35 results about "Flutter" patented technology

A method and control for predicting the remaining useful life of an airfoil for a gas turbine engine includes the steps of monitoring conditions of the blade such as flutter, leaning, etc. A measured amount of deflection of the airfoil is compared to tabulated data to predict an expected crack length which is likely causing the deflection, etc. Once a predicted crack length has been identified, the amount of accumulated damage to the airfoil at the crack is monitored and stored. The amount of useful life for the blade can be predicted by compiling the accumulated damage over time. The useful life remaining can be displayed such that flight plans or maintenance schedules for the aircraft can be modified as appropriate.

The invention discloses a flight test determination method of the multi-input and multi-output equivalent pneumatic servo elastic robust stability, used for solving the technical problems of the conservation and complex calculation of the traditional robust analysis method. Based on the technical scheme, the flight test determination method comprises the following steps of obtaining an open-loop transfer function frequency characteristic matrix of a multi-input and multi-output system through a frequency sweeping flight test and the like; lagging a phase angle through serial gaining at each loop; obtaining a scalar judgment formula of phase and amplitude margin between two adjacent flutter frequency regions according to direct equivalence of the closed-loop frequency characteristics by carrying out characteristic decomposition on the system open-loop transfer function frequency characteristic matrix; and calculating the phase margin and the ASE stability at the flutter frequency in a mode similar to a single-input and single-output system. The invention simplifies the problems through characteristic linear conversion, obtains the scalar judgment formula, gives a calculation method of the flutter boundary stability, the margin and the safety during elastic flight of a canard wing aircraft and reduces the conservation of the traditional method.

The invention relates to a flutter test safety protection method, belongs to the technical field of aviation aerodynamic wind tunnel tests, and aims to solve the problem of damage to a model or internal parts of a wind tunnel caused by excessive vibration of a test model in the wind tunnel flutter test process. According to the method, the flow field stability condition is monitored through a dataacquisition and analysis module when a wind tunnel control module starts a test, the value of an acceleration sensor on the test model is monitored, calculation and analysis are carried out, a safetyprotection device is used to carry out safety protection on the model, the control difficulty of the aircraft in the pneumatic simulation flight test is reduced, the control efficiency is improved, the problems of large size of a development board and occupation of model space are solved, the weight of the model is effectively reduced, and the development difficulty is greatly reduced.

The embodiment of the invention discloses a control method and a control device of an electromagnetic valve and a vehicle. The control method of the electromagnetic valve comprises the following steps: generating a pulse width modulation signal, wherein the frequency of the pulse width modulation signal is greater than the response frequency of the electromagnetic valve and the duty ratio of the pulse width modulation signal fluctuates periodically within a preset range so that the pulse width modulation signal is superposed with a flutter signal; and transmitting the pulse width modulation signal to the electromagnetic valve. The amplitude and the frequency of the flutter signal are respectively determined by modulating the periodic fluctuation of the duty ratio of the PWM signal in the preset range and the fluctuation range and the fluctuation period of the duty ratio of the PWM signal so that the independent adjustment of the amplitude and the frequency of the flutter signal is realized. The algorithm of the flutter signal is simplified on the basis of solving the problem of hysteresis of the electromagnetic valve due to static friction of the system or parasitic flutter of theelectromagnetic valve, and the amplitude and the frequency of the flutter signal can be better controlled according to the requirement of the hysteresis effect and thus the dynamic response characteristic of the electromagnetic valve can be better improved.

The embodiment of the invention discloses a reconstruction method for an aerodynamic model in a ground flutter test, and the reconstruction method achieves the reconstruction of the aerodynamic modelthrough modal shape excitation, physical coordinate and modal coordinate conversion and interpolation point polycondensation. In a reconstruction process, vibration mode excitation is carried out on amode participating in flutter; obtaining an aerodynamic force influence coefficient matrix under frequency domain and modal coordinates; the method comprises the following steps: performing conversion of physical quantities under modal coordinates and physical coordinates, performing optimal configuration on interpolation points through a genetic algorithm to realize polycondensation of the interpolation points, and finally fitting frequency domain aerodynamic force to a time domain through a rational function fitting mode to obtain an aerodynamic force model meeting the requirements of a ground flutter test. The reconstruction method provided by the embodiment of the invention has the advantages that the precision is high, the calculation efficiency is high, and the model reconstructionprocess is simple and feasible; the test precision of the ground flutter test in a transonic region is improved, so that the effect of expanding the application range of the ground flutter test is achieved.

The invention relates to a method for measuring a full-motion horizontal tail rotating mode of an airplane, which comprises the following steps of: firstly, identifying an initial dimensionless frequency of a full-motion horizontal tail rotating mode by adopting a phase separation technology, and then obtaining a force frequency curve of the full-motion horizontal tail rotating mode by adopting aphase resonance mode extraction technology near the frequency and gradually increasing exciting force from small to large; when the force frequency curve of the full-motion horizontal tail rotating mode is obtained, processing the force of the exciting force by a resultant moment M. According to the method, the nonlinear influence of friction and gaps on the full-motion horizontal tail rotation mode can be more completely considered, the force-frequency curve of the full-motion horizontal tail rotation mode measured by the method is used as the input of horizontal tail flutter calculation, theflutter calculation can be considered more comprehensively and reasonably, and the flutter safety of an airplane is effectively guaranteed.

The invention provides a system and a method for testing the dynamic stiffness of an airplane horizontal tail servo actuating mechanism. The system comprises ground maintenance and detection equipment(1), a flight control computer (2), a system tester (3), a dynamic stiffness test bed and the like (4), wherein the dynamic stiffness test bed (4) comprises a dynamic stiffness test bed body (41) anda dynamic stiffness test bed control system (42). The method comprises the following steps: drawing a required dynamic stiffness amplitude-frequency characteristic curve through loading force and corresponding position feedback, and finding out the lowest point of the dynamic stiffness amplitude and the corresponding frequency value according to the drawn characteristic curve. The system can truly and effectively measure the dynamic stiffness amplitude-frequency characteristic curve, and according to the testing method, the lowest point of the dynamic stiffness amplitude of the horizontal tail servo actuator and the frequency corresponding to the lowest point can be found out, and necessary data support is provided for analyzing the airplane servo flutter phenomenon and optimizing the horizontal tail servo actuator and a gap eliminating connecting mechanism between the horizontal tail servo actuator and a control surface.

The invention discloses a method for optimizing static aeroelasticity and flutter of an all-moving horizontal tail. Firstly, a finite element model is established, then static aeroelasticity, flutterand balance weight are optimized one by one, the matching degree is high, the minimum-weight full-motion horizontal tail which meets the requirements for static aeroelasticity and flutter and does notaffect the appearance of the horizontal tail is obtained, constraint conditions are met, and the problems of static aeroelasticity and flutter of an airplane are effectively solved.

The invention discloses a method for solving the linear flutter speed of a three-dimensional wing based on a multi-body system transfer matrix method. The method comprises the following steps: deriving a bending-torsion coupling beam transfer matrix of the wing based on the multi-body system transfer matrix method; establishing a total transfer equation of the wing, and solving and obtaining the inherent frequency and the vibration mode of the wing; according to the Theodorsen unsteady aerodynamic force theory, establishing the relation between wing vibration displacement and aerodynamic force; establishing a body dynamics equation of the wing, and converting the body dynamics equation into a frequency domain to obtain a flutter frequency domain model of the wing; and performing frequencydomain solving on the wing flutter frequency domain model to obtain the flutter speed of the wing. The method can achieve the quick solving of the flutter speed of the wing.

The invention belongs to the field of structural mechanics and relates to a method for determining the dimension of a rectangular beam section of a low-speed flutter main beam model of an airplane. The method is characterized by adapting to the situation that the half width a of a rectangular section is smaller than the half height b of the rectangular section or the situation that J / Ix<1.69, wherein Ix refers to vertical inertia moment and J refers to polar inertia moment. The method for determining the dimension of the rectangular beam section with a lug piece comprises the following steps of: calculating the half width ar and the half height br of the rectangular section when the rectangular beam without the lug piece has the vertical inertia moment Ix of a preset value and the polar inertia moment J of a preset value; adjusting the polar inertia moment J with the preset value; calculating the half width a and the half height b; and calculating the half width 1 of the rectangular beam section with the lug piece when the rectangular beam with the lug piece has lateral inertia moment Iy of a preset value and the thickness of the lug piece is t. The method disclosed by the invention has the advantages of increasing the rigidity precision of the section of the model, reducing the uncertainty of model design, shortening the time for determining the dimension of the section, and increasing the design efficiency of the flutter model.

The invention provides a folding rudder capable of monitoring and optimizing the flutter phenomenon. The folding rudder comprises a folding module, a monitoring module, a processing and analyzing module and a driving module. According to the folding rudder, the flutter phenomenon is relieved or eliminated in a passive monitoring and active intervention mode, the real-time performance is achieved, the vibration condition of the folding rudder in the flight process of a guided weapon can be analyzed more accurately, and in addition, when the folding rudder of the guided weapon flutters, the problem can be solved in an active intervention mode, so that structural damage is avoided; and therefore, the vibration condition of the folding rudder in the flight process of the guided weapon can be monitored in real time, in addition, a computer analyzes and processes the data, so that the vibration of the folding rudder is relieved or counteracted in an active intervention mode, and therefore the flutter phenomenon is effectively prevented. The folding rudder has the advantages of being low in design cost, high in real-time performance, high in reliability, convenient in installation and free from influence on the original structure and equipment of the guided weapon.

The invention belongs to the field of aviation structural mechanics, and relates to a design method for obtaining the overweight ratio of a flutter model in the design of the flutter model. The method comprises the following steps of: calculating the equivalent mass mBE of a reference beam frame according to a reference beam frame model, and calculating the equivalent mass ratio nBE of the reference beam frame in combination with the component mass mA; calculating a design scale coefficient ndss; specifying a section shape coefficient ns; specifying the mass ratio kb of a flutter model beam; and calculating the overweight ratio nG of the flutter model. In According to the method disclosed by the invention, a reverse design method is used for replacing a trial and error method adopted in the conventional model design, so that various factors possibly influencing model weight and mathematical relations thereof are considered at the start of model design, the overweight ratio of the model is further obtained, the uncertainty of model design is reduced, and the model design period is shortened greatly.

The invention discloses an unsteady flow control method based on flutter winglets, and belongs to the technical field of fluid machinery. A flutter winglet is arranged near a separation point where a controlled pneumatic component generates flow separation, the flow direction of the flutter winglet is within the range of 5% L in front of and behind the separation point, L is the characteristic length of the controlled pneumatic component, and the distance between the spanwise direction of the flutter winglet and the separation point is 5%-50% of the height h of a separation area. Energy sources such as an external air source or a power source are not needed, a complex air path or a circuit system is not needed, energy can be extracted from a main stream to generate flutter only through the structure of the device, and therefore unsteady excitation used for restraining flow separation is generated, performance indexes such as the pressure ratio, the efficiency, the total pressure recovery and the stability margin of a controlled pneumatic component are improved, and the reliability of the controlled pneumatic component is improved. The system has the advantages of simple structure, no external energy source, high engineering practicability and the like.

The invention discloses a design method for improving aerodynamic elastic stability based on a beam structure of an ultrahigh-speed aircraft. The invention aims to solve the problems that flutter boundaries of conventional uniform beam structures of aircrafts in the prior art are insufficient, flutter is prone to occurring, and disastrous structural damage is often caused by flutter. The method comprises the steps that the basic size of a beam structure is designed according to the use requirement of the beam structure in ultra-high-speed flight, and meanwhile material attributes and materialparameters are determined; under the action of the determined aerodynamic force, MATLAB simulation software is used for programming simulation, and the chatter order and the chatter boundary of the uniform beam are obtained; based on the chatter order and the chatter boundary of the uniform beam, the aeroelastic stability of the beam structure is improved through an additional concentrated mass method, a spring constraint method or an axial functionally graded material design method. The invention belongs to the field of aerospace.

A decoy body that is configured with one or more flutter assemblies that make the decoy appear as though the wings / feathers are moving. The flutter assemblies include a suspension arm and lightweight flutter elements configured to be operable in very low wind conditions. The flutter assembly can be used on new or existing decoys. Movement of the wings attracts live game to the decoy spread.

The invention provides a multi-loop aircraft model cluster flutter restraining composite root locus multistage PID (Proportion Integration Differentiation) robust controller design method. The method comprises the following steps: directly determining to obtain a model cluster constructed by amplitude frequency and phase frequency features in a full envelope by means of frequency sweeping flight test under the condition that different heights and Mach numbers are given; giving a closed loop pole distribution limitation index under corresponding root locus description through multi-loop model equivalence according to an amplitude frequency margin in the flight envelope and a phase margin military standard requirement, determining the stage number and parameter value of a multistage PID robust controller by adding the multistage PID robust controller, the closed-loop pole distribution limitation index in the full envelope of an air craft and a model identification method in system identification; designing a low-altitude flight aircraft which is accordant with the full envelope, can be used for restraining flutter, has low overshot and is stable on the basis of the concept of closed-loop pole distribution limitation under root locus description.

The invention provides an aircraft flutter characteristic analysis method with no additional aerodynamic damping, which comprises: building an aircraft aeroelastic model at first, calculating the sub-critical damping and frequency in each mode at each flying speed point within the flying speed scope containing an aircraft flutter critical speed, drawing a change curve of the sub-critical damping and frequency along with the flying speed, and determining the aircraft flutter critical speed. The aircraft flutter analysis equation with no additional aerodynamic damping is introduced, and the damping iteration method is adopted for analysis of sub-critical damping characteristics according to the corresponding nonlinear characteristic value of the equation. Compared with the conventional flutter analysis method with the additional aerodynamic damping, no additional aerodynamic damping needs to be introduced, the obtained sub-critical damping characteristics of aircrafts are more accurate,the determination of the aircraft flutter characteristics can be facilitated, and a reliable basis can be provided for trial flutter flights.

Methods and systems for analyzing and predicting aeroelastic flutter on configurable aircraft are disclosed herein. The method may include the steps of: a) flying a known aircraft type above ground, wherein the aircraft has a payload in a known configuration; b) acquiring data from at least one sensor on the aircraft while flying above ground; c) repeating steps a) and b) with a different payload configuration; d) training a machine learning predictive model for the aircraft type for aeroelastic flutter using the collected data; and e) using the predictive model to predict when aeroelastic flutter may occur on the aircraft type when the aircraft has a payload in a new configuration for which data from sensors was not previously collected with the aircraft in flight.

The invention relates to an aeroelastic system flutter signal abnormal data expansion method. Based on a time sequence signal generation algorithm of a generator in a generative adversarial network, aflutter data set can be expanded through a pre-trained generator network by only needing a small number of flutter signals with coupling resonance occurring in structural response signals of an aeroelastic system or an aircraft, normal data and abnormal data are balanced as much as possible, and the influence of data samples on the method in the method verification process is reduced. Compared with the prior art, the flutter signal data set extension method based on the deep learning algorithm is provided by introducing the deep learning algorithm into the data set extension of the vibrationsignals of the aeroelastic system. Based on a pre-trained generative adversarial network, conversion from random signals based on Gaussian distribution to flutter signals with flutter characteristicsis achieved through a small number of actually-measured wind tunnel flutter test flutter signals of coupling resonance, and the purpose of expanding a data set is achieved.

The invention relates to a method for expanding abnormal data of flutter signals of an aeroelastic system. Based on a time series signal generation algorithm of a generator in a generative adversarial network, only a small amount of flutter signals with coupled resonance appear in structural response signals of aeroelastic systems or aircraft. It is possible to expand the flutter data set through the pre-trained generator network, so as to balance the normal and abnormal data as much as possible, and reduce the impact of data samples on the method during the method verification process. Compared with the prior art, the present invention introduces the deep learning algorithm into the data set expansion of the vibration signal of the aeroelastic system, and provides a flutter signal data set expansion method based on the deep learning algorithm. Based on the pre-trained generative adversarial network, through a small amount of coupled resonance measured wind tunnel flutter test flutter signals to realize the transformation from random signals based on Gaussian distribution to flutter signals with flutter characteristics, so as to achieve the purpose of expanding the data set.

The invention relates to an aircraft dynamics monitoring method based on frequency response analysis, and belongs to the technical field of aeroelastic test. The aircraft dynamics monitoring method based on frequency response analysis comprises the following steps: step one, acquiring the reference frequency fri and the inspection frequency fti of an aircraft structure in an ith-order mode; and step two, acquiring frequency weigh errors of an aircraft flutter characteristic according to a weight error formula. According to the external field dynamics monitoring method based on frequency response analysis, dynamics monitoring can be carried out on an external field by simple frequency examination through application frequency response function analysis and mode weight error analysis, thus,the dynamic characteristic of an aircraft in a using process is obtained and is applied to assessment of the flutter characteristic of the aircraft, and finally, the use envelope of the aircraft is given to guarantee flight safety of the aircraft.

The invention belongs to the field of aeroelastic mechanics, and particularly relates to an emergency method for flutter suppression in a low-speed flutter wind tunnel test. The emergency method comprises the steps that 1 an emergency system for flutter suppression is constructed, the emergency system comprises a low-speed flutter model and a skin damage mechanism, the low-speed flutter model comprises a frame section and a dimensional skin, and the skin damage mechanism is installed at a flutter sensitive part in the low-speed flutter model; and 2 when flutter occurs, the dimensional skin of the low-speed flutter model is damaged through the skin damage mechanism. According to the emergency method for flutter suppression in the low-speed flutter wind tunnel test, an emergency flutter suppression measure is provided, when flutter occurs, the skin damage mechanism is adopted to damage the dimensional skin of the low-speed flutter model, and the purpose of protecting the safety of the overall structure of the model at a low cost is achieved. The method is clear in principle and easy to implement, and can protect the model safety when other reversible flutter suppression measures are invalid.

The application belongs to the field of flutter wind tunnel tests, and in particular relates to a flutter model protection method and structure based on safe and controllable damage. The protection method includes: obtaining a flutter wind tunnel model, installing an anti-flutter structure on the flutter sensitive part of the flutter wind tunnel model, and the anti-flutter structure is designed according to the following principles: when flutter occurs, the The anti-flutter structure is prior to the destruction of other parts of the flutter wind tunnel model; after the anti-flutter structure falls off from the flutter wind tunnel model, the flutter speed of the flutter wind tunnel model will increase; After the anti-flutter structure falls off from the flutter wind tunnel model, it will not cause severe wind tunnel damage to the flutter wind tunnel model. This application designs the anti-flutter structure, so that the flutter wind tunnel model can achieve the protective effect under the premise of ensuring normal blowing, especially suitable for continuous wind tunnels, which is conducive to safer and smoother flutter model wind tunnel tests Research.

The invention belongs to the field of aviation structural mechanics, and relates to a design method for obtaining the overweight ratio of a flutter model in the design of the flutter model. The method comprises the following steps of: calculating the equivalent mass mBE of a reference beam frame according to a reference beam frame model, and calculating the equivalent mass ratio nBE of the reference beam frame in combination with the component mass mA; calculating a design scale coefficient ndss; specifying a section shape coefficient ns; specifying the mass ratio kb of a flutter model beam; and calculating the overweight ratio nG of the flutter model. In According to the method disclosed by the invention, a reverse design method is used for replacing a trial and error method adopted in the conventional model design, so that various factors possibly influencing model weight and mathematical relations thereof are considered at the start of model design, the overweight ratio of the model is further obtained, the uncertainty of model design is reduced, and the model design period is shortened greatly.

To overcome the problem of a complex flutter model under influence of aerodynamic force and intensity change cannot be effectively expressed in the prior art, the invention provides an aircraft flutter analysis network model Laguerre modeling method. According to the method, a plurality of grid points are selected in aircraft body shafting, a complex flutter network model ix expressed according toa body shafting decomposition method under influence of aerodynamic force and intensity change of different flight speeds, atmosphere density, airflow environment and different temperatures, requirements of installing a sensor and recording data and images are proposed according to a requirement of establishing the model, data are obtained through an effective flutter flight test, an excitation function is obtained through a measured value of an airflow sensor, a Laguerre function is adopted to perform approximate and equivalent description of a vibration variable, solving of three axial vibration equations at the body shafting coordinate grid points is determined according to an identification method at the same time, and the technical problem that a complex flutter model under influenceof aerodynamic force and intensity change cannot be effectively expressed in the prior art.

The invention relates to a method for rapidly obtaining the speed of transonic flutter of a composite airfoil surface. The method comprises the following steps of: on the basis of the subsonic flutter optimization, selecting a finite number of airfoil surface sensitive areas to be known as sensitive elements, describing mechanical characteristics of the sensitive elements by using three-dimensional rigidity, further establishing three-dimensional rigidity feasible regions of the sensitive elements, selecting characteristic points in the rigidity feasible regions to calculate the unsteady aerodynamic force flutter, and respectively establishing transonic flutter speed response surfaces between each sensitive element and the whole composite airfoil surface according to a calculating result. In the subsequent flutter resolving, analyzing and optimizing process, the transonic flutter speed of the airfoil surface is obtained by inquiring the response surfaces, thus the long-time unsteady aerodynamic force flutter is replaced and the purposes of reducing the calculated amount and shortening the computing time are achieved.