31 results about "Aeroelasticity" patented technology

The invention discloses a rapid static aeroelasticity analysis method for a high-aspect-ratio unmanned aerial vehicle, and the method comprises the steps: firstly, carrying out the measurement of a geometric model, so as to construct an aerodynamic model in the form of a standard data card, and building a simplified beam model of the high-aspect-ratio unmanned aerial vehicle; carrying out aerodynamic solution based on a curved surface vortex grid method and statics solution based on a classical beam theory; and loading an aerodynamic load obtained by solving aerodynamic force into the structural model according to a force translation principle, solving torsion and bending of a beam through a beam theory, mapping structural parameters into aerodynamic parameters, and repeating the operation until calculation is converged. And finally, outputting final displacement, aerodynamic distribution and other results. Compared with a calculation result based on a CFD / CSD high-precision numerical method, the static aeroelasticity calculation result has the advantages that the maximum error of elastic deformation does not exceed 0.3%, and the precision is high.

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 application provides a heavy-duty gas turbine compressor aerodynamic strength vibration structure coupling design method and system. The method comprises the following steps: determining first flow-through parameters and first blade profile parameters of the compressor according to aerodynamic performance indexes; optimizing the first flow-through parameters and the first blade profile parameters of the compressor according to structural manufacturing and assembly indexes and the aerodynamic performance indexes, to obtain second flow-through parameters and second blade profile parameters of the compressor; optimizing blade root geometric parameters of the compressor according to strength vibration life indexes and the structural manufacturing and assembly indexes, and simultaneously designing and optimizing geometric parameters of a blade locking structure of the compressor, to obtain optimized blade root geometric parameters of the compressor and the geometric parameters of the blade locking structure; and designing the heavy-duty gas turbine compressor based on the second flow-through parameters and the second blade profile parameters. The technical scheme provided by the application improves the design efficiency, optimizes the performance balance, and reduces the design cost.

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.

The invention relates to an efficient elastic aircraft large maneuvering process simulation method, belongs to the technical field of aircraft aeroelasticity and flight mechanics coupling simulation, and solves the problems that in the prior art, a high-precision CFD method is low in calculation efficiency, an aerodynamic derivative cannot be rapidly derived, and the calculation efficiency is low. The method comprises the following steps: S1, establishing an average body shaft system, a dynamic equation set and a kinematics equation of the elastic aircraft; s2, performing small disturbance linearization on a rigid-elastic coupling equation of the elastic aircraft, establishing a disturbance equation of the elastic aircraft, and extracting an aerodynamic derivative; s3, constructing a small disturbance model of the elastic aircraft; s4, a state matrix and an input matrix changing along with the incidence angle are obtained in advance; s5, the small disturbance model of the elastic aircraft is popularized under the large maneuvering working condition, and a rigid-elastic coupling model is obtained; s6, the established rigid-elastic coupling model is embedded into a simulation system, simulation calculation is executed, and a simulation result is obtained.

The application discloses a two-stage composite tension adjusting device, system and method for a large-chord-ratio flexible aircraft wing, and belongs to the technical field of aircraft structure design, active control of aeroelasticity and intelligent material structure; the system integrates a first-stage self-locking large-load slow adjusting mechanism and a second-stage lever small-load high-speed adjusting mechanism. In the first stage, a worm gear self-locking principle is used to set and lock a wing basic pre-tension in a high-load and low-power consumption mode, so that quasi-static configuration adjustment adapting to a task is realized. In the second stage, a mechanical gain mechanism such as a lever or a symmetrical cantilever is used to perform slight dynamic modulation on the tension in a high-frequency response and high-efficiency mode, so that dynamic load and flutter suppression are realized. Through function decoupling and cooperation, the application simultaneously gives the flexible wing the ability of macro-configuration optimization and micro-dynamic stability at the lowest system weight and energy consumption cost, and significantly expands the flight performance envelope and task adaptability.

The present invention relates to the field of wind tunnel testing of large wind turbine blades and aircraft wings, and specifically to a two-degree-of-freedom aeroelasticity experimental device, comprising a wind tunnel test model, a two-degree-of-freedom support mechanism, and a signal acquisition system; the wind tunnel test model comprises a wing section model, a pressure measuring tube, an end plate, a central torsion shaft, a locking mechanism, and a center of mass adjustment mechanism; the two-degree-of-freedom support mechanism comprises a wind tunnel fixing plate, a wind tunnel anti-spoiler cover, a support plate, a linear bracket, an angular displacement sensor fixing flange, a torsion shaft fixing rod, a torsion support rod, a smooth linear straight rod, a sinking and buoying elastic mechanism, a pitch linear spring, a pitch spring connecting column, a linear sliding bearing, an outer spherical bearing, a sinking and buoying limit device, and a pitching limit device. The present invention can be used to study the aeroelastic characteristics of a wind tunnel test model under two-degree-of-freedom support conditions of pitch and sinking, that is, to provide an experimental device for exploring phenomena such as flow separation on the wing section surface, dynamic stall, and flutter during aeroelasticity testing.

The invention relates to a flutter analysis modal tracking method based on modal confidence, and belongs to the technical field of aeroelasticity. According to the flutter analysis modal tracking method, modal eigenvalues and right eigenvectors are matched and tracked in combination with the modal confidence and eigenvalue continuity under different airspeeds, and the flutter analysis modal tracking method based on the modal confidence is obtained. Therefore, the correctness and robustness of the corresponding relation between the characteristic values are maintained, the problems of large calculation amount, insufficient precision and failure under specific conditions of a modal tracking method in the prior art are solved, and the precision and reliability of modal tracking are improved.

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.

The invention relates to a blisk equivalent aeroelasticity test system. The blisk equivalent aeroelasticity test system comprises a calculation-acquisition-control system and a vibration exciter-structure system, the calculation-acquisition-control system comprises an aerodynamic force calculation module, a data acquisition module and a centralized aerodynamic force loading control module; the vibration exciter-structure system comprises an eddy current displacement sensor and a test bed; the eddy current sensor collects displacement signals of specified collection points of all the blades in real time, converts the displacement signals into voltage signals and inputs the voltage signals into computer hardware of the aerodynamic force calculation module through a data transmission line. According to the method, a displacement acquisition and input-real-time calculation-aerodynamic force output system is constructed, displacement signals of all blades are acquired by utilizing an eddy current sensor, real-time aerodynamic force is calculated by utilizing real-time calculation hardware and an aerodynamic force reduced-order model, and the real-time aerodynamic force is output to a blisk through a vibration exciter; and equivalent simulation of the fluid-solid coupling effect of the blisk in the inner flow field is realized.

The invention discloses a method and a device for realizing aeroelasticity analysis of a gas compressor, and the embodiment of the invention relates to a method for analyzing aeroelasticity of a gas compressor based on one-dimensional and three-dimensional modeling, which greatly improves the analysis efficiency of the vibration characteristics of a blade under a surge condition. By optimizing the modeling process and reducing the calculation cost, the surge problem of the gas compressor is recognized more quickly, optimization is conducted in the design stage, and the reliability and stability of the gas compressor are improved.

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 present invention relates to a method for efficiently obtaining gust loads of an unmanned aerial vehicle (UAV), belonging to the technical field of aircraft and aeroelasticity. The present invention first determines a nonlinear wing beam model with strain as a variable, and obtains a cross-sectional constitutive relationship matrix of each beam cross section based on the nonlinear wing beam model; secondly, a nonlinear static aeroelastic model is constructed based on an unsteady vortex lattice method, gust disturbances are introduced into the model, and aerodynamic forces corresponding to each beam section are obtained, and then a numerical solution is performed based on the nonlinear beam model to obtain corresponding updated strain quantities; finally, the wing section of each beam section is determined, and the corresponding wing section load, i.e., the gust load, is obtained based on the updated strain quantities of each beam section and the corresponding cross-sectional constitutive relationship matrix.

A two-degree-of-freedom three-dimensional aeroelasticity measurement device comprises: a four-component balance, a blade flapping mechanism, a blade pitching mechanism and an end plate structure, wherein: the four-component balance is arranged at the bottom end of the interior of a wind tunnel test section and is connected to the blade flapping mechanism outside the wind tunnel test section; the blade flapping mechanism is located between the four-component balance and the blade pitching mechanism to achieve left and right tilting operation; the blade pitching mechanism is arranged below the blade flapping mechanism to achieve pitching operation; and the end plate structure is arranged at the top of the wind tunnel test section to suppress blade tip vortices. The present invention can achieve high-precision measurement of the nonlinear aerodynamic load, pitch angle change response and flapping angle change response of a forced oscillating blade in a pitch-flapping two-degree-of-freedom three-dimensional aeroelastic system. The present invention has a convenient spring replacement function, can flexibly change the stiffness coefficient in the flapping direction, and can perform parametric simulation of special resonance phenomena such as combined resonance and sub-resonance.

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 present invention proposes a method for hybrid optimization design of the number, size, and angle of the radial structure of an aircraft considering aeroelasticity. Based on the actual working environment of the aircraft, constraints such as aerodynamic heat, allowable stress of materials, and structural weight are comprehensively considered. At the same time, hybrid optimization of the number, size, and angle of the structure is achieved, and the correlation factors between various optimization objectives are considered, which is in line with engineering practice. In addition, for the traditional optimization process based on the basic structure method, the layout of the determined basic structure will not be adjusted, which requires a high level of initial basic structure layout design. In the optimization process of the present invention, the angle variable of the basic structure is introduced as an optimization variable, which reduces the difficulty of the initial layout of the basic structure.

The invention relates to a high-aspect-ratio unmanned aerial vehicle wing flutter suppression method considering wingtip deflection angle influence, belongs to the technical field of aeroelasticity of aircrafts, and solves the problems that in the prior art, flutter critical speed prediction is not accurate, and passive flutter suppression measures are low in efficiency. And the flutter energy cannot be counteracted in real time. According to the high-aspect-ratio unmanned aerial vehicle wing flutter suppression method, the coupling influence of a structural power subsystem and an aerodynamic subsystem is considered, the critical speed of high-aspect-ratio wing flutter with a wingtip deflection angle can be predicted, and a relational expression between a control surface control signal and flutter control influence factors is obtained; the method has a good effect of suppressing the flutter of the wing of the high-aspect-ratio unmanned aerial vehicle considering the influence of the wingtip deflection angle, and meanwhile, the theoretical method provided by the invention is simple and clear and is convenient for engineering application.

The invention relates to an aerodynamic force prediction method based on a homotopy theory and a modified surface element method, belongs to the field of aerodynamics and aeroelastic mechanics analysis, and solves the defect that an aerodynamic force prediction tool in existing aeroelastic analysis cannot meet the requirement of difficulty in balancing calculation efficiency and calculation precision. Compared with a traditional CFD calculation method, the method has the advantages that the calculation efficiency is high, the form is simple, the method can be easily coupled with a structural equation, and the overall efficiency of aeroelasticity analysis can be effectively improved. Besides, in the process of constructing the homotopy aerodynamic model, different weighting functions are selected, so that the method disclosed by the invention has more flexible and abundant constitution forms, and aerodynamic nonlinear factors can be taken into consideration.

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 present invention discloses a method for attitude-elastic-aerodynamic coupled flutter control of an aircraft with a flying wing layout, which relates to technical fields such as mechanics, aeroelasticity, and control. The control method includes: integrating the Lagrangian equation in the general body axis system to establish a full-coupling dynamic model, and through small perturbation linearization, establishing a dynamic equation of perturbed motion for flight control system design; based on system identification flight tests, correcting the dynamic equation of perturbed motion; designing an attitude-elastic-aerodynamic coupled flutter suppression controller based on traditional or modern control theory and embedding it into the flight control system; verifying the effectiveness of the flutter suppression controller through flight tests. The present invention significantly solves the problems that existing numerical analysis and wind tunnel experiments cannot accurately reflect the complex coupling effects during flight, cannot effectively suppress the occurrence of flutter during flight, and cannot expand the flight envelope, etc., and can effectively improve the flight safety and stability of aircraft with a flying wing layout.

The application provides a center-of-mass-adjustable anti-chattering air rudder and a manufacturing method. A counterweight of a first mass is riveted and fixed at a position before a pneumatic pressure center of a skeleton of the air rudder; strength analysis and chattering analysis are performed on a rudder surface of the air rudder to obtain a strength margin and a chattering margin of the current rudder surface; if the margin is sufficient, the mass of the counterweight is reduced to a second mass to reduce the weight of the rudder surface, and then the strength analysis and the chattering analysis are performed again until the strength margin and the chattering margin can meet the design requirements; if the margin is insufficient, the mass of the counterweight is increased to a third mass to increase the weight of the rudder surface, and then the strength analysis and the chattering analysis are performed again until the strength margin and the chattering margin can meet the design requirements. The application adjusts the center-of-mass position of the rudder surface to improve the aeroelasticity and avoid the occurrence of the chattering problem under the premise that the air rudder has sufficient thermal strength.

The present invention provides a design optimization method for the aeroelastic stability of a wind turbine blade. First, a segmented model of a complete blade including the aerodynamic profile and the detailed structural ply is established; then, the overall physical parameter characteristics of the blade are calculated through the blade parameter model to obtain the corresponding mechanical properties of the blade, and the distributions of the torsional stiffness and the flapwise stiffness of the blade are output; at the same time, through the wind turbine whole machine model, the rated operation condition of the wind turbine is simulated and calculated to obtain the variation curves of the rotational speed and the pitch angle; then, the stability of the wind turbine rotor aeroelasticity is calculated, and according to the results, the distributions of the torsional stiffness and the flapwise stiffness are optimized and improved through different optimization curves to improve the rotor stability; finally, the above-mentioned cyclic optimization steps for the torsional stiffness distribution of the blade are repeated to coordinate the coupling relationships between various objectives and between constraints and objectives, and to promote the modification of its torsional stiffness and flapwise stiffness variables to converge to the optimal solution or the optimal solution set of the aeroelastic stability of the rotor.

The application discloses a fan design method and system for a small and medium flow civil turbofan engine, which simultaneously carries out sensitive parameter analysis in three aspects of flutter, aerodynamics and noise after determining a preliminary design scheme of the fan, and respectively determines flutter suppression optimization design parameters, aerodynamic optimization design parameters and noise reduction optimization design parameters, so that sensitive parameters of aerodynamic optimization, flutter optimization and noise reduction optimization can be accurately identified, targeted parameter adjustment and optimization are facilitated, and optimization efficiency is improved. Moreover, overall optimization design is comprehensively considered in the aerodynamic optimization, the flutter suppression optimization and the noise reduction optimization to determine a new optimization design scheme, flutter stability evaluation, aerodynamic performance evaluation and aerodynamic noise evaluation are simultaneously carried out based on the optimization design scheme, parallel coupling iterative design of three disciplines of aerodynamics, aeroelasticity and acoustics is further improved, design efficiency is further improved, and a fan design scheme with comprehensive optimization of aerodynamic performance, aeroelastic stability and noise level can be obtained.

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.