1026 results about "Aerodynamic force" patented technology

The invention relates to an autonomous energy-saving soaring flight path planning method for a small-sized low-cost aircraft, belongs to the technical field of aircraft trajectory planning, solves the problem of low-cost wind field energy acquisition of the small-sized low-cost aircraft in the prior art, and comprises the following steps: S1, configuring a sensor for the aircraft, and measuring through the sensor to obtain observation parameters; s2, establishing a state vector of the aircraft; s3, establishing an aerodynamic force model, introducing a dynamic equation and a state transition equation, and performing accurate modeling on aerodynamic force; s4, performing multi-source data fusion by adopting extended Kalman filtering, establishing an extended Kalman filter of a nonlinear system, and executing real-time wind vector high-precision sensing; and S5, performing global wind field modeling, estimating a wind field environment, and performing energy-obtaining flight path planning to obtain an optimal energy-obtaining soaring flight path planning scheme.

A method of using computer implemented neural network for simulation of aerodynamic performance of technical object having geometry, method includes: training neural network using plurality of sets of pre-computed computational fluid dynamics encodings, CFD, outputs, wherein training is generated using inputs including: geometry of at least one training technical object; spatial locations of neural network input nodes as node attributes; relationship between geometry of at least one training technical object and neural network input node locations; associated boundary conditions; operating conditions; and computed outputs including flow fields and aerodynamic performance parameters; training using loss function evaluating error between neural network and pre-computed CFD outputs to produce trained neural network; using trained neural network with new inputs to generate as output predicted aerodynamic performance of technical object, wherein relationship between geometry of at least one training technical object and neural network input node locations includes vector with at least two parameters.

The invention discloses a multi-dimensional aeroelastic calculation method and device based on a nonlinear Euler-Bernoulli beam and a storage medium, and belongs to the technical field of wind turbine aerodynamics, the method comprises the following steps: inputting the stiffness of each section of a blade into a nonlinear Euler-Bernoulli beam equation to obtain a modal-displacement transformation matrix; the rigidity, the weight and the external force of each section of the blade are converted to obtain a generalized rigidity matrix, a generalized mass matrix and a generalized force; starting circulation, substituting into a forced coupling vibration equation, solving modal displacement of each order at the current moment, and inputting into a modal-displacement inverse transformation matrix to obtain blade displacement; the generalized force is substituted into a blade element momentum theoretical formula to solve aerodynamic force at the next moment, the generalized force is updated to obtain updated generalized force, and aerodynamic force at each moment is obtained; and calculating to obtain the multi-dimensional aerobomb. According to the method, the calculation of the blade deformation and the load considers the change of the torsional degree of freedom, and more accurate and more efficient load and deformation calculation can be carried out.

The invention provides a new class of transportation with a partially aerostatically supported ram air cushion ship (PASRACS) that can provide safe, fast, efficient global transport services with extraordinary comfort, luxury and amenities. The PASRACS employs inventive synergistic combinations of lift from aerostatic, aerodynamic, hydrostatic and hydrodynamic forces for different modes of operation ranging from a stationary floating configuration on a water surface, to flight in ground effect at high speed. A propulsion system can utilize hydrogen as an energy source driving fluid-dynamic thrusters, to enable zero carbon emissions operations. A transition method is provided for a PASRACS to transition from a floating mode to a takeoff mode to a flight in ground-effect mode in an inventive optimized manner. A transport method for multimodally transporting payload is provided with PASRACS vehicles and systems, with quick turn time operations enabled by the use of payload transfer transport modules and transfer vehicles.

The invention discloses an unmanned helicopter control method based on a fuzzy neural network extended state observer, and belongs to the technical field of transverse unmanned helicopter flight control, and the structure of a transverse unmanned helicopter flight controller comprises an FNNESO-RBSC controller, a control strategy and a transverse unmanned helicopter flight dynamics model; the FNNESO-RBSC controller calculates transverse, longitudinal, course and vertical channel control quantities according to system input and output information, the control quantities are distributed to all control surfaces of the transverse unmanned helicopter through a control strategy and act on a controlled object of the transverse unmanned helicopter, object state response of the transverse unmanned helicopter is changed under the action of aerodynamic force, and the control surfaces of the transverse unmanned helicopter are controlled. And the target flight state is controlled. According to the method, the fuzzy neural network is adopted to transform the extended state observer, and the adaptability of the extended state observer is improved, so that the unknown total disturbance needing to be compensated in the backstepping control law is more effectively obtained.

The invention discloses a wind turbine blade flutter characteristic calculation method based on joint simulation, and relates to the technical field of wind power generation. The method comprises the following steps: S1, reading structural body data of a wind turbine blade, and extracting parameters such as a spanwise position, an airfoil profile coordinate and aerodynamic characteristics of the blade; s2, full-automatic parametric modeling of the blade is achieved through a parametric modeling tool, and a related input file is generated; s3, based on an unsteady blade element momentum theory, aerodynamic force and torque responses of the blade at different spanwise positions are calculated step by step; s4, standardized calculation steps are packaged through a programmed method, personal errors are reduced, and calculation efficiency is improved; s5, designing a cabin interface, a tower interface and a floating platform interface, and realizing flexible data interaction and expansion among modules; and S6, dynamic response analysis is combined, and flutter characteristics of the wind turbine in a complex load environment are simulated. According to the method, the calculation process of wind turbine blade flutter characteristic analysis is accelerated, and the analysis precision and the calculation efficiency are improved. Through programmed and modular design, the flutter simulation of the wind turbine blade can be quickly and accurately completed, and a scientific basis is provided for structural optimization of the blade and safety design of a wind turbine system.

The invention belongs to the technical field of multi-component aerodynamic load measurement sensors, and discloses a method for applying a constant force to a half-mode balance calibration system. The method for applying the constant force comprises the following steps: installing the constant force value loading device; accurate force value loading is carried out; carrying out displacement real-time compensation; constant force value output is kept at all the force application device point positions; and completing the ground static calibration of the half-mode balance. According to the method for applying constant force, the displacement of the electric cylinder is calculated according to the value of the applied force, after the corresponding displacement of the electric cylinder is completed through closed-loop control of the driving electric cylinder, the steel spring inevitably generates creep displacement in the process of waiting for arrival of a plurality of force application device point positions, and the creep displacement is changed into the creep displacement. The applied force value corresponding to the creep deformation is synchronously measured by the high-precision force sensor, the compensation displacement of the piezoelectric ceramic is correspondingly calculated, the creep deformation displacement is dynamically compensated in real time, the constant force value loading device is ensured to be always at the required constant force value, and the high-precision calibration of the half-mode balance is realized.

The invention relates to an aerodynamic force-thermal-structure multi-field coupling intelligent prediction method for a high-speed aircraft, and belongs to the technical field of high-speed aircraft design. According to the flight path of the aircraft, sample parameters used for calculating aerodynamic force and aerodynamic heat states are extracted; modeling the aircraft, and carrying out analog calculation to obtain an aerodynamic force and aerodynamic heat data set; calculating an aircraft temperature field along the trajectory; training the aerodynamic force data set to obtain an aerodynamic force prediction model; calculating structural stress displacement field data of the aircraft model; extracting concerned aircraft feature point corresponding data from the aerodynamic data set, the aircraft temperature field and the structural stress displacement field data of the aircraft model to obtain a feature point data set; training the feature point data set to obtain an accurate prediction result; according to the method, after artificial intelligence algorithm offline learning modeling is carried out on the aerodynamic force-heat-structure coupling data of analog simulation, the flight path is input, and the data, changing along the flight path, of the feature point data of the output aircraft is rapidly predicted.

The invention belongs to the field of dynamics of high-speed rotor aircrafts, and particularly relates to a method for calculating rotation flutter damping of a tilt rotorcraft. The method comprises the following steps: acquiring a full-aircraft modal shape and modal frequency of the tilt-rotor helicopter; screening out a mode which can be excited by a load transmitted by a rotor wing from a full-machine mode vibration mode of the tilt rotor wing helicopter; extracting modal displacement vectors and angular displacement vectors of the centers of the left and right pairs of rotor hubs in respective modals from the modals excited by the loads transmitted by the rotors, and converting the modal displacement vectors and the angular displacement vectors to a fuselage coordinate system; a rotor dynamics simulation model is established in CAMRAD, and the modal displacement vector and the angular displacement vector which are converted to the fuselage coordinate system are used for replacing the fuselage modal; a fuselage horizontal tail and a wing aerodynamic force output interface are respectively connected with the centers of a left propeller hub and a right propeller hub by using a CAMRAD CORE command.

The invention discloses a Gaussian process modeling method for a controlled separation steerage of an assembly aircraft, and belongs to the technical field of aerodynamics. In order to achieve the steering aerodynamic force obtaining precision of the combined aircraft, the method comprises the steps that a steering aerodynamic force database of a secondary aircraft in a disturbing flow field in the separation process of the combined aircraft is obtained through a wind tunnel grid force measurement test; the obtained first-stage aircraft attack angle, second-stage aircraft attack angle, two-stage aircraft gap and second-stage aircraft rudder deflection angle are combined into an input feature matrix, and the obtained pitching rudder deflection moment of the second-stage aircraft in the interference flow field of the first-stage aircraft serves as an output vector to obtain a training set, a verification set and a test set; and importing the training set into a Gaussian process regression model, training the Gaussian process regression model, verifying the trained Gaussian process regression model by using the verification set, and then testing by using the test set to obtain the Gaussian process regression model of the controlled separation steerage of the combined aircraft to predict the steerage aerodynamic torque.

The invention belongs to but not limited to the technical field of interception nets, and particularly relates to a multi-point traction type interception net launching device which comprises a base, a power device, a launcher and an interception net. The base can be adjusted and positioned at multiple angles, the borne impact force is at least not less than 15 KN, the base is designed into a folding structure, the packaging size is reduced, and transportation and storage are facilitated; the power device adopts compressed air as a power source, an air storage chamber is filled with high-pressure air, when a switch is turned on, energy of the high-pressure air is released instantly, and huge aerodynamic force acts on the pressed area of an object, so that the object moves in an accelerated mode from a static state and is thrown out. The power device body is composed of a launching tube, an air chamber, a valve element assembly, a firing device, a trigger, a carbon fiber air cylinder, a shell and the like, the launching tube, the carbon fiber air cylinder and the shell are distributed on the two sides of the air chamber, the firing device is arranged above the air chamber, and the trigger is pressed on the firing device; the launcher is in a horn shape, and four cylindrical storage bins connected with the launching tube are arranged on the edge of the launcher.

The invention discloses an aircraft body freedom degree flutter time domain analysis method, device and equipment and a medium, and relates to the field of aircraft dynamic aeroelastic analysis, and the method comprises the steps: determining the initial steady state of an aircraft through static balancing by using an aircraft aerodynamic grid model and an aircraft structural grid model; performing motion deformation coupling analysis on the aircraft after the initial disturbance is applied under the initial steady state condition, and determining the displacement of the aircraft surface node relative to the balance position based on the analysis result; updating a pneumatic grid through displacement, obtaining surface aerodynamic force of the aircraft based on the updated pneumatic grid, mapping the aerodynamic force to a structural grid by using a pneumatic structural coupling interface, and obtaining an aerodynamic load; and carrying out motion deformation coupling iteration and pneumatic structure coupling iteration based on the pneumatic load, carrying out coupling time propulsion when a coupling residual error meets a preset condition, judging the dynamic stability of the aircraft according to the displacement in the time propulsion process, and completing the time domain analysis of the degree-of-freedom flutter of the aircraft body. And high-precision analysis of the aircraft is realized.

An aircraft including at least one unducted turbine engine for the propulsion of the aircraft. The turbine engine comprising: a rotor and a stator comprising a plurality of stator blades extending radially with respect to the longitudinal axis, each stator blade being defined, in a plane transverse to the longitudinal axis, by an angular position; and at least one aerodynamic obstruction positioned close to the turbine engine. The stator of the turbine engine comprises stator blades having a first chord, referred to as conventional blades, and at least one stator blade having a second chord larger than the first chord, referred to as the elongate blade, said at least one elongate blade being positioned in an interference angular range defined opposite the aerodynamic obstacle, so as to increase the straightening of the airflow from the rotor in the interference angular range.

The invention relates to the technical field of helicopter aerodynamic force simulation, and discloses a sea rescue helicopter rotor aerodynamic force simulation method and system. The method comprises the following steps: acquiring a rotor wing downwash airflow historical speed data set, a historical sea surface wave data set and a historical aerodynamic pressure data set monitored within a preset historical time range; executing a dynamic volatility analysis process on the three types of historical data sets to generate a speed volatility index, a sea surface wave volatility index and an aerodynamic pressure volatility index; taking the three types of volatility indexes as guide parameters, performing adaptive search operation in a rotor aerodynamic force simulation parameter database, and determining an optimal simulation time interval; and performing numerical calculation simulation on pneumatic coupling between the rotor wing and the sea surface according to the optimal simulation time interval to generate a simulation data sequence, performing feature extraction and classification on the simulation data sequence by using an aerodynamic force recognition model, and outputting a rotor wing aerodynamic force simulation result.

The invention provides a hypersonic transition and forced motion coupling dynamic derivative prediction method, and belongs to the technical field of aerospace data processing, and the method comprises the steps: constructing an initial transition model based on a standard Reynolds shear stress transport turbulence model framework by using an intermittent factor transport equation and a transition initial momentum thickness Reynolds number transport equation; carrying out primary correction on a transportation equation of the initial transition model, and carrying out compressibility and transverse flow effect correction on the initial transition model according to a compressible boundary layer theory and transverse flow correlation to obtain a corrected transition model; carrying out coupling calculation on the corrected transition model and forced vibration by utilizing a rigid dynamic grid technology to obtain unsteady aerodynamic force and torque; and identifying the unsteady aerodynamic force and torque by using an integral method to obtain a predicted combined dynamic derivative. The method has a full-dimension dynamic derivative analysis capability, breaks through speed domain limitation, and realizes full coverage from subsonic velocity or transonic velocity or supersonic velocity to hypersonic velocity working conditions.

The ball is a thin-walled hollow sphere that includes an array of apertures with dimples in a pattern that are concentric to the aperture array. Dimples are provided on the exterior surface of the ball between the apertures with the pattern and dimensions of the dimples varying based on proximity to the apertures. Internal protuberances are provided on the interior surface of the hollow ball such that the interior surface also manages aerodynamic airflow through the ball. Channels are also provided around each of the apertures within the ball which may intersect proximate dimples or be spaced therefrom. The ball also includes apertures having converging and diverging sidewalls that form a venturi shape to increase inflow air into the apertures and outflow air out of the apertures. Finally, the ball also maintains a weight:diameter ratio within a fixed range.

The invention discloses a small sample aerodynamic modeling method based on multi-task learning. Comprising the following steps: constructing a multi-task prediction model, integrating an auxiliary task network and a target task network, and enhancing the feature extraction capability of an encoder through an SE layer and an attention layer; acquiring multi-source aerodynamic force data and carrying out standardized preprocessing on the multi-source aerodynamic force data; a dynamic weight mechanism is designed, the influence of auxiliary task prediction on target task output is adaptively adjusted according to input features, and effective fusion of low-fidelity data and high-fidelity data is achieved; two-stage training is carried out to ensure the efficiency and stability of multi-task learning; the prediction capability of the model under the limited sample condition is verified; through the synergistic effect of multi-task knowledge migration, feature enhancement and dynamic task integration, in combination with the feature optimization characteristics of an SE layer and an attention mechanism, high-precision aerodynamic prediction is realized under the condition of limited samples, and an efficient and accurate aerodynamic prediction method is provided for aircraft design.

The invention relates to a CFD (Computational Fluid Dynamics) method for nearly continuous hypersonic high-temperature thermochemical unbalanced flow numerical simulation, which comprises the following steps of: 1, introducing a Park double-temperature model into an NS equation to describe a thermochemical unbalanced effect at high temperature, and closing a shear stress tensor, a heat flow vector and a mass diffusion flux by adopting an equivalent constitutive model to obtain a closed shear stress model; capturing speed slippage and temperature jump phenomena on the wall surface by adopting corresponding slippage boundary conditions to obtain a macroscopic control equation and boundary conditions suitable for near-continuous hypersonic high-temperature thermochemical non-equilibrium flow; 2, discretely solving the macroscopic control equation and the boundary conditions given in the step 1 by adopting a finite volume method; and step 3, performing post-processing on a solution result obtained in the step 2 to obtain a streaming flow field, a wall surface aerodynamic force and a wall surface heat flow of the calculated appearance.

A land vehicle having an aerodynamic system includes a rear air diffuser positioned at the lower rear of the land vehicle and configured to redirect airflow from under the land vehicle upward behind the land vehicle while the land vehicle is traveling in a forward direction, and a rear wing positioned at the upper rear of the land vehicle and configured to generate downforce when the land vehicle is traveling in a forward direction. The rear air diffuser and the rear wing may be configured and arranged to merge the flow of air from underneath the land vehicle with the flow of air from above the land vehicle.

Some implementations relate to methods, systems and computer readable media to provide an aerodynamic force model for real-time distributed physics simulation. According to one aspect, a computer-implemented method includes receiving a description of a mechanism that includes physically coupled geometric assemblies within a virtual experience, with each geometric assembly defining a surface mesh. The description includes motion data of the mechanism. The method further includes, for each geometric assembly, identifying exposed surface areas of the surface mesh. The method further includes evaluating an aerodynamic force model based on the exposed surface areas and the motion data of the mechanism, where the aerodynamic force model includes a pressure coefficient that varies based on an angle of attack and based on windward and leeward facing surfaces. The method further includes integrating the aerodynamic force models into a physics simulation to refine the motion data of the mechanism.

The invention provides a carrier rocket control method based on wind field characteristics, and the method comprises the steps: obtaining a wind field model generated through the fitting of actually measured wind field data; according to the speed of the carrier rocket at the current moment under the launching coordinate system, the wind speed at the current moment and the first conversion relation, the relative speed of the carrier rocket relative to the atmosphere at the current moment in the aircraft coordinate system is determined; influence parameters are calculated through the relative speed, and the aerodynamic force at the current moment is calculated according to the influence parameters; integrating the mass center motion differential equation set of the carrier rocket by combining the aerodynamic force, the preset initial value of the preset control parameter and the flight program angle expression of the carrier rocket changing along with time, and circularly calculating the preset state variable of each moment; calculating a target function of the carrier rocket according to the preset state variable at each moment; and on the basis of a sequential quadratic programming algorithm, target functions corresponding to different preset control parameters are calculated through the target functions and preset constraint conditions, so that optimal values of the preset control parameters are determined.

The invention discloses a patrol aircraft flight simulation verification method and system based on a digital model, and belongs to the technical field of aircraft flight simulation. The system comprises a launch control device digital model, a scene simulation digital model, a seeker model, an autopilot model, an aerodynamic model, a steering engine model, a simulation solution model and an inertial navigation model, and digitizes the environment, a target, a seeker, a steering engine, inertial navigation, a power device, an autopilot and other modules in the flight process of the cruising aircraft. The operation hardware platform of each module is virtualized, on the premise that the original system communication mode is not changed, guidance and control of the flight process of the cruising aircraft and evaluation of performance such as miss distance can be completed, and the influence of various environment changes and structure changes on the flight of the cruising aircraft can be quickly verified through flight simulation of the cruising aircraft.

The invention belongs to the technical field of aircraft simulation, and particularly relates to an aircraft maneuvering load simulation method and device considering a downwash effect. The method comprises the following steps: S1, carrying out 1g level flight balancing on the aircraft, and determining a balancing attack angle of the aircraft and a balancing elevator skewness of the aircraft; s2, superposing the aerodynamic coefficient increment brought by the downwash angle of attack to the aerodynamic coefficient of the aircraft body, and determining the aerodynamic force of the aircraft; s3, dynamic response solving is carried out on aerodynamic force of the aircraft, and the aerodynamic attack angle, the flight speed and the pitch angle speed are obtained; s4, calculating elevator skewness at the current moment according to elevator instruction input; and S5, re-determining the aerodynamic force of the aircraft according to the flight speed, the pitch angle speed, the aerodynamic angle of attack and the elevator skewness until the simulation is finished. According to the method, the effect of the downwash influence in the aircraft maneuvering load determination process is reasonably considered, the simulation process is simple, the precision is high, and the engineering practice requirement is met.

The invention discloses a flight-oriented fan wing trailing edge adjusting device, and relates to the technical field of aircrafts, the flight-oriented fan wing trailing edge adjusting device comprises a trailing edge angle adjusting mechanism, a rectification angle adjusting mechanism, a circular end plate and a rectification plate, the trailing edge angle adjusting mechanism comprises a steering engine, a first connecting rod mechanism, a trailing edge plate and a rotating shaft; the steering engine is fixedly installed on the circular end plate and controls the rear edge plate to move through the first connecting rod mechanism, the rotating shaft is arranged on one side of the rear edge plate, and the rear edge plate is hinged to the rectifying plate through the rotating shaft. By freely adjusting the trailing edge angle and the rectification angle, the aerodynamic force of the fan-wing aircraft is changed, the lift force and thrust distribution is improved, the aerodynamic effect is enhanced, the energy consumption is reduced, and the flight efficiency is improved. The fan wing trailing edge adjusting device is applied to a serial fan wing, forward coupling gain between two wings is achieved by changing the size of a trailing edge angle and a rectification angle, different flight conditions are met, various flight requirements are flexibly met, and the load performance and the endurance are optimized.

An actuator assembly for moving a movable aerodynamic surface of an aircraft is disclosed including an input section for introducing an externally provided rotary motion into the actuator assembly, and a geared rotary actuator having at least one actuator slice having a planetary gear with a sun gear couplable to the input section, a fixed gear attachable to a fixed component of the aircraft and an output gear couplable with the movable aerodynamic surface. At least one clutch is connected to the geared rotary actuator and is configured to couple the movable surface with the input section and to selectively decouple the movable surface from the input section, such that it is freely movable.

The invention provides a rotating speed determination method in aero-engine complete machine three-dimensional simulation, and belongs to the technical field of aero-engine simulation, and the method specifically comprises the steps: constructing a fluid-structure interaction control equation on each grid unit of an aero-engine complete machine geometric model; forming an integral coupling matrix, and solving the integral coupling matrix; simulating the calculated thermal state geometric model; the gas compressor and the turbine can only rotate around the rotating shaft; calculating aerodynamic forces acting on the compressor blades and the turbine blades; in each time step, the rotating speed is updated according to the angular acceleration; and after a plurality of iterations, when the difference value of the aerodynamic torque borne by the gas compressor and the aerodynamic torque borne by the turbine accords with a preset range, the corresponding rotating speed is the steady-state working rotating speed of the engine under the thermal state geometry. According to the processing scheme, the engine can automatically match and dynamically determine the working rotating speed in the simulation process, the inaccuracy of the preset rotating speed is avoided, and the precision and reliability of a simulation result are improved.

The invention discloses an aircraft unsteady aerodynamic force prediction method, system and device and a medium, and relates to the technical field of aircraft modeling and time sequence prediction. According to the method, an LSTM neural network prediction model is constructed, an aircraft unsteady aerodynamic force time sequence data set processed by a sliding window is used as a training data set, the LSTM neural network prediction model is trained, and the aircraft unsteady aerodynamic force time sequence data set is predicted by using the trained LSTM neural network prediction model. And the change of the unsteady aerodynamic performance of the aircraft under new incoming flow disturbance is predicted online in real time. According to the method, the prediction response to the attitude angle and the unsteady aerodynamic performance of the aircraft under the incoming flow disturbance is fast, the accuracy is high, and meanwhile reference can be provided for reducing the engineering problems of aircraft vibration, flight attitude control and the like.

The invention discloses a large flexible wing static aeroelasticity analysis method based on a physical information neural network, and belongs to the field of aircraft design. The method comprises the following steps: for the wing with the super-high aspect ratio, firstly, dispersing the wing into n spanwise sections based on a strip theory, and initializing a reference angle-of-attack parameter of each spanwise section; then, a physical information neural network (PINN) is built, unsteady flow field solving is carried out on each section, for the current ith iteration, the physical information neural network is used for sequentially traversing and extracting pressure distribution of the airfoil surface of each section, respective aerodynamic load distribution is calculated, and the generalized force of the airfoil is obtained; and solving a structural dynamics equation based on the generalized force and structural information of the wing, calculating the deformation of the large-deformation wing by using a modal rotation method, and finally, carrying out solving iterative circulation on aerodynamic force and structural deformation until the attack angle is converged, so as to obtain the actual lift force L of the wing. According to the method, the calculation precision and efficiency are remarkably improved, and different flight conditions and wing design can be adapted.

The invention provides a self-adaptive cross damping flutter suppression method for blades of an offshore wind turbine, and belongs to the technical field of wind power generation based on computer data processing. The method comprises the following steps: firstly, establishing an aeroelastic model containing a non-diagonal damping item, and coupling sinking and floating motion and pitching motion through a cross damping coefficient; secondly, designing a self-adaptive control law based on a differential Riccati equation, and dynamically adjusting a damping parameter to compensate the time-varying aerodynamic force; and finally, proving closed-loop stability through a Lyapunov theory, and verifying a control effect by utilizing numerical simulation. An external pneumatic executing mechanism is not needed, the structure is simple, response is fast, the vibration convergence time can be shortened to be within 4 seconds at the wind speed of 8 m / s, the steady-state error approaches zero, and the method is suitable for flutter suppression of the blades of the large offshore wind turbine.