages that the distance between the propeller plane and the front edge of the wing can be adjusted by adjusting the installation position of the hanging frame sliding block on the third nacelle beam, and then the influence of the distance between the propeller plane and the front end of the wing on the downward washing effect is analyzed. The length of the propeller-pod pylon can be adjusted and controlled by adjusting the butt joint hole positions of the bottom end of the pylon beam and the upper end of the pylon beam, and then the influence of the length of the pod pylon on the aeroelastic stability of the propeller-pod-wing system is analyzed.

The invention relates to a nonlinear structure reduced-order model construction method based on sparse recognition and a mixed mode, belongs to the technical field of structural dynamics analysis and aeroelastic mechanics analysis, and solves the problem that complex motion caused by geometric nonlinearity under large deformation cannot be accurately described in the prior art. Comprising the following steps: S1, establishing a nonlinear finite element model of a target large flexible wing to obtain a training data set; s2, solving a mixed modal basis based on the displacement residual error and SVD (Singular Value Decomposition); s3, establishing a nonlinear stiffness coefficient solving problem model, introducing LASSO regression to establish an LASSO regression optimization objective function, and solving to obtain a sparse nonlinear stiffness coefficient; s4, based on the sparse nonlinear stiffness coefficient and the structural kinetic equation, establishing a nonlinear structure reduced-order model; and S5, applying the nonlinear structure reduced-order model to statics response solution and dynamics response solution of the large flexible wing to obtain statics response and dynamics response results.

The application discloses a method for suppressing transonic buffeting load and response based on agent cooperation, and relates to the technical field of aeroelasticity, and aims at solving the problem of transonic buffeting of a wing of an aircraft. The method comprises the following steps: firstly, calculating the elastic wing tip acceleration response based on a one-way fluid-solid coupling method; secondly, configuring a multi-target reinforcement learning agent, taking the variable-camber trailing edge and the telescopic micro-bulge as control mechanisms, taking the maintenance of the aerodynamic lift-drag ratio and the suppression of the vibration response of the structure as targets, and letting the agent adopt a proximal policy optimization algorithm to be trained to obtain an optimal control law capable of cooperatively driving the telescopic micro-bulge and the variable-camber trailing edge; and thirdly, deploying the trained cooperative control law on the target wing, and observing the suppression effect of the wing buffeting through quantitative evaluation of the slowing rate. The application trains the agent through a deep reinforcement learning method, so that the agent can guide the telescopic micro-bulge to suppress the diffusion in the separation zone, reduce the shock motion amplitude, drive the variable-camber trailing edge to regulate the flow at the trailing edge and compensate the overall lift fluctuation, and cooperatively suppress the buffeting load and response.

The invention discloses a bidirectional clearance coordination simulation device for a control surface aeroelasticity test. The bidirectional clearance coordination simulation device is composed of an inner side protection device (A), a radial clearance coordination simulation device (B1), an axial clearance coordination simulation device (B2) and a control surface structure (C). The function of the inner side protection device (A) is that the radial clearance coordination simulation device (B1) is fixedly connected with a wind tunnel structure through the supporting connecting shafts (1) and the clamping plates (2) at the two ends. The radial clearance coordination simulation device (B1) has the function of realizing quantitative simulation of the radial clearance through matching of shaft sleeves (8) with different thicknesses and circular ring clamping grooves (10). The function of the axial clearance coordination simulation device (B2) is that quantitative simulation of the axial clearance is realized through cooperation of hole sleeves (6) with different thicknesses, through holes (5) and positioning pins (7). The control surface structure (C) serves as a wind tunnel test object, and a test piece is connected with the gap coordination simulation device through an outer side clamping plate (11) and a clamping plate (12) at the root of a control surface. In the testing process, shaft sleeves (8) with different thicknesses are replaced to be matched with the circular ring clamping grooves (10), hole sleeves (6) with different thicknesses are replaced to be matched with the through holes (5) and the positioning pins (7), coordinated quantitative simulation of the radial clearance and the axial clearance is achieved, and quantitative control over the radial clearance and the axial clearance of the full-motion rudder system is achieved.

This invention discloses a method for analyzing the coupled influence of structural deformation and detection performance on sensor aircraft, belonging to the fields of aircraft design and aeroelasticity. The specific steps of the method are as follows: establishing an aerodynamic and structural model of the aircraft; establishing state-space aerodynamic equations; defining a flight dynamics model; obtaining rigid-elastic coupled dynamic equations based on the Lagrange equations; performing aeroelastic dynamic response calculations through coupled solution to obtain the physical deformation of the wing structure; and calculating antenna performance based on this deformation to analyze the impact of structural deformation on detection performance. This invention, by establishing a rigid-elastic coupled aeroelastic model, achieves a full-process coupled analysis from aeroelastic excitation to structural response and then to antenna performance changes. It solves the problem that traditional methods cannot accurately consider the impact of structural deformation on detection performance under actual flight conditions, providing a more accurate analytical means for the integrated design of sensor aircraft.

This invention discloses an intelligent optimization algorithm for curved fiber wing that integrates physical information buckling proxy, relating to the field of aircraft design technology. The method includes: equivalently planarizing the spatially twisted quadrilateral; deriving the Jacobian matrix through a double-triangle affine mapping to construct a variable stiffness total potential energy functional; embedding the mechanical control equations using a physical information neural network (PINN) to quickly solve for the buckling critical load factor and modes, replacing traditional finite element calculations; parameterizing the fiber path and thickness to construct design variables; and employing a multi-stage adaptive differential evolution algorithm for global optimization, integrating process, static / dynamic aeroelasticity, and strength constraints. This invention achieves a closed-loop modeling-evaluation-optimization process, significantly reducing computational costs while maintaining mechanical consistency, solving the convergence problem of high-dimensional variable coupling, and greatly improving the automation and engineering efficiency of curved fiber wing design.

This invention belongs to the field of aircraft aeroelasticity, and specifically discloses a method for analyzing the spiral flutter of multirotor aircraft. It is designed for the configuration characteristics (hingedless small-diameter propeller) and flight modes (vertical takeoff and landing, large pitch angles in forward flight) of multirotor aircraft. The 2-DOF model in the preliminary design stage does not require complex modeling and software operation, supports rapid adjustment of structural, overall, and aerodynamic parameters, and can complete the analysis of multiple sets of parameters in a short time, meeting the needs of rapid iteration in the preliminary design. The 2-DOF (rapid analysis) and 4-DOF (precise verification) models can be flexibly switched according to the design stage, taking into account both analysis efficiency and calculation accuracy, and adapting to the needs of the entire process from preliminary design to detailed design.

The invention discloses a seaplane buoy fault state divergence analysis method, which comprises the following steps of: 1, establishing a finite element structure model and an aerodynamic model of wings, buoys and supporting rods in a normal state according to the actual shapes of the wings and the buoys of a seaplane and the movement of the supporting rods; step 2, establishing a buoy fault state model based on a fault state encountered by the stay bar; 3, on the basis of the buoy fault state model, the divergence speed of the buoy in the fault state is solved through a static aeroelasticity method; 4, for the buoy fault state model of which the divergence speed exceeds the aeroelastic stability speed envelope, calculating a buoy divergence motion mode; 5, determining a main supporting rod for supporting the buoy in the divergent motion mode of the buoy; calculating the maximum deformation of the main supporting rod in the pitching direction under the maximum damage bending moment; and judging the interference condition of the stay bar with the wing after the stay bar elastically deforms based on the maximum deformation. According to the invention, the flight safety of the airplane can be evaluated, so that the safety and reliability of the seaplane are improved.

The invention relates to the field of wind tunnel equipment, and discloses a continuous transsonic wind tunnel system based on a heavy gas medium and a use method thereof.The system comprises a wind tunnel body, a heavy gas injection module is arranged at the gas inlet end of the wind tunnel body, and a heavy gas recovery module is arranged at the gas exhaust end of the wind tunnel body. The injection module sequentially comprises a heavy gas storage tank, a delivery pump, a vaporizer and a pressure regulating and gas distributing unit; the recovery module sequentially comprises a vacuum pump set, a compressor, condensing equipment, a mixed gas storage tank, cold box freezing equipment and adsorption equipment. Heavy gas is adopted to replace traditional air to serve as a test medium, the structural dynamic characteristics of an aircraft aeroelasticity model can be accurately simulated, the test safety is remarkably improved, energy consumption is reduced, and the requirement for a future large military and civil aircraft full-aircraft aeroelasticity ground simulation test with the high composite material proportion is met.

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.