145 results about "Aeroelasticity" patented technology

The invention discloses a CFD (computational fluid dynamics) / CSD (circuit switch data) coupled solving nonlinear aeroelasticity simulation method, comprising: firstly, carrying out calculation initialization; then solving the CFD of unsteady aerodynamic force; converting boundary load information; solving the CSD of nonlinear structure dynamic response; judging whether to exist computation or notby taking the response of structural finite-element mesh point displacement as a convergence criteria; if not exiting, carrying out the dynamic mesh deformation of a flow field after the surface displacement information of a computation model is converted; and then, continuing to compute until the convergence criteria is satisfied, thus finishing computation. By means of the simulation method disclosed by the invention, the problem of pneumatic nonlinearity and structural nonlinearity coupling and the problem of geometric large deformation dynamic mesh can be solved; and the simulation methoddisclosed by the invention can be applied on the nonlinear aeroelasticity analysis of aerospace aircrafts and the nonlinear aeroelasticity analysis of civil high-rise buildings and bridges.

An aircraft aeroelasticity measurement system according to the invention can be used to perform in-flight wing twist measurements of an aircraft or any moving vehicle or structure on which its sensors can be mounted. The system includes a centralized processing unit, a reference navigation unit mounted in the aircraft body, a plurality of measurement navigation units mounted in the aircraft wings, and a global positioning system (“GPS”) receiver that generates GPS data for enhancing the accuracy of the navigation data. The system collects navigation data from the reference navigation unit and the measurement navigation units, generates navigation solutions from the navigation data, and performs a stochastic alignment and flexure estimation procedure to generate a corrected measurement solution that indicates the aeroelasticity of the aircraft. Wing twist and wing deflection information can be derived from the corrected measurement solution.

The invention provides a geometric nonlinear static aeroelasticity analysis method based on a structure reduced-order model. The method includes: building an aircraft wing or full-aircraft finite element model, presetting follow-up test load, and using finite element software to solve corresponding test deformation; under a given nonlinear rigidity coefficient form, performing regression analysison the follow-up test load and the test deformation to solve a nonlinear rigidity coefficient and built the structure reduced-order model; applying the structure reduced-order model to aircraft geometric nonlinear static aeroelasticity analysis, and building aircraft wing geometric nonlinear static aeroelasticity deformation calculation flow and full-aircraft geometric nonlinear static aeroelasticity trim analysis flow. The method has the advantages that the cubic polynomial nonlinear rigidity coefficient is given, an aerodynamic force follow-up effect and extension-direction deformation underaircraft wing geometric nonlinear large deformation are considered, the reasonability and accuracy of geometric nonlinear static aeroelasticity analysis are increased, and efficient analysis and calculation of geometric nonlinear static aeroelasticity deformation and trim are benefited.

The invention discloses a multi-objective optimization design method considering the aeroelasticity restraint and for a wing made of a composite material. The method includes the following steps that (1) a finite element model is built and analyzed; (2) weight is optimized; (3) gust response is optimized; (4) weight is optimized for the second time. The strength, rigidity and aeroelasticity restraint is considered, multi-objective optimization of structure weight and gust response is achieved, and the problem that the gust response amplitude is increased when structure weight optimization is conducted by using a traditional optimization method for the wing made of the composite material is solved. In the optimization process, in the gust response optimization stage, only the composite material auxiliary layer direction is used as a design variable, and gust response design requirement can be met under the condition that the wing weight is not changed.

A system is provided for control of a structure immersed in a fluid flow regime. The system comprises a plurality of flow sensors disposed on a surface of the structure, the flow sensors including at least one flow bifurcation point sensor. The system further comprises at least one structural sensor attached to the structure for measurement of a structural parameter. A data processor in communication with the flow sensors and the at least one structural sensor is configured for processing the data signals and for generating and transmitting one or more actuator command signals to effect a desired flow and structural state for the structure. The system also comprises at least one actuator in communication with the data processor. Each of the at least one actuator is configured for effecting, in response to an actuator command signal, one of the set consisting of a structural deformation in at least a portion of the structure, a change in configuration of a control surface of the structure, and an operational change in a flow control device embedded in the structure.

The invention relates to a surface element correction and grid beforehand self-adaption calculation method. According to the method, segmented linearizing correction is carried out on a low-order surface element method with multiple groups of CFD aerodynamic force load data under different attack angles, and distribution of surface element calculation grids is optimized by adopting a self-adaption simulated annealing algorithm. By means of self-adaption optimization of the surface element grids, the accuracy of the surface element correction method is improved, and the defect that a traditional surface element correction method depends on grid distribution is overcome. When the optimized grids are applied to the surface element correction method, the advantage of high calculation efficiency of the surface element method is kept, the accuracy of overall wing stress close to the CFD data can be ensured, and the accuracy and efficiency in the aeroelasticity optimal iteration design process are effectively improved. The accuracy error between the obtained aerodynamic load data and CFD calculation results is within 2% and effectively extends to the non-linear segment in which aerodynamic loads change along with the attack angle, and the aerodynamic load calculation efficiency in the changing process of structural rigidity parameters is improved.

The invention discloses a pneumatic / structure / static aeroelasticity coupling optimizing method for a high-aspect-ratio wing. The pneumatic / structure / static aeroelasticity coupling optimizing method is based on the system decomposition strategy in multidisciplinary optimization for system-level decomposition and subsystem-level decomposition. In pneumatic subsystem-level decomposition, the pneumatic performance under the cruising appearance serves as an optimal object, and the rigidity-appearance pneumatic performance in traditional pneumatic optimization is replaced; after pneumatic subsystem-level decomposition is completed, and pneumatic loads, pneumatic appearance design parameters and structural configuration parameters serve as the input conditions for structure subsystem-level optimization; after pneumatic subsystem-level decomposition and structure subsystem-level decomposition are completed, the target quantity of system-level decomposition is output, system-level decomposition is returned, and therefore the pneumatic / structure / static aeroelasticity multidisciplinary comprehensive optimization design process is achieved through iterative optimization of the pneumatic appearance parameters and the structure parameters.

The invention relates to a biomimetic micro air vehicle with 8-shaped flapping tracks, which makes use of a flexible carbon fiber rod and a parylene film as an aerofoil framework and an aerofoil covering skin respectively for a palm-sized micro air vehicle. Even through a flapping (gear connecting rod) driving mechanism is only designed into a single degree of freedom four-bar (mechanism), the linkage of the upward and downward flapping is successfully coupled to flow direction forward and backward vibrations through the material and the shape design of a flexible aerofoil, thus three-dimensional 8-shaped flapping tracks are assembled at the aerofoil tip; and the flapping frequency of the generation of the 8-shaped tracks is close to or even more than 20 Hz, thus a micro air vehicle with a simple structure is formed and can generate the 8-shaped tracks automatically according to the essence of aeroelasticity without getting help from the design of a complicated delicate movement mechanism with multiple degrees of freedom.

An active wing and lift surface control system for an aircraft is described. The wing and lift surface control system includes an aeroelasticity measurement system configured to provide at least one of real time wing twist measurements and real time measurements of aircraft body bending, at least one actuator mechanically coupled to a control surface of the aircraft, and a control system communicatively coupled to the aeroelasticity measurement system and to the at least one actuator. The control system is operable to receive the measurements from the aeroelasticity measurement system and generate control signals, based on the real time measurements, to operate the at least one actuator to adjust a drag associated with one or more of the wing and the aircraft body.

The invention belongs to the field of airplane aeroelasticity tests, and relates to an airplane ground test scope, in particular to an automatic amplitude modulation frequency scanning method for an airplane servo elasticity frequency response test. Control plane overloads are fed back to automatically adjust a frequency scanning amplitude value, on the premise that safety of a control pane structure is guaranteed, control plane deflection exciting force of all frequency scanning points is kept maximum so that sufficiently large airplane structure response can be excited, the signal to noise ratio of a response signal of an airplane-flight control combined circuit is improved, test errors are reduced, and reliability of a test result is improved. Meanwhile, the adjusting process of an excitation signal amplitude value is controlled through construction of a hyperbolic tangent-exponent function, violent response of a control plane caused by sudden changes of the excitation signal amplitude value is prevented from happening in the process of frequency scanning, and safety of the control plane structure is ensured. In addition, the advantages of a real-time simulation system of a computer are brought into full play, in the process of the test, people only need to slightly click a mouse, so that parameters such as control limit overloads and excitation signal cycles can be adjusted, and the automatic amplitude modulation frequency scanning method is high in practicability.

The invention relates to the field of aeroelasticity, in particular to a wing aerodynamics modeling method. The airplane stiffness and mass distribution can be obtained in the early airplane design period under the condition of undefined wing structure arrangement. According to the wing aerodynamics modeling method provided by the invention, firstly, a finite element model of a wing pure metal plate element is built; the obtained aerodynamic lift is calculated; then, the thickness of the metal plate element is used as a variable, the wing tip deformation, the gravity center, the modal frequency, the flutter speed and the like are used as constraint conditions, and the wing weight is used as an optimization objective; finally, a group of metal plate element thickness distribution values is obtained through optimization calculation, so that the airplane stiffness and mass distribution is simulated; and the wing stiffness and mass distribution can be fast obtained in the early airplane design period under the conditions that the wing structure arrangement is not clear and only the wing pneumatic appearance and the weight are known, so that the wing aerodynamics feature analysis is carried out.

The invention provides a scaling modeling method based on aeroelasticity constraints. The method includes the steps that (A) an aerodynamic force analytical model and a structure finite element model are established according to a standard wing model; (B) a structure flutter analytical model is established according to the aerodynamic force analytical model and the structure finite element model; (C) the flutter characteristic of the model is solved according to a boundary condition; (D) whether the flutter characteristic meets a predetermined requirement or not is judged according to a predetermined constraint condition. The scaling modeling method has the advantages that in the wind-tunnel scaling modeling method of a composite material wing, the dynamics constraint, the pitch line constraint and the flutter constraint are comprehensively considered, the flutter model is optimally obtained, and thus the dynamics and flutter characteristic of a true structure can be better simulated; by optimal modeling of a low speed and high subsonic speed flutter model, flutter scaling modeling procedures based on an optimization method are respectively established, the modeling efficiency can be improved, the model can be made to well reflect the aeroelasticity characteristic of an entity, and the method is applicable to aeroelasticity wind-tunnel modeling of civil aircraft composite material wings.

The invention provides a simulation method for an elastic waverider hypersonic flight vehicle. The method includes the following steps that considering influences of aerodynamic heating and variable section inertia moment, an elastic model of a hypersonic flight vehicle free beam structure is established; a model superposition method is used for solving an elastic vibration equation, the intrinsic frequency, damping ratio and intrinsic vibration mode of each mode are solved, and on this basis, an elastic vibration generalized coordinate equation of the waverider flight vehicle is obtained; a computational fluid dynamics method is used for obtaining aerodynamic force and engine propulsion of the flight vehicle; on the basis of analyzing thrust, aerodynamic force and aeroelasticity, a rigid body-elastic coupling model of the hypersonic flight vehicle is established. The simulation method is suitable for performing modeling and simulation on the waverider flight vehicle. On this basis, the established model of the waverider hypersonic flight vehicle is more accurate, and the aerodynamic / propelling / elastic coupling characteristics of the flight vehicle can be reflected better in the process of model simulation.

The invention relates to a scheme stage wing parameter estimation method and belongs to the field of wing aeroelasticity design. The method includes the steps that the mass of a wing is determined; root bending rigidity and root torsional rigidity of the wing are calculated; the axial position of an elastic shaft of the wing is determined; bending rigidity distribution of the elastic shaft of the wing and torsional rigidity distribution of the elastic shaft of the wing are determined; a wing dynamical model is built; an optimization module is built according to the wing dynamic model to obtain optimal elastic shaft rigidity distribution, and the optimization module takes minimum wing beam equivalent volume of the wing as an objective function and wing root rigidity distribution in the third step and curve shape parameters of elastic shaft rigidity distribution in the fifth step as variables for optimization. Through the method, rigidity and mass distribution of the wing are obtained quickly, and the dynamic model of the wing is built.

The invention discloses a vehicle-mounted aeroelastic test platform system, and provides a comprehensive test platform system for model support, air flow generation and adjustment and movement test in an aeroelastic test. The vehicle-mounted aeroelastic test platform system adopts a test model in aeroelasticity (static aeroelastic divergence, vibration, aeroservoelasticity and the like) researching or other aerodynamics researching as an application object, and can be used for the blowing tests such as the principle exploring, performance test, theory verification and the like of aeroelasticity or aerodynamics. The whole system comprises a vehicle subsystem, a data test and control subsystem and a model support subsystem.

The invention relates to an aircraft aeroelasticity inertia sensor layout method. According to the method, the aircraft structure dynamic response is solved by employing the finite element method, a finite element model is corrected via a fluid-solid interaction method, and a fluid-solid coupling analysis of all the flight states and disturbance factors is conducted; the relatively precise structure dynamic response is calculated, and the response of each node in the finite element model is obtained; standardization processing of a data set is performed, and an elasticity wing structure dynamic response data set is generated; and data clustering by employing a clustering method based on distance measurement in pattern recognition is finally performed, and the final positions of the sensors are optimized and obtained. According to the method, the maximization of measuring information of sensors is regarded as the overall goal, optimization balance between the quantity of the sensors and the optimal layout positions is performed, and valuable aeroelasticity characteristic values can be acquired at the arranging position of each sensor.

The invention discloses a synchronous wind tunnel testing device for aeroelasticity and wind pressure considering the aerodynamic interference effect, and is applied to the field of wind tunnel testing devices. The device comprises a testing bench fixed in a wind tunnel, and a testing unit arranged on the testing bench. According to the synchronous wind tunnel testing device for aeroelasticity andwind pressure considering the aerodynamic interference effect, a wind pressure test and a aeroelasticity test can be synchronously carried out, and the influence of model vibration on wind load is comprehensively reflected; the air inertial force in the model vibration is measured, so that the obtained result of the aerodynamic force actually caused by wind is more accurate; the simulation of a multi-building structure layout is realized in one testing device, and the aerodynamic interference effect factor existing in a building group is taken into consideration; and meanwhile, it is convenient to adjust the relative position of each building structure, namely the distance between every two buildings, and the included angle relative to the wind direction, so that the research on the influence of the position layout of different buildings on aeroelasticity characteristics and wind pressure characteristics in one testing device is achieved, and the more practical theoretical research support is provided for the technical field of aeroelasticity characteristics and wind pressure characteristics.

The invention belongs to the field of aircraft strength tests, and relates to the category of aircraft aeroelasticity tests, in particular to a method for preventing control plane integral saturation in an aircraft ground servo elasticity test. A real-time emulation system is connected between a flight control sensor and a flight control computer in series. The real-time emulation system comprises a main control computer and an emulation computer. After parameters of a high-pass filter are set and coded in the main control computer, the coded parameters are transmitted to the emulation computer. The parameters of the high-pass filter are designed with the dynamic characters of the control law within the range of the aircraft elastic vibration frequency as the restraint, the time of control plane integral saturation is prolonged, fast control plane integral saturation is effectively restrained, and meanwhile the effect on the test result is reduced to the maximum extent. The filter is placed in an external emulation machine, hardware and software of an onboard control system do not need to be adjusted, the parameters of the filter can be adjusted according to the actual situation in the test, and the method is high in adaptability and practicability.

The invention provides a method for aeroelastic tailoring of a composite wing. The method for aeroelastic tailoring of the composite wing comprises the steps that aeroelastic tailoring of the composite wing is achieved by means of aeroelasticity optimization, and solution of aeroelasticity optimization is conducted according to a genetic / sensitivity-based hybrid optimization method. The genetic / sensitivity-based hybrid optimization method comprises the steps that (A) a group of individuals are generated randomly according to a genetic algorithm; (B) performance evaluation is conducted on the individuals in the group one by one; (C) according to a user definition, whether sensitivity-based optimization is applied or not is judged; (D) whether the end condition is met or not is judged; (E) if the end condition is not met, a new generation of excellent individuals are generated; if the end condition is met, operation is ended. By the adoption of the method for aeroelastic tailoring of the composite wing and the genetic / sensitivity-based hybrid optimization method, under the condition that multiple requirements for flutter, diffusion, inherent frequency, deformation, aileron efficiency, strength, the flight load, weight and the like are met, tailoring design of the structure is achieved and optimized.

The invention discloses a method for optimizing flight dynamics of a small unmanned aerial vehicle based on a bionic flexible wing, and belongs to the field of overall design and flight dynamics of small fixed-wing unmanned aerial vehicles. According to the method, the bionic flexible wing is designed by imitating geometric and structural features of wings of large gliding birds; based on the bionic flexible wing, the modal analysis is performed on the wing to obtain first few modes of torsion which are very significant; an obtained first few mode equation set and a flight dynamics equation set are synthesized, and aerodynamic coupled terms generated by the elastic deformation are added; the influence of the aeroelastic effect of the bionic wing on the coupled of the flight dynamics is obtained, and critical parameters are refined, signs and sizes of critical parameters are adjusted by adjusting the wing structure design without changing the fuselage structure arrangement and overall pneumatic layout, the coupling form of the aeroelasticity of the wing and the flight dynamics is adjusted to achieve the goals of reducing the short-cycle frequency, increasing the short-cycle dampingratio and slow downing the gust disturbance of the small unmanned aerial vehicle.

The invention discloses a testing device for a static aeroelasticity test of a wind tunnel, and the device comprises a pedestal which is provided with an accommodation cavity with an upper opening and is also provided with a longitudinal symmetric surface; a cushion block which is disposed in the accommodation cavity, wherein a first side wall of the cushion block is closely attached to and detachably connected with one side wall, parallel to the longitudinal symmetric surface, of the accommodation cavity; a balance which is disposed in the accommodation cavity, and is detachably connected to a second side wall, forming an inclined angle with the first side wall, of the cushion block; a missile wining model which comprises a main body part and a root part located below the main body part, wherein the main body part vertically extends above the pedestal, and the root part stretches into the accommodation cavity, and is connected to the balance. Through the selection of the cushion block from a plurality of cushion blocks with different inclined angles, the installation angle of the missile wining model relative to the longitudinal symmetric surface. The method can precisely measure the aerodynamic force and torque characteristics of the missile wining model at different attack angles under the action of an aerodynamic load after elastic changes.

The invention relates to a flexible wing flutter model variable in trailing edge and belongs to the technical field of aeroelasticity tests. The flexible wing flutter model comprises a wing beam, a leading edge dimensional-shape frame section, a trailing edge frame section, a trailing edge control connection rod and a steering engine. The leading edge dimensional-shape frame section is fixed to the wing beam and hinged to the trailing edge frame section, the steering engine is arranged on the wing beam, and accordingly the trailing edge frame section can be driven to deflect relative to the leading edge dimensional-shape frame section through the steering engine; and one end of the trailing edge control connection rod is arranged on the leading edge dimensional-shape frame section, and the other end of the trailing edge control connection rod is arranged on the trailing edge frame section so that the deflection directions of multiple frame section bodies of the trailing edge frame section can be controlled. According to the flexible wing flutter model variable in trailing edge, a hinge rod mode is adopted, the wing trailing edge can be deformed through movement of the connection rod, wing unsteady aerodynamic distribution is changed so that a wing can dissipate energy outwards, and accordingly the flutter inhibition purpose is achieved.

The invention discloses a real-time measurement method of the deformation of a test piece free end in a wind tunnel test, relating to the technical field of aircraft aeroelasticity. The method concretely comprises a step of selecting a reference point at a free end of a test piece, installing a first camera and a second camera at a side wall and a top in a wind tunnel by reference to the reference point, and setting a first mark tag and a second mark tag at side walls opposite to the first camera and the second camera, a step of applying external force to the test piece such that the test piece is in vertical deformation and horizontal deformation, recording positions of the reference point on the first mark tag and the second mark tag viewed through the first camera and the second camera, and a step of carrying out a blowing test on the test piece, measuring relative positions of the reference point in the first mark tag and the second mark tag through screen captures, and calculating the actual deformation of the test piece. The method has the advantages of simple operation and convenience in use, and the measurement precision can meet the requirements of the engineering.

The invention discloses a magnetic gas double-suspension breaking-type cone bearing with porous media. The cone bearing with an air static pressure supporting system and an electromagnetic suspensionsupporting system comprises first and second bearing end covers, a rotor shaft, first and second stators, first and second magnetic conductive sleeves, first and second enameled coils, an adjustment ring, first and second dust-proof seal rings, and the first and second porous media; according to the air static pressure supporting system, air enters from air inlet holes of the two stators, passes through air inlet cavities to act on the two magnetic conductive sleeves through the porous media, and forms tapered aeroelastic potential energy between the two stators and the two magnetic conductivesleeves, and accordingly, axial and radial support of the bearing is realized; the electromagnetic suspension supporting system is realized by the physical principle of electromagnetic attraction onthe two stators. The magnetic gas double-suspension breaking-type cone bearing has the advantages that the abrasion of the bearing can be remarkably reduced, the rotation speed can be increased, separate arrangement of auxiliary supports and circulating cooling systems is not needed, and the adjustment capability and the precision of a bearing system are improved.

The invention relates to a longitudinal freedom degree simulation full aircraft flutter wind tunnel model support system and belongs to the technical field of aeroelasticity tests. The longitudinal freedom degree simulation full aircraft flutter wind tunnel model support system comprises a support beam rack and a model suspension device, wherein the model suspension device is arranged on the support beam rack and is used for suspending a flutter model, a model suspension device can be arranged on a main beam rack in a sliding manner, and displacement of the model suspension device along an axis of the main beam rack can be restricted via an ups and downs brake block arranged on a lateral part of the model suspension device; the model suspension device also comprises a pitching brake block and a model suspension shaft, the model suspension shaft is used for suspending the flutter model so as to make the flutter model rotate around the model suspension shaft, the pitching brake block can press tightly on or get away from the flutter model, and therefore pitching angle displacement of the model suspension device can be restricted. According to the longitudinal freedom degree simulation full aircraft flutter wind tunnel model support system, freedom of a full aircraft flutter in an up and down direction and a pitching direction in a wind tunnel can be realized via the ups and downs brake block and the pitching brake block, and therefore a real flight state of an aircraft can be simulated.

The invention provides a fin and wing aeroelasticity test device of a high-speed wind tunnel. The fin and wing aeroelasticity test device of the high-speed wind tunnel comprises a protection mechanism, a supporting mechanism, a reflection board, a supporting baffle, a base and a bottom surface platform. The protection mechanism is arranged in front of the supporting mechanism and mainly comprises a protection cover, a guide rail and a cylinder. The cylinder is fixed to the lower portion of the base. The base is installed on the bottom surface platform, and the position of the base can be adjusted front and back. The protection cover is connected with the cylinder and the guide rail in a mode of moving vertically. The supporting mechanism is fixed to the base and used for supporting a model. The reflection board is arranged at the root of the model through the supporting baffle. The reflection board is provided with an opening matched with the protection cover and the model so that the reflection board cannot make contact with the protection cover or the model. The fin and wing aeroelasticity test device of the high-speed wind tunnel can be used for carrying out fin and wing aeroelasticity test research in the high-speed wind tunnel. The protection mechanism is used for protecting the model against impact from a wind tunnel flow field, and when the vibration amplitude of the model is large, vibration damage to the model can be prevented.

The invention provides a bidirectional intelligent selection fluid-solid coupling analysis method. The technical problems that an existing interpolation method is poor in accuracy, low in data transmission precision, low in interpolation efficiency, low in automation degree and the like can be solved. The method comprises the following steps: S1, establishing a flow field CFD grid and a structurefinite element grid; S2, intelligently matching and grouping all structure finite element grid nodes according to the flow field object planes numbered in sequence through a set intelligent matching judgment criterion; S3, training the structural finite element grid nodes after matching and grouping to obtain a structural feature point set {QFEM}, and establishing a data transfer relationship {TFC} between the {QFEM} and the corresponding flow field object plane nodes; S4, obtaining the aerodynamic force of the flow field; S5, interpolating aerodynamic force, calculating elastic deformation, and interpolating the deformation to a CFD grid; S6, carrying out the flow field calculation of the CFD grid for the deformed flow field grid; and S7, judging whether a calculation result meets a static aeroelasticity convergence condition or not.