56 results about "Helicopter dynamics" patented technology

The invention relates to a simple rotor mechanism of a coaxial dual-rotor helicopter test stand. The rotor mechanism comprises an inner shaft, an outer shaft, an actuator group, a lower control system, an upper control system, an upper rotor propeller hub and a lower rotor propeller hub, wherein the upper rotor propeller hub comprises a first propeller hub seat, and a first paddle arm arranged on the first propeller hub seat in a full articulation manner; the lower rotor propeller hub comprises a second propeller hub seat and a second paddle arm arranged on the second propeller hub seat in a full articulation manner; the upper control system is connected with the first paddle arm; the lower control system is connected with the second paddle arm; the upper control system is connected with the lower control system via an intermediate draw bar; and the actuator group is connected with the lower control system to control synchronous motion of the first paddle arm and the second paddle arm. The mechanism can achieve collective pitch control and cyclic pitch change of coaxial dual rotors, and effectively achieves various dynamic tests of the helicopter such as upper and lower rotor aerodynamic characteristics of the coaxial dual rotors, dynamic stability of the rotors, and moving part loads of the rotors.

The invention provides a coaxial rigid rotor coupling pneumatic elastic response analysis method, and belongs to a helicopter kinetic analysis technology. The method is characterized in that according to the operation and control characteristics of a coaxial rigid rotor helicopter, a balance equation optimization solution method is built, and an optimization solution is found according to the given target and constraint conditions. According to the characteristic of complicated coaxial rigid rotor flow field environment, a rotor flow field is calculated by using a computational fluid dynamic method based on an Euler / N-S equation; and then, the coupling pneumatic elastic response is calculated by using a computational fluid dynamic / computational structural dynamic loose coupling analysis method. The coaxial rigid rotor coupling pneumatic elastic analysis method has the advantage that when the analysis method is used, balance operation and control parameters, blade response, rotor loads and the like of the coaxial rigid rotor helicopter can be calculated. The method has good analysis precision and engineering applicability; the relying of the development process on the test can be reduced; and the design period and the development cost can be greatly reduced.

The invention belongs to the field of miniature rotor type unmanned aerial vehicle autonomous flight control and research, and provides a method for controlling an unmanned aerial vehicle and for making an unmanned helicopter quickly and accurately stabilized and controlled. A controller less depends on prior knowledge of models and has good robustness on uncertainty of a system. According to the technical scheme, a nonlinear robust control method of the posture of the single-rotor unmanned helicopter based on fuzzy feedforward includes the following steps that firstly, the form of a single-rotor unmanned helicopter dynamical model and the form of a single-rotor unmanned helicopter posture dynamical model are as follows; secondly, the nonlinear semicontinuous robust controller based on fuzzy feedforward is designed (img file=' DDA0000463767750000011.TIF' wi=' 754' he=' 72' / ) ( / maths). The method is mainly applied to the design and manufacture of the miniature rotor type unmanned aerial vehicle.

The invention discloses a four-rotor flight control method based on a sliding mode variable structure. First, based on the fact that controlled variables respond to control input at different speeds, the controlled variables are classified as very fast variables, fast variables and slow variables, and a four-rotor unmanned helicopter dynamics model is divided into three subsystems, namely, a very fast loop system, a fast loop system and a slow loop system; then, a control law based on a dynamic inversion control method is designed for the very fast loop system, the fast loop system and the slow loop system respectively; and finally, a control law based on a variable sliding mode structure control method is designed for the very fast loop system, the fast loop system and the slow loop system respectively to enhance the robustness of the system to parameter perturbation and external disturbance. The method of the invention can improve the static characteristic and dynamic characteristic of four-rotor unmanned helicopters.

The invention discloses a synthetic modeling method of rotor and airframe coupling dynamics modality, and belongs to the dynamic design technology of helicopters. The synthetic modeling method of the rotor and airframe coupling dynamics modality is suitable for the dynamic stability and dynamic response analysis modeling of single-rotor and multi-rotor coupling. The synthetic modeling method analyzes the air resonance instability of a rotor blade by aiming at the rotor blade operation pneumatic environment of a flight state stipulated by a flight envelope rule which needs to considers to cover and properly exceed the performance requirements of the helicopter, researches the influence of high-speed high-overload flight state parameters, looks for the structural design parameters of rotors, blades and a steering control system which cause or affect the generation of the air resonance, and determines the instability boundary of the air resonance. Through the sensitivity analysis of the air resonance by the structural design parameters, a structural design parameter range which affects the air resonance is determined, and a design measure for expanding the security boundary of the air resonance is put forward.

The invention provides an improved particle swarm algorithm optimized fuzzy PID unmanned helicopter attitude control method. The method comprises steps as follows: S1, a fuzzy PID attitude controlleris designed according to unmanned helicopter dynamics and kinematics model obtained with a mechanism modeling method, an error of an unexpected attitude angle and an actual attitude angle and an errorchange rate are controlled by a controller, and a parameter adjustment amount of the fuzzy PID attitude controller is obtained; S2, a quantification factor and a scaling factor in the fuzzy PID attitude controller are optimized with an improved particle swarm algorithm; S3, the optimized quantification factor and the optimized scaling factor are assigned to the PID attitude controller.

The invention discloses a helicopter / engine comprehensive simulation model. The invention introduces a spin state rotor rotordynamic differential equation into the helicopter / engine comprehensive simulation model, thereby accurately simulating helicopter dynamic change in a helicopter spin training process. The invention further discloses a method and a system for controlling a turboshaft engine based on the model. In a normal flying stage, a spin starting stage and a stable spin stage, and before a clutch is closed in a spin recovery stage, a univariance controller using fuel oil flow as a control variable is used for controlling the turboshaft engine; and after the clutch is closed in the spin recovery stage, a bivariate controller using fuel oil flow and a guide vane angle as control variables is used for controlling the turboshaft engine. The helicopter / engine comprehensive simulation model provided by the invention can make the overshoot of a free turbine rotating speed of the turboshaft engine in the helicopter spin training process not more than 3%, and greatly improves dynamic quality of the fuel oil.

The invention discloses an elastomeric shimmy damper model and application thereof to a helicopter system and belongs to the field of helicopter dynamics design. The elastomeric shimmy damper model is applied to dynamic stability design of the helicopter system by the aid of an excitation frequency modifying formula, and the application specifically includes: firstly establishing a balance equation of a rotor and helicopter body coupling system of a helicopter with an elastomeric shimmy damper; then modifying a blade lagging natural frequency under the condition of single frequency; solving dynamic displacement of the elastomeric lag damper under the condition of forward flight and modifying an excitation frequency of the elastomeric lag damper and a blade shimmy natural frequency under the condition of double frequency; and finally solving modal damping of a rotor and helicopter body system under the condition of the double frequency so as to judge dynamic stability of a star flexible hub helicopter system. The excitation-frequency modified elastomeric shimmy damper model has high applicability, can be used under the condition of the single frequency or the double frequency and can be applied to design of hinged, hinge-less and bearing-less rotor helicopters with the elastomeric shimmy damper.

The invention discloses a device for improving the aeroelastic stability of a bearing-free rotor and a design method of the device, and belongs to the field of helicopter dynamics. According to the invention, a shimmy pin is added between a rotor hub and a torsion sleeve. The design method comprises the following steps of: firstly determining design variables and the upper limit and the lower limit of each design variable; dividing each design variable at equal intervals between the upper limit and the lower limit into a plurality of design points so as to form a design space; solving the aeroelastic stability of the bearing-free rotor corresponding to each design point in the design space, thereby obtaining parameters which represent the aeroelastic stability of the bearing-free rotor and correspond to the design point one by one in the design space; and finally solving the function taking the upper limit and the lower limit of the design variables as a restriction condition, thereby obtaining a bearing-free rotor structure with the optimum aeroelastic stability. According to the invention, the arrangement of the shimmy pin can restrict the movement of the torsion sleeve, and the shimmy damping of a paddle can be increased, and further the aim for improving the aeroelastic stability of the bearing-free rotor is achieved.

The invention belongs to the technical field of helicopter dynamic tests, and discloses a helicopter flight control system transmission characteristic test method. The method comprises the steps of respectively performing sweep frequency excitation on a rotor time domain total distance and a periodic variable pitch through a control system, operating an excitation signal through a hydraulic actuator cylinder, and measuring the rotor time domain total distance and the periodic variable pitch under excitation; and performing FFT on the time domain total distance and the periodic variable pitch of the hydraulic actuator cylinder and a paddle to obtain a control system total distance and periodic variable pitch frequency domain transfer function. The method can be used for providing data for establishing a helicopter flight control system mathematical model based on tests, and providing basic technical support for establishing a helicopter rotor and body coupling stability model considering a flight control system and stability comprehensive analysis. The method has a certain engineering reference value for design and analysis of ground resonance and air resonance of advanced helicopters in China, and has far-reaching significance for guaranteeing safe flight and stability improvement of modern advanced helicopters.

Disclosed is a coaxial double-rotor unmanned helicopter modeling method based on optical cable laying. The method comprises flexible optical cable dynamics modeling at a constant tension release state and an optical cable laying unmanned helicopter dynamics model for a flight control system. The flexible optical cable dynamics modeling at the constant tension release state comprises stress calculating method of a single optical cable segment and a whole optical cable dynamics model establish method. The optical cable laying unmanned helicopter dynamics model for the flight control system comprises an unmanned helicopter platform dynamics model and an unmanned helicopter parameter identification test system, and it is set that: a), a whole optical cable to be laid is composed of N optical cable segments; b), each segment is sequentially released from a release mechanism, and the speed of the currently releasing optical cable segment is consistent with the speed of the just released optical cable segment; c), each optical cable segment is taken as a rigid rod, and the mass is concentrated at one end point of each optical cable segment; and d), the rigid rods are connected through twists. The coaxial double-rotor unmanned helicopter modeling method provided by the invention provides support for development of the flight control system of an unmanned helicopter.

The invention discloses a method and a device for identifying unmanned helicopter kinetic parameters. The method comprises steps that initial estimates of to-be-identified parameters are acquired, a first system matrix and a first control matrix of the initial estimates containing the to-be-identified parameters are acquired, a response function of the system measurement data is acquired according to the acquired data, a system matrix containing the optimal solution and a coefficient matrix are acquired by utilizing a state space mode, the response function and a cost function according to the flight data, the first system matrix and the first control matrix, in combination with the update strategy, a second system matrix and a second control matrix after update are acquired; whether the second system matrix and the second control matrix satisfy the preset condition is determined, if not, the second system matrix and the second control matrix are taken as the first system matrix and the first control matrix to re-actuate the previous steps till the preset condition is satisfied, and present values of to-be-identified parameters in the second system matrix and the second control matrix are taken as identification accomplishment target values.

The invention discloses an instant automatic releasing device and belongs to testing devices of the technical field of helicopter dynamics. The instant automatic releasing device is used for achieving the instant automatic releasing function when an anti-crashing test is carried out on a whole helicopter or components of the helicopter. The instant automatic releasing device comprises a bearing frame, an electromagnet, a connecting rod mechanism, a claw hook and a movable plate. The movable plate is of a strip-shaped structure and arranged between two hook-type pieces in the claw hook. Gears on the two sides of the movable plate are meshed with gear discs in the two hook-type pieces of the claw hook. When the claw hook is in an opened state, the movable plate can be pushed and triggered to enable the claw hook to be closed automatically. When the claw hook is in a closed state, the electromagnet is powered on and triggered, and the claw hook is made to be in an opened state through the connecting rod mechanism. The instant automatic releasing device achieves the automatic releasing function of a tested piece in the anti-crashing test of the whole helicopter or the components of the helicopter, and test risks are lowered. The instant automatic releasing device has the advantages of being high in control precision, high in bearing capacity, stable and reliable in operation and the like, and hook opening operation and hook closing operation are smooth and stable.

The invention relates to a helicopter control loading simulator having touch guidance. The helicopter control loading simulator is characterized in that a table board base is composed of a lower bottom board of the table board base and a table board of the table board base, wherein the top surface of the lower bottom board of the table board base is fixedly connected with the table board of the table board base, a fixed base is fixedly connected to the top surface of the table board of the table board base, a transverse motion servo motor is fixedly connected with the table board of the table board base though the fixed base, a transverse motion plate-type reducer is arranged at the front end of the transverse motion servo motor, an outer lug of the transverse motion plate-type reducer is connected with the fixed base, a rotating bracket is of a right-angle-shaped structure, the bottom surface of the rotating bracket is fixedly connected with a flange face of the transverse motion plate-type reducer, and the helicopter control loading simulator can accurately provide touch guide force in real time according to a helicopter kinetic model and an operating lever model, so that the high-fidelity immediate force sense is simulated. Meanwhile, aiming at different helicopter types, the programming method can be adopted for simulating the small differences of the helicopter types in the aspect of force sense, and thus the helicopter control loading simulator has good adaptation and universality.

A three-degree-of-freedom helicopter anti-saturation attitude tracking control method includes the steps of: 1. giving an expected tracking value: an expected attitude angle x<d>=[rho<d>epsilon<d>]<T>; 2. performing attitude angle tracking error calculation: an error between the expected attitude angle and an actual attitude angle z<1>=x-x<d>; 3. performing virtual control law design: designing v<d> and solving a derivative of v<d> to time as a state quantity in a neural network; 4. performing neural network design: the formula is as described in the specification, approaching an uncertain term in a 3-DOF helicopter dynamics equation; 5. performing auxiliary control system design: designing a state quantity u, and solving the problem of input saturation; 6. performing model control law design: utilizing the state quantity u of the auxiliary control system to correct an attitude angle speed tracking error term, and combined with the neural network, estimating the uncertain term, thereby finally obtaining a model control law; and 7. applying the model control law to a 3-DOF helicopter nonlinear model to perform attitude angle tracking simulation. The method can approach the model uncertain term, inhibit external disturbance influence, resist execution mechanism saturation and track any expected attitude, thereby guaranteeing asymptotic stability of a system, and simplifying a calculation process.

The invention discloses a helicopter rotor and fuselage coupling stability modeling method, and belongs to the helicopter dynamics modeling and analysis technology. The method includes: adopting a modal synthesis technology to perform comprehensive modeling on a rotor and fuselage dynamics coupling system, describing the motion of a fuselage and rotors in different coordinates, establishing structural dynamics finite element method models of isolated rotor blades and an airframe structure respectively, adopting an aerodynamic model to model aerodynamic force, and deducing a rotor / airframe coupling system kinetic equation by applying the Hamilton principle; establishing an automatic flight control system control model, and deriving an aerodynamic load matrix related to the pitch variable, thereby establishing a rotor wing and aircraft body coupling aeroelastic stability analysis model considering the flight control system; and finally, solving a characteristic value through a characteristic value method, and judging the stability of the coupling system through a characteristic value solution. The model can be used for calculating and analyzing the'ground resonance 'and'air resonance' stability of all advanced helicopters, and key technical support is provided for model design and development.

The invention discloses a layout for an inter-blade shimmy damper of a rotor wing of a helicopter, belonging to the field of the dynamic design of helicopters. According to the layout, the inter-blade shimmy damper is connected between the front edge of a previous rotor blade and the rear edge of a next rotor blade; and an included angle between a mounting support arm of the inter-blade shimmy damper and a propeller hub plane is located between -180 DEG and -150 DEG. With the adoption of the layout, the shimmy regressive lag modal damping of the rotor wing of the helicopter is increased, and the steady load of the inter-blade shimmy damper is reduced, so that the dynamic stability of a rotor wing system can be improved.

The invention discloses a four-degree-of-freedom helicopter dynamic flight simulator. Simultaneous simulation of three modes of space yaw, rolling and pitching can be achieved under the centrifugal field, and overload simulation is achieved; mounting components on a rotating arm are arranged in a centralized manner, a balancing design method is adopted, a counterweight unit is designed according to the trimming principle and mounted on the rotating arm, the mass center of the dynamic flight simulator is placed on the revolution axis, and vibration of the simulator caused by imbalance is reduced. The asymmetric rotating arm structure is adopted to reduce the rotational inertia of the dynamic flight simulator, and energy consumption of the dynamic flight simulator is reduced; a transmissionunit is adopted to prevent vibration of a main reduction motor from being directly transmitted to the rotating arm to make the entire dynamic flight simulator vibrate excessively, the unbalanced forceof the rotating arm and the rotating arm mounting components is transmitted to the civil part through the transmission unit, and running stability of the main reduction motor is enhanced.

The invention discloses a design method capable of expanding an air resonance security boundary of a helicopter, relates to a design method capable of expanding the security boundary of air resonance appearing in the helicopter in the flying speed and force lifting overloading state, and belongs to the technical field of dynamic design of the helicopter. According to the rotor blade operation pneumatic environment of the flying state stipulated by a flight envelope which needs to consider a cover and properly exceeds the performance requirement of the helicopter, unstability of the air resonance is analyzed, and the pitch control stiffness is improved to the air resonance insensitive range; the width of paddle rear edge adjusting sheets and distribution of the paddle rear edge adjusting sheets in the expanding direction are increased, so that an equivalent aerodynamic center moves backwards; and the paddle aeroelasticity coupling characteristics are improved so as to expand the air resonance security boundary.

The invention provides a dynamic load measurement device and relates to the technical field of helicopter dynamic testing. The device comprises an upper plate, eight single-dimension mechanisms and abaseplate. The upper plate and the baseplate are arranged in parallel, and four lug sheets are arranged in the middle of one side of the upper plate and connected with the four single-dimension mechanisms; the other ends of the two single-dimension mechanisms are connected to the two corners, at the same side of the upper plate, of the baseplate, the other end of each of another two single-dimension mechanisms is connected to the middle of the baseplate, in a similar way, the other four single-dimension mechanisms are installed between the middle of the other side of the upper plate and the baseplate, and a load sensor is arranged on each single-dimension mechanism. The eight load sensors form measurement circuits of eight Wheatstone bridges, and dynamic loads in all directions can be obtained through decoupling calculation. The dynamic load measurement device is high in practicability, has the advantages of high precision, low coupling and the like and can be successfully applied to falling vibration tests and anti-crash tests of helicopter undercarriages of various types.

The invention discloses an unmanned helicopter flight control improved PSO algorithm verification method based on embedded hardware. Based on flight control, the PSO algorithm is improved. An unmannedhelicopter simulation computer is in communication connection with an unmanned helicopter airborne flight control computer kernel through a serial port, and is used for receiving steering engine control information sent by a flight control computer kernel. The unmanned helicopter dynamic model is operated at a fixed step length, the attitude and position state information of the unmanned helicopter is output in real time, so closed-loop control is realized; an MATLAB / SIMULINK platform is adopted, an embedded hardware object is connected into a test loop through a serial port, whether a control algorithm can meet the performance requirement or not can be verified, meanwhile, rapid parameter adjustment of a control system can be facilitated, and the development efficiency is improved. The verification method can directly verify the effectiveness of the configured flight control improved PSO algorithm parameters, and can provide reference for semi-physical simulation experiment parameterconfiguration and subsequent experiment flight decision.

The invention belongs to the technical field of helicopter dynamics, and relates to a method for adjusting vibration in a rotating surface of a rotor hub of a helicopter. The method comprises two parts, specially, the adjustment of a hub shimmy damper and the adjustment of a hub weight stack are carried out, when the vibration eccentric force in the rotating surface of the rotor hub reaches 60-100kg or the speed value of the vertical vibration in a cabin reaches 0.6-1 IPS, and the hub shimmy damper is adopted to adjust; when the vibration eccentric force in the rotating surface of the rotor hub is less than 60 kg or the speed value of the vertical vibration in the cabin is less than 0. 6g IPS, and the hub weight stack is adopted to adjust. According to the method, the defects of a currentsingle adjustment mode can be overcome, and the vibration environment of the helicopter is ensured to be kept at a comfortable level.

The invention belongs to a dynamic design technique of helicopters, relates to a dynamic modeling method of a multi-transmission rotor hub structure and aims to prevent ground resonance of a bearingless rotor and expand the air resonance boundary. The accurate modeling analysis and design means is provided in terms of rigidity configuration, displacement coordination for a flexible beam, a control lever, a damper and a sleeve; dynamic finite element modeling of a multi-transmission rotor hub structure is performed first, and degrees of freedom of non-dependent joints of component parts of the structure are expressed in joint degree-of-freedom linearity; according to the dependency and relativeness of the joint degrees of freedom of the component parts of the structure, the independent joints, the relative joints and the displacement coordination relation between the component parts of the structure and the connective joints are determined so as to establish a dynamic rigidity matrix of the structure. By the use of the dynamic modeling method, the key difficulties in the design and analysis of the complex multi-transmission rotor hub structure are broken through.

A helicopter dynamics simulation method based on Unignine comprises the steps that firstly, each component of a helicopter body is abstracted into a mass point, and the mass centers of all the mass points are calculated and serve as the mass center of a helicopter; and then a helicopter body motion equation, a rotor mathematical model and a tail rotor mathematical model are respectively constructed so as to obtain a dynamic model of the whole helicopter. According to the inherent attribute, the external environment and the current state of the helicopter to stimulate .stress of the helicopter,so that the helicopter can feed back the air environment in time. The attitude characteristics of the helicopter in flowing air are accurately reproduced, and the phenomena of load increase, load decrease, load movement and the like of the helicopter are accurately simulated by adopting a centroid dynamic calculation mode. Helicopter power under different speeds, different models and different environments can be simulated, and the helicopter attitude can be simulated more accurately.

The invention belongs to a tool designing technology, particularly relating to a main propeller hub and propeller shell assembly dismounting tool and method. The repairing technological level of helicopter dynamic components directly affects the quality of heavy repairing of helicopters in factories, and meanwhile, along with continuous improvement of the capacity of a company, repairing and maintaining operations of main propeller hub products are increasingly complicated and overloaded. A dismounting method and tool for disassembling a main propeller hub and propeller shell assembly of a helicopter and preventing parts from being damaged are required to be researched urgently so as to provide a convenient and effective working means for disassembling the main propeller hub and propeller shell assembly of the helicopter. At present, ready-made methods and tools do not exist, problems of deformation of upper and lower star plates, pulling damage of bearings, scratching damage of inner holes of the star plates and cross axles and the like are always caused in a disassembling process, parts are quite easily discarded as useless, and repairing quality and progress of products are affected. The invention provides the brand-new main propeller hub and propeller shell assembly dismounting tool and method. The tool is simple in structure and convenient to operate, and can meet requirements of dismounting of a main propeller hub, in a disassembling process, the bearings, the inner holes of the star plates and the cross axles cannot be damaged, the upper and lower star plates do not deform, and improvement of dismounting quality and dismounting efficiency of the propeller shell assembly during disassembling is facilitated.

The invention belongs to the technical field of helicopter vibration absorption, and relates to a helicopter dynamic vibration absorber. The vibration absorber comprises an upper supporting plate (1)and a lower supporting plate (2). One end of the upper supporting plate (1) and one end of the lower supporting plate (2) are fixed. The other end of the upper supporting plate (1) is fixed to the helicopter cockpit floor. A mass block is fixed to the other end of the lower supporting plate (2) and is suspended. The dynamic vibration absorber is small in structural size, simple and easy to install, can generate large antiresonance load, and has a good vibration absorbing effect. The service life of the vibration absorption structure is unlimited, and safety and reliability are achieved. When the dynamic vibration absorber is installed under the floor, a pilot is not disturbed, other control and system structures are not interfered, and the dynamic vibration absorber can well adapt to a helicopter.

A longitudinal helicopter rotor and fuselage coupling stability modeling method belongs to the helicopter dynamics modeling and analysis technology, adopts a fully-articulated rigid blade model, and comprises the following steps: firstly, establishing coordinate systems of all systems and relationships among the coordinate systems, establishing a fuselage motion model, a rotor motion model and anaerodynamic model; after rotor coupling response solving and multi-blade coordinate transformation, combining a rotor body undercarriage blade model mass damping stiffness matrix, establishing a tandem double-rotor helicopter ground resonance analysis model, calculating ground resonance by adopting a characteristic value method, and judging ground resonance stability through a characteristic valuereal part. The model can be used for calculating and analyzing the ground resonance stability of the tandem double-rotor helicopter, and key technical support is provided for model design and modification development.

The invention provides a helicopter moving part vibration signal data quality calculation method, and the method comprises the following steps: A, setting vibration sampling frequency and sampling time, and synchronously acquiring a vibration signal and a rotating speed signal; B, averagely dividing the vibration signal and the rotating speed signal into n sections, intercepting each section of vibration signal according to the synchronous rotating speed signal, and carrying out time domain average processing; C, performing FFT conversion on the n time domain synchronous average signals in sequence, and extracting amplitudes and phases of the n rotating speed same-frequency vibration signals; D, eliminating abnormal values of the amplitudes of the n rotating speed same-frequency vibrationsignals by utilizing a 3[delta] rule; E, calculating dynamic balance data quality according to the maximum value and the minimum value of the amplitude of the residual rotating speed same-frequency vibration signals; and F, if the mass of the dynamic balance data is smaller than the mass of the dynamic balance data, calculating the average value of amplitudes and phases of residual rotating speedsame-frequency vibration signals, otherwise, returning to the step A. The accuracy of the helicopter dynamic balance test result can be effectively improved, and the number of times of field dynamic balance adjustment is reduced.

The invention discloses a method for calculating the dynamic threshold value of a helicopter dynamic component based on Bayesian inference. The method comprises the following steps of 1) carrying outthe data preprocessing to filter noise in the data; 2) carrying out the feature extraction of respectively extracting features of normal data and fault data; 3) calculating the dynamic threshold, andusing Bayesian reasoning method to calculate the dynamic threshold of normal data and fault data. The invention has the advantages that in the feature extraction method, four methods of extracting inner loop feature frequency energy, extracting outer loop feature frequency energy, extracting ball feature frequency energy and extracting M6A are respectively adopted to carry out feature extraction on normal data and different types of fault data, so that the vibration characteristics of normal and fault data can be effectively reflected. In the calculation of dynamic threshold, the Bayesian reasoning method is used to consider not only the probability density of normal data, but also the probability density of different types of fault data, which improves the calculation accuracy of dynamicthreshold.