9851results about "Sustainable transportation" patented technology

The invention discloses an interior noise analysis and prediction method of a high speed train. The method comprises the following steps: establishing a train reconditioning train body model, a statistical energy analysis model of a body in white structure and an interior and exterior vocal cavity statistical energy analysis model, and carrying out simplification and subsystem division; obtaining the statistical energy analysis parameters of the train body structure and an interior vocal cavity model, and loading the statistical energy analysis parameters onto a train body structure model plate subsystem and a vocal cavity model subsystem; and obtaining exterior vocal excitation source energy borne on the train body, applying the exterior vocal excitation source energy onto the exterior vocal cavity statistical energy analysis model, causing the exterior vocal excitation source energy to reach an interior vocal cavity after the exterior vocal excitation source energy is attenuated by the sound insulation property of a structural plate in a body in white structure model so as to obtain structural noise energy radiated into the train by the reconditioning train body under the function of the two-line suspension force of a compartment, and then, carrying out interior noise analysis and prediction. The problem that the interior noise of the train is difficult in prediction and the problems of the upper limit boundedness of a frequency domain, complex calculation flow, incomplete motivation consideration in the traditional method are overcome, calculation efficiency and prediction accuracy are improved, and development and test cost is lowered.

A method for designing an engine case static structure of a gas turbine engine includes creating signals representing an engine case static structure knowledge base of information. The knowledge base has a plurality of design rule signals with respect to a corresponding plurality of parameter signals of associated elements of an engine case static structure, wherein the knowledge base comprises at least one data value signal for each of the plurality of design rule signals. A desired data value signal is entered for a selected one of the plurality of parameter signals of an associated element of the engine case static structure and compared to the corresponding data value signal in the knowledge base for the corresponding one of the plurality of design rule signals. Signals representative of a geometric representation of the selected parameter signal may be created.

In a flight simulator program, a flight dynamics editing module enables the user to input parameters for modifying an existing aircraft design or creating an aircraft design. Starting with the type and purpose of aircraft, the user is able to specify parameters defining the configuration, and various other aspects of the aircraft, including the number and type of engines, properties of the flight controls, type of landing gear, etc. Once the user has input the parameters, an aerodynamic coefficients generator module included with the flight simulation program determines aerodynamic coefficients for the aircraft design, using classical formulas and determining the coefficients in an appropriate order. The aerodynamic coefficients and certain parameters input by the user are then output as two flight model data files, in a format usable by the flight simulator program, so that the user can evaluate the aircraft design by flying it within the simulation.

A computer implemented method for visualization of a 3D model of an object, the method includes:generating and displaying a first view of the 3D model;receiving an user input, the user input are one or more interaction commands comprises interactions for understanding particular functionality of the 3D model, wherein functionality of the 3D model is demonstrated by automatic operation of the part / s of the 3D model which operates in an ordered manner to perform the particular functionality;identifying one or more interaction commands;in response to the identified command / s, rendering of corresponding interaction to 3D model of object with or without sound output using texture data, computer graphics data and selectively using sound data of the 3D-model of object; anddisplaying the corresponding interaction to 3D model,wherein operating in Ordered manner includes parallel or sequential operation of part / s.

A computer-aided method suitable for assisting in the design of an aircraft by providing the dimensioning aerodynamic forces and other relevant values comprising the following steps: a) Selecting a set of parameters of the aircraft, being said aerodynamic forces and other relevant values dependant of said parameters; b) Performing flow field CFD RANS computations for a number N1 of different combination of values of said parameters; c) Obtaining said aerodynamic forces and other relevant values for whatever combination of values of said parameters through a reduced-order model, generated by computing the POD modes of the flow variables, expanding the flow variables using said POD modes and obtaining the POD coefficients of said expanded flow variables using a genetic algorithm that minimizes the error associated to the expansion of the Navier-Stokes equations. The invention also refers to a system for carrying out said method.

The invention relates to a pneumatic elastic analytical method, and provides a pneumatic elastic mechanical characteristic analytical method of a hypersonic speed aircraft in thermal environment. According to the method, the piston theory is used for carrying out frequency domain nonsteady aerodynamic calculation on a full-aircraft finite element model of the hypersonic speed aircraft; on the basis, the pneumatic thermal effect on the model in the hypersonic speed thermal environment is considered; the weak coupling effect of pneumatic thermal input in aerodynamic input and elastic force input is ignored; only the structure temperature steady-state characteristics under the pneumatic thermal load effect are considered; after the steady aerodynamic calculation, the thermal flux density on the surface of the aircraft is obtained by adopting a reference enthalpy method; further, the steady temperature distribution on the surface of the aircraft is calculated; the practical equivalent rigidity matrix of the structure at the moment is obtained; and the critical flutter speed is calculated by an engineering method. The pneumatic elastic mechanical characteristic analytical method has the advantages that the pneumatic elastic analytical problem of the hypersonic speed aircraft in the pneumatic thermal environment is solved, and the pneumatic elastic performance of the hypersonic speed aircraft is improved through flutter speed analysis.

A PoF based method for calculating mission reliability of an electronic product comprises steps as follows: step one, information of all mission profiles of the product is collected, and an environment profile of each mission is determined; step two, thermal simulation and vibration simulation of environmental stress of each mission are performed, and a local response of the product to an environmental load is obtained; step three, a product simulation model is established; step four, simulation calculation of the product in all the mission profiles is completed, and the mean time to failure and a main failure mechanism of the product are obtained; and step five, the mission reliability of the product is calculated according to the mean time to failure. According to the PoF based method for calculating the mission reliability of the electronic product, all missions of the product during lifetime use are considered, the environmental stress of each mission is simulated, and the mean time to failure and the mission reliability of the product are comprehensively calculated. By means of a PoF model, the direction relation between parameters of a product material, structure, process and the like and the reliability can be obtained, and a design improvement direction is clearly and directly provided for the product.

A scenario is defined that including models of vehicles and a typical driving environment. A model of a subject vehicle is added to the scenario and sensor locations are defined on the subject vehicle. Perception of the scenario by sensors at the sensor locations is simulated to obtain simulated sensor outputs. The simulated sensor outputs are annotated to indicate the location of obstacles in the scenario. The annotated sensor outputs may then be used to validate a statistical model or to train a machine learning model. The simulates sensor outputs may be modeled with sufficient detail to include sensor noise or may include artificially added noise to simulate real world conditions.

The invention discloses a method for predicting the size of acceleration condition in-vehicle noise contribution. The method comprises the following steps that a vibration source generating in-vehicle noise and the transfer path thereof are measured so as to obtain a noise fitting model; and the excitation source or the transfer function in the noise fitting model is modified so as to predict the size of contribution of the corresponding noise excitation source or the transfer function on the in-vehicle noise. The in-vehicle noise model is established by using the measured acceleration excitation source and the noise transfer function thereof and correlation analysis is performed on the relevant excitation source, the path transfer function and the noise frequency band to determine the strongly correlated vibration source of the in-vehicle noise and the path transfer function thereof so as to perform troubleshooting, and the method is simple and rapid, the computing burden can be substantially reduced and the processing efficiency can be enhanced.

The invention discloses an optimization design method of fibre enhanced composite material marine propeller blades, which relates to an optimization design method of propeller blades and aims at solving the problem of imperfect design method of the fibre enhanced composite material marine propeller blades. On the basis of offsets data of the original high-speed metal propeller blades, the method is calculated by using a fluid-solid coupling method combining with the implementation of predeformation strategy, a blade structure is composed by mixed fibre composite materials, the surface of a blade adopts fibre glass enhanced composite material, the interior of the blade adopts the mixture of carbon fiber and Kelvar fiber enhanced composite material, and the specific mixed paving mode and the mixing ratio are determined according to the hydro elastic design result of a fibre enhanced composite material marine propeller. The optimization design method is used for designing the propeller blades.

The invention provides a digital ship model plane motion organization experimental method based on computation fluid dynamics CFD software FLUENT. The method includes applying a FLUENT pre-process software GAMBIT to establish a research object model and a control domain; arranging a triangle grid on the model surface, then arranging a unstructured grid in the control domain; setting the boundary condition, joining an user-defined function UDF document, introducing dynamic grid technology, using a finite volume method based on the complete unstructured grid, realizing a pure swaying movement, a pure rising and falling movement, a pure oscillatory movement, a pure pitching movement and a pure rolling movement carried on by a plane motion organization experiment; applying a science computation software MATLAB to do Fourier expansion of a force and a moment coefficient obtained by the FLUENT, being fitted by EXCEL least squares method, obtaining a hydrodynamic force coefficient of a vertical plane and a horizontal plane as well as the related hydrodynamics analysis by dimensionless. The invention carries on the digit plane motion organization experiment using the CFD software, can satisfy the request of hydrodynamic force data in the design initial period of a submersible.

The invention belongs to the technical field of aircraft simulation, and particularly relates to a simulation analysis method and system for gust response of an elastic aircraft. The method comprisesthe steps of according to the dynamics and the structure correspondence of the elastic aircraft under gust excitation, firstly, modeling an aircraft six-degree-of-freedom system, the atmospheric turbulence, the gust excitation and the aerodynamic force data according to the flight state; secondly, under the condition of static elasticity, analyzing the deformation of the aircraft wing under the aerodynamic force load and the change of an angle of attack and the influence of the angle of attack on an aerodynamic coefficient, correcting the aerodynamic coefficient to enable the aerodynamic coefficient to better conform to the real flight state of the elastic aircraft, and finally, connecting all the modules into a system to realize the corresponding analysis under the excitation of gust of the elastic aircraft. The system achieves a good effect in cases of a general aircraft and a civil large-scale passenger plane.

The present invention discloses a helicopter aerodynamic layout optimization method capable of reducing the adverse effect of aerodynamic interference. According to the method, aiming at the defects of a wind tunnel test and a CFD method and based on a discrete vortex method and a surface element method, a coupled aerodynamic interference high-precision analysis model for a helicopter is set up, so that the influence of aerodynamic interference on the helicopter can be considered only by simulating calculation without carrying out the wind tunnel test; compared with the CFD method, the method has the advantages of greatly reducing the requirements of a simulation model for computing resources and shortening computing time. In order to further reduce calculated amount and a optimization region, a helicopter fuselage shape parameter model is set up by a parametric method and a combined optimization method based on an agent model is adopted, so that the optimal helicopter aerodynamic outline and layout are obtained and the purpose of reducing aerodynamic interference to improve the flight characteristics of the helicopter is achieved. The helicopter aerodynamic layout optimization method can be applied to the conceptual design phase of the helicopter so as to improve the flight characteristics, the balancing characteristic, stability and manoeuvrability of the helicopter.

The invention relates to virtual prototype and MATLAB control system joint simulation with the aim of improving dynamic performance of vehicles. The technical problems solved by the invention are to provide a modified sky-hook control simulation technology of the four-axle heavy truck suspension system coupling the virtual prototype and MATLAB, and improve the road friendliness and smoothness of vehicles, aiming at the shortcomings of the prior art. The method of the invention is based on the ADAMS / View platform to build a four-axle heavy truck virtual prototype model and all grades of bilateral track multipoint random excitation pavement models; a modified sky-hook control algorithm is provided according to the modern control theory, ADAMS and Matlab / Simulink joint simulation are adoptedto realize the semi-active control and active control of modified sky-hook of suspension system damping. The important innovation of the invention is to realize the ADAMS / MATLAB joint simulation of the four-axle heavy truck suspension system; and the designed and modified sky-hook control algorithm is applicable to the semi-active control and active control of the suspension system, and can effectively improve the road friendliness and smoothness of the four-axle heavy truck and have stronger robustness.

A method and associated system for multi-disciplinary optimization of various parameters associated with a space vehicle that experiences aerocapture and atmospheric entry in a specified atmosphere. In one embodiment, simultaneous maximization of a ratio of landed payload to vehicle atmospheric entry mass, maximization of fluid flow distance before flow separation from vehicle, and minimization of heat transfer to the vehicle are performed with respect to vehicle surface geometric parameters, and aerostructure and aerothermal vehicle response for the vehicle moving along a specified trajectory. A Pareto Optimal set of superior performance parameters is identified.

The invention relates to the field of modeling of aerocrafts in the aerospace field, in particular to a pneumatic modeling method which is suitable for a hypersonic aerocraft and based on CFD, a feasible thought is provided for pneumatic modeling with high experiment costs, the analysis and design progresses of the hypersonic aerocraft are promoted, a comprehensive and reliable modeling method is provided for development of a novel hypersonic aerocraft model, the development cost is reduced, and great application prospects and economic value are realized. According to the technical scheme, the CFD pneumatic modeling method for the hypersonic aerocraft is composed of three parts, namely the pre-treatment part of carrying out numerical modeling and mesh division, the solver part for calculation and solution and the after-treatment part for carrying out result analysis, wherein the pre-treatment part includes the two sub-parts of building a geometric model and dividing calculation meshes. The CFD pneumatic modeling method for the hypersonic aerocraft is mainly applied to design and manufacturing of aerospace aerocrafts.

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 for a wheel, relates to the technical field of vehicle engineering and aims at solving the problems that the optimization space is limited and automatic optimization design is not carried out through joint transferring of calculation software with various properties due to the fact that only limited performance of the vehicle is analyzed and optimized and parametric modeling is lost at present. The method comprises the steps of topological optimization of a wheel fatigue test, combined topological optimization modeling, fatigue performance analysis, impact performance analysis, aerodynamic performance analysis and multi-objective optimization. Five working conditions of a dynamic bending fatigue test of the wheel, a dynamic radial fatigue test, an impact test of an impact hammer opposite to a spoke, an impact test of the impact hammer opposite to a window and air dynamics analysis of the wheel are set; the properties, such as the strength, the rigidity, the vibration characteristic, the fatigue life, the fatigue life security coefficient, the impact performance, the air dynamic property and the lightweight performance of the wheel are comprehensively considered; various performance parameters are calculated through finite element software emulation; and optimization design is carried out on the wheel by setting the performance parameters to be constraint conditions or objective functions.

A method of design optimization includes: a) producing a first computer model of a first geometry, the first computer model including: i) a set of nodes; and ii) a set of construction points each having a position defined by a local co-ordinate system; b) producing a computational fluid dynamics (CFD) mesh from the first computer model; c) using the CFD mesh to produce a measure of the fluid dynamic performance of the first geometry, such as drag count or lift coefficient; d) producing a new computer model of a new geometry by: i) moving one or more of the nodes; and ii) recalculating the positions of the construction points in response to movement of the one or more nodes; e) producing a computational fluid dynamics (CFD) mesh from the new computer model; f) using the CFD mesh to produce a measure of the fluid dynamic performance of the new geometry, such as drag count or lift coefficient; and g) identifying an optimal geometry using the fluid dynamic performance measurements of the first and new geometries produced in steps c and f.

The invention discloses a full-firing-direction autonomous reentry guidance method based on a three-dimensional reentry trajectory analytical solution. The full-firing-direction autonomous reentry guidance method comprises the following steps of 1, building of a kinematical equation and a dynamical equation, 2, summarizing of the reentry guidance method, 3, descent stage guidance strategies, 4, steady gliding stage guidance strategies and 5 height-adjustment stage guidance strategies. The full-firing-direction autonomous reentry guidance method has the advantages that reentry guidance of an aircraft with the large lift-drag ratio can be guided effectively, low-damping and long-period trajectory oscillation caused by the high lift-drag ratio can be restrained effectively, the peak value ofthe heat flux density is reduced, and tracking of a reference track is facilitated; inertia force caused by earth rotation and actual aerodynamic force are combined into equivalent aerodynamic force,an equivalent aerodynamic section is planned through the analytical solution, and thus the aircraft has the omni-directional reentry task processing capability; and an equivalent aerodynamic section fitting formula based on the inverse proportional function is put forward, the reentry trajectory analytical solution based on the complex equivalent aerodynamic section form is further deduced, and thus high precision, all-dimensional applicability and high robustness are achieved.

The invention discloses an aeroengine remaining service life prediction method based on an LSTM network and an ARIMA model. The method comprises the following steps: n engine health index (LSTM-HI) models based on an LSTM depth neural network are established according to engine historical degradation data; According to the engine sensor data, the ARIMA model is trained and the engine sensor parameters are predicted backward in many steps. By predicting the sensor parameters, according to LSTM-HI index is used to evaluate whether the engine is degraded to failure, and the remaining service lifeand its probability distribution are obtained. The invention provides a novel method for predicting the remaining service life of an aeroengine, which has high accuracy and feasibility, and plays anactive role in promoting the real-time health management of the aeroengine and reducing the maintenance cost.

The invention discloses a method for designing a water-spraying boost pump hydraulic model for an efficient and large-power-density ship. The method comprises the following steps that type selection design of water-spraying boost pump hydraulic parameters is conducted, two-dimensional axial face projective geometry is determined, three-dimensional geometrical shapes of an impeller and guide blades are obtained according to a parameterized ternary reversal design method, a three-dimensional geometrical shape of a shrinkage nozzle is obtained through circumferential rotation, the three-dimensional geometrical shapes of the impeller, the guide blades and the shrinkage nozzle are combined to obtain the water-spraying boost pump hydraulic model. The method has the outstanding advantages that the design period is short, the design quality is high, and the water-spraying boost pump hydraulic model which has the advantages of being high in efficiency, compact and large in power density can be rapidly and reliably designed. The geometrical shapes of the impeller and the blades are obtained according to the parameterized ternary reversal design method, so that on the condition that the water-spraying boost pump hydraulic model is kept compact, the efficiency is improved, cavitation is restrained, and the power density is increased in the researching and developing process of a high-speed ship water-spraying boost pump.

The invention discloses an optimization method of a smooth and steady gliding trajectory of a hypersonic flight vehicle. The optimization method specifically comprises the steps that 1, a reentry dynamics equation and a trajectory process constraint are established; 2, a trajectory damping control technology is extended; 3, the trajectory damping control technology is utilized to improve the dynamics equation; 4, a hp self-adaptive Radau pseudo-spectral method is utilized to solve an optimal control problem. The optimization method effectively inhibits trajectory jumping by changing the dynamic characteristics of a reentry trajectory and establishes the optimal control problem having smooth and steady state characteristics by changing the dynamic characteristics of the reentry trajectory, the difficulty of numerical solution is greatly reduced, the high-precision smooth and steady gliding trajectory meeting multiple constraint requirements can be obtained within short time, the optimization efficiency of the smooth and steady gliding trajectory is improved, and the regularity of the trajectory optimization is greatly strengthened.

The invention discloses an unmanned helicopter blade load flight actual measurement system which comprises a blade strain acquisition system, an unmanned helicopter flight state measurement system, adata sending and transmitting system, a ground receiving system and a data processing system. The blade strain acquisition system is mounted on an unmanned helicopter rotor hub through a fixture discand synchronously rotates along with a rotor to record and send rotor blade strain parameters. The unmanned helicopter flight state measurement system is mounted in a mission cabin of an unmanned helicopter to acquire, record and send measured flight state parameters of the unmanned helicopter. The ground receiving and data processing system receives the flight state parameters and the rotor bladestrain parameters sent by the data sending and transmitting system and analyzes and processes to obtain blade load and flight state parameters. The unmanned helicopter blade load flight actual measurement system is small in size, light in weight and high in precision.

A computer-aided method suitable for assisting in the design of an aircraft by providing relevant dimensioning values corresponding to an aircraft component in transonic conditions inside a predefined parameter space by means of a reconstruction of the CFD computations for an initial group of points in the parameter space using a POD reduced-order model, comprising the following steps: a) Decomposing for each flow variable the complete flow field into a smooth field and a shock wave field in each of said computations; b) Obtaining the POD modes associated with the smooth field and the shock wave field considering all said computations; c) Obtaining the POD coefficients using a genetic algorithm (GA) that minimizes a fitness function; d) Calculating said dimensioning values for whatever combination of values of said parameters using the reduced-order model. The invention also refers to a system able to perform the method.

The invention provides a hypersonic speed guided missile multi-field coupling dynamics integrated simulation analysis method, which comprises the following steps of: establishing a coupled structural motion equation; improving an aerodynamic heating calculation method; improving an unsteady aerodynamics calculation method; establishing an integrated simulation analysis technique process; designing a relationship between an integrated simulation analysis operation model and data; and realizing an integrated simulation analysis system. Compared with the prior art, the aerodynamic heating calculation method and the unsteady aerodynamics calculation method are improved; various techniques of flutter analysis, thermal flutter analysis, pneumatic servo elastic analysis and thermal pneumatic servo elastic analysis are integrated; and an effective integration method for integrally solving and optimally designing a high-speed winged missile all-missile combination body coupling dynamics problem is provided.

The invention discloses a structural finite-element parametric modeling method applicable to a grating-configuration rudder surface. According to the method, a mapping transformation method based on a finite element model and a parametric approach to achieving conversion from two-dimensional mesh parameterization division to three-dimensional appearance expansion are adopted. The method includes the steps of specifically conducting two-dimensional plane projection on the grating-configuration rudder surface, extracting characteristic parameters, and conducting two-dimensional parameterization division; building a mapping relation between a two-dimensional finite element mesh and a three-dimensional finite element mesh, designing a numbering rule for finite element mesh points, and achieving conversion from the two-dimensional mesh to the three-dimensional appearance expansion; by means of a high-level computer language program, implementing the structural finite-element parametric modeling procedures for the grating-configuration rudder surface. By using the method, in a conceptual design phase or a preliminary design phase, the structural modeling efficiency can be greatly increased, the costs of time and manpower are low, parameters of self-compiled programs can be adjusted quickly, analysis applicability is high, the obtained model is applicable to analytical calculation of structural vibration, structural dynamics and the like, and the method is applicable to grating-configuration rudder surfaces and grating-configuration airfoils.

The invention relates to a rolling bearing fault diagnosis method based on a vibration signal. A CEEMDAN algorithm is adopted to decompose the vibration signal, de-trending fluctuation analysis is carried out on an obtained intrinsic mode function, a scale function value of each IMF component is calculated, and a noise dominant IMF component is selected to carry out de-noising processing; the noise can be better removed, and the distortion degree of the signal is reduced; calculating correlation coefficients and kurtosis values of all orders of IMF components, selecting IMF components with relatively large correlation coefficients and kurtosis values to perform signal reconstruction, performing Hilbert envelope spectrum analysis on reconstructed signals, extracting fault feature frequency, introducing a grey wolf algorithm to optimize initial parameters of multi-scale permutation entropy, performing MPE value calculation on the reconstructed signals, selecting a proper MPE value to construct a rolling bearing fault feature set, and inputting a fault feature vector into the trained support vector machine to carry out rolling bearing fault recognition, so that the entropy discrimination degree is high, the constructed fault feature vector is better, and the recognition rate is higher.

The invention discloses a hypersonic transition prediction method based on a simplified three-equation transition model. On the basis of a gamma-Re theta t transition model, a new empirical transitionrelation is constructed by introducing the localized pressure gradient parameter, and the transport equation of transition momentum thickness Reynolds number Re theta t is successfully eliminated. Onthis basis, the simulation and prediction of hypersonic boundary layer flow transition are realized by the coupled turbulence / transition model compressible correction method. Compared with the gamma-Re theta t transition model, the hypersonic transition prediction method has a simpler empirical transition relation, and the computational complexity decreases. The hypersonic transition prediction method can accurately predict the onset of hypersonic boundary layer transition, the length of transition region, and the heat flux and friction coefficient on the surface of the hypersonic body.

The invention provides a ship track active-disturbance-rejection control method based on the slip form of a multi-modal nonsingular terminal. The method includes the steps of constructing an expectedship heading equation according to navigational parameters of a ship, deducing a discretization nonlinear structural model of a two-order differentiator, processing expected ship heading signals in the expected ship heading equation to obtain expected ship heading differential signals required for controlling the rudder angle by the control law, obtaining the output rudder angle on the basis of the switchable error feedback control law, constructing an expanded state observer in combination with the error feedback control law, tracking the output deviation according to the expanded state observer, estimating the heading angle and inner and outer disturbance errors controlled by the control law, and feeding the heading angle and the inner and outer disturbance errors back into the control law. By improving the expanded state observer, a linear function and a nonlinear function can be switched in real time; the switchable segmented slip form face is designed; through the segmented slip form face, the switchable error feedback control law is designed, and the under-actuated ship track control is well realized.