251 results about "Inertial coordinate system" patented technology

The invention discloses a four-rotor unmanned aerial vehicle model prediction control method based on a deep Q-network (DQN). The method comprises the following steps: firstly, establishing an inertial coordinate system and a body coordinate system, and establishing a kinematic equation and a kinetic equation of the four-rotor unmanned aerial vehicle based on a Newton second law and an Euler-Lagrange equation; according to the method, an unmanned aerial vehicle robust controller based on model predictive control (MPC) is designed, optimal control input is generated through dynamic optimization, and the trajectory tracking performance of the unmanned aerial vehicle is improved. For various disturbances existing in the whole tracking process, a DQN reinforcement learning algorithm and an MPC method are combined to design a flight control system, control rate errors caused by the disturbances are compensated, the stability of attitude control is enhanced, and the tracking precision and the anti-interference capability are improved. And finally, verifying the robust performance of the flight control system through a simulation experiment.

The invention discloses an ultra-low orbit determination method and device for a near-earth satellite, a medium and a product. Comprising the following steps: acquiring a thrust vector of a satellite engine under a satellite body coordinate system under the condition that the thrust of an ion propeller of a satellite is constant; obtaining a first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and converting the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain thrust components of the thrust vector on three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on-off state of the engine and the thrust components on the three coordinate axes to obtain a small thrust perturbation item; and adding a small thrust perturbation item on the basis of the atmospheric resistance perturbation item and the conventional resistance perturbation item to obtain an input parameter, and inputting the input parameter to the orbit mechanical model, so that the orbit mechanical model determines the satellite orbit element based on the atmospheric resistance perturbation item, the conventional perturbation item and the small thrust perturbation item.

The invention discloses an initial orbit determination method for passive angle measurement of a ring fire orbit for Mars sampling return, and the method comprises the steps: building an absolute dynamic model, obtaining an absolute motion state of a spacecraft, and deducing a relative motion state X; constructing a data set by using the measurement quantities H and X of the sensor; a network architecture is designed, input data I are position and speed coordinates of the tracker in a Mars equator inertial coordinate system and angle information converted from H, and output data are X; and finally, adjusting an observation interval dt and times k according to a ring fire observation condition, training a physical information neural network by using limited observation data, defining a total loss function F comprising a loss function FY and a physical law penalty term FD, and optimizing a weight coefficient and bias by minimizing the loss function through an F gradient descent method. According to the method, the neural network is trained offline, a network mapping model from a target line-of-sight angle measurement sequence to a target orbit is constructed, and the network mapping model is solidified on a returner for online use, so that sampler short-arc passive measurement rapid orbit determination is realized.

The invention discloses a gas-solid two-phase flow numerical simulation method for bidirectional coupling of non-spherical particles and a flow field, and belongs to the technical field of biomass fuel utilization, and the method comprises the following steps: constructing a flow field model in Fluent, and correspondingly setting an initial boundary condition of the flow field; setting initialization conditions of non-spherical particles in the UDF; establishing an inertial coordinate system and a particle coordinate system, and coupling the two coordinate systems; and obtaining a translational motion table / rotational motion characterization equation, determining the position and direction of the particles by solving the two equations, correcting flow field parameters by using the positions and direction of the particles, and iteratively calculating particle motion parameters to realize bidirectional coupling of the non-spherical particles and the flow field. According to the numerical simulation method, through Fluent-UDF numerical simulation, the interaction between the non-spherical particles and the flow field can be accurately simulated, the accuracy of gas-solid two-phase flow numerical simulation is guaranteed, understanding of the complex process of direct combustion coupling of biomass particles of a coal-fired boiler is deepened, and theoretical support is provided for optimization of combustor design and the combustion process.

The invention, which relates to the technical field of the radar system, discloses a swash plate machine phase-sweeping radan installation angle adaptive correction method comprising the following steps: acquiring three-axis attitude angle data of an installation reference surface, and synchronously acquiring azimuth angle rotation data of a radar antenna; based on the three-axis attitude angle data of the installation reference surface, establishing a coordinate transformation matrix from a radar antenna coordinate system to an inertial coordinate system, and mapping the three-axis attitude angle data of the installation reference surface to the inertial coordinate system; obtaining a beam scanning conical surface deviation angle according to the coordinate transformation matrix and azimuth angle rotation data of the radar antenna; generating antenna attitude correction according to the beam scanning conical surface deviation angle; and inputting the antenna attitude correction quantity into an antenna driving controller, adjusting the attitude of two axes of pitching and orientation of the antenna, and compensating the beam scanning conical surface deviation. According to the method, the precise coordinate transformation matrix and the angle difference compensation model are established, and the inertial measurement unit and a closed-loop feedback control mechanism are combined, so that the self-adaptive real-time correction of the mounting angle is realized.

The invention relates to a UUV (Unmanned Underwater Vehicle) control method based on a feed-forward PID (Proportion Integration Differentiation). The objective of the invention is to solve the problems of poor precision and poor stability of a traditional PID control algorithm in UUV control. The method comprises the following steps: 1, setting an initial position and a target position of the UUV in an inertial coordinate system; 2, real-time sea wave interference information and the real state of the UUV are obtained; 3, outputting angular acceleration by the PID attitude controller; 4, outputting the optimized depth information by the filter; 5, outputting linear acceleration by the PID position controller; 6, the angular acceleration and the linear acceleration act on an executing mechanism, the thrust and the torque of a propeller are changed by controlling the rotating speed of the executing mechanism, the thrust and the torque act on the UUV, and the underwater movement track of the UUV is changed; 7, judging whether the UUV reaches the target position or not; if yes, the UUV stops sailing; and if not, repeatedly executing the steps 2-6 until the UUV reaches the target position. The method is applied to the field of UUV control.

The invention is suitable for the technical field of multi-unmanned aerial vehicle cooperative control, and provides a multi-quadrotor formation maneuvering flight event triggering cooperative control method and system, and the method comprises the steps: firstly building a kinetic model of a quadrotor unmanned aerial vehicle in a distributed formation system under an inertial coordinate system, designing a multi-quadrotor unmanned aerial vehicle cooperative formation synchronization error, and carrying out the cooperative formation of the quadrotor unmanned aerial vehicle; establishing a preset performance control function to constrain the formation synchronization error, and performing corresponding equivalent error transformation; designing a dynamic event trigger track subsystem controller; designing a robust attitude subsystem controller; and constructing a Lyapunov function, and verifying whether a closed-loop system is stable or not and whether an expected formation effect can be achieved or not. The method can effectively restrain the synchronization error of the multi-quadrotor unmanned aerial vehicle formation system, reduces the error overshoot, improves the coping capability for uncertainty, environmental interference and abrupt trajectory change, and guarantees the stable formation flight of the multi-quadrotor unmanned aerial vehicle system.

The invention discloses a method for establishing a kinetic model of an underwater operation robot, and relates to the field of underwater robotics.The method comprises the steps that firstly, a motion model of the underwater operation robot is established and simulated, a robot model is designed through a three-dimensional tool, and a terrain model of an underwater environment is established in combination with a geographic information system (GIS); meanwhile, an inertial coordinate system and a coordinate system of the robot are further defined according to the established model, the position and the posture of the robot in the space are analyzed, and the position and the posture of the robot in the space are obtained by analyzing and calculating the translational motion coefficient Py and the rotational motion coefficient Xz of the robot and then presetting a posture threshold value Z; and performing comparison evaluation on the attitude change coefficient Jq of the robot, and if the attitude of the robot is found to be abnormal, namely the attitude change coefficient Jq of the robot exceeds a preset attitude threshold value Z, further analyzing the influence of the underwater environment on the attitude of the robot, and performing real-time regulation and control on the power output of the robot according to the underwater environment.

The invention provides a satellite inertial installation error calibration method and system based on iterative filtering, and aims to solve the problem that the attitude measurement precision of a star sensor is influenced by an installation error between inertial navigation and the star sensor. The method comprises the following steps: constructing a state space model and a measurement model of iterative extended Kalman filtering through a star vector under a star sensor body coordinate system, an ephemeris vector under an inertial coordinate system and inertial information, and deducing and calculating to obtain a measurement Jacobian matrix; and substituting the measured Jacobian matrix into the measurement model of the iterative extended Kalman filter to iteratively estimate the satellite inertial installation error, namely the installation error of the inertial astronomical integrated navigation system, so that the function of calibrating the installation error of the inertial astronomical integrated navigation system is realized. The method eliminates the influence of the installation error between the inertial navigation system and the star sensor on the attitude measurement of the star sensor, is suitable for the application occasion of combined navigation, and is good in adaptability and high in precision.

The invention relates to a method for determining an initial orbit of a satellite with an embedded physical knowledge neural network, which is used for determining an initial orbit of a short arc segment under an optical observation condition and providing support for subsequent precise orbit determination. The network comprises an encoder-decoder structure, an encoder predicts the motion state of a satellite in an orbit according to observation information, the decoder converts the predicted satellite motion state into an observation value, the prediction accuracy of the encoder is evaluated, physical information of satellite orbit dynamics is introduced into a loss function, and the prediction accuracy of the satellite orbit is improved. And satellite orbit determination based on the neural network is realized. The input of the encoder is the observation angle quantity (namely azimuth angle and pitch angle) of the satellite relative to an observer, the observation time and the position information (namely longitude, latitude and altitude) of the observer; the output of the encoder, namely the input of the decoder, is the absolute motion state quantity (namely the position and the speed of the satellite in a geocentric inertial coordinate system) of the satellite in the orbit and the position information of an observer, and the output of the decoder is the observed quantity of the satellite relative to the observer. When the network is trained, firstly, a decoder is trained by taking a mean square error between a predicted observation value and a real observation value as a loss function, and a transformation relation from a satellite motion state to the observation value is learned; fixing the structure and parameters of a decoder, training the whole encoder-decoder network, calculating the derivative of the predicted satellite absolute state quantity relative to the observation time, and calculating the physical information loss of a speed item and an acceleration item; adding the physical information item loss and the mean square error of the observed quantity, constructing a total loss function, and optimizing network parameters through error back propagation; in the reasoning stage, an encoder network is used for reasoning, and the absolute motion state quantity, predicted by an encoder, of the satellite is converted into an orbit element form, namely, the initial orbit is determined.

The invention discloses an unknown environment-oriented four-rotor unmanned aerial vehicle autonomous obstacle avoidance system, which comprises an autonomous obstacle avoidance module and a flight control unit, the autonomous obstacle avoidance module is connected with the flight control unit, the flight control unit is also connected with a remote control unit, and the autonomous obstacle avoidance module comprises a trajectory planning unit and an environment sensing unit. The trajectory planning unit obtains output information of the environment sensing unit to generate a speed instruction to the flight control unit; the trajectory planning unit is used for constructing an inertial coordinate system and an airframe coordinate system to obtain the position and Euler angle of the unmanned aerial vehicle relative to the inertial coordinate system, performing path planning based on a potential field method, and outputting a position control instruction. Compared with the prior art, the four-rotor unmanned aerial vehicle autonomous obstacle avoidance system for the unknown environment has the advantages that the four-rotor unmanned aerial vehicle autonomous obstacle avoidance system is based on the potential field method obstacle avoidance principle, and path planning and tracking control are integrated.

The invention relates to an indoor airship-oriented finite time trajectory tracking control system, platform and method, the system comprises an airship dynamic model, a disturbance estimation module, a sliding mode filter and a self-adaptive controller, the airship dynamic model is formed by constructing a motion stipulation inertial coordinate system and an airship body coordinate system of an airship; the disturbance estimation module estimates unknown disturbance based on a double-layer neural network according to an airship dynamic model and updates the network in real time; the sliding mode filter estimates the influence of noise on the motion speed of the airship according to the unknown disturbance condition of the airship and updates the network in real time; the adaptive controller adjusts the virtual control input of the airship motion trail by controlling the convergence condition of the tracking error within the finite time; according to the method, the disturbance problem caused by uncertainty and dead zones in the airship and the deviation problem of speed signal estimation caused by noise are solved.

The invention discloses a two-stage installation error outfield calibration method for a space target fine-following camera and a two-dimensional rotary table, and belongs to the technical field of error calibration. The method comprises the following steps: extracting shooting information of each angle in an inertial coordinate system according to a plurality of starry sky images at different shooting angles; according to the inertial sight vector truth value of the equator coordinate under the inertial coordinate system, calculating a horizon sight vector truth value under the east-north-sky coordinate system by combining a transfer matrix from the east-north-sky coordinate system to the inertial coordinate system; calculating a horizon sight vector measurement value of the corresponding angle according to the rotary table rotation angle of each angle and the equipment installation matrix; and inputting the errors of the truth values and the measured values of the horizon sight line vectors under all the shooting angles into a least square iterative algorithm for optimization calibration, and outputting an optimal rotary table installation matrix and an optimal camera installation matrix. Calibration of the fine-following camera of the ground test two-dimensional rotary table can be realized, and the pointing precision of the fine-following camera is improved.

The invention relates to a satellite-borne microwave radiometer cold air calibration attitude planning method based on multi-objective optimization, and belongs to the field of satellite microwave remote sensing, and the method comprises the steps: building a satellite attitude kinematics model and an attitude dynamics model through a geocentric equator inertial coordinate system, an orbital coordinate system and a satellite body coordinate system; establishing a brightness temperature sum model received by the microwave radiometer in the cold air calibration process; according to the method, a microwave radiometer receiving brightness temperature sum, brightness temperature fluctuation, platform maneuvering capability, energy and a satellite sensor field of view are comprehensively considered, a multi-target optimization model of microwave radiometer cold air calibration consideration constraints is established, solving is performed, constraint conditions are established, and a solution or a solution set is obtained. According to the method, the brightness temperature sum and the brightness temperature fluctuation in the observation area of the microwave radiometer in the cold air calibration process are optimized by adopting multi-objective optimization, and the on-orbit cold air calibration effect and the system reliability of the satellite-borne microwave radiometer are improved.

The invention provides a non-singular sliding mode control method, device and equipment for a quad-rotor unmanned aerial vehicle, and a medium, and relates to the technical field of unmanned aerial vehicle flight control. According to the method, a quadrotor unmanned aerial vehicle body coordinate system and an inertial coordinate system are established, and an attitude model of the quadrotor unmanned aerial vehicle is established by considering air resistance and external unknown disturbance conditions according to a rigid body motion law; introducing an auxiliary equation to construct a predefined time disturbance observer, and rapidly estimating external disturbance by adjusting predefined time in combination with a Lyapunov function; then designing a non-singular predefined time sliding mode controller; and controlling the quadrotor unmanned aerial vehicle to perform attitude tracking according to the nonsingular predefined time sliding mode controller, and controlling the attitude tracking error of the quadrotor unmanned aerial vehicle to converge within the predefined time by adjusting predefined time parameters. According to the method, the external disturbance can be quickly estimated within the preset time, and the convergence rate of the four-rotor unmanned aerial vehicle is flexibly controlled.

The invention discloses an inertial navigation system periodic oscillation suppression method based on time-frequency transformation, and relates to the technical field of underwater navigation, and the method comprises the following steps: S1, modeling an error propagation model of an INS, and defining coordinate systems required by INS navigation, including an inertial coordinate system, a navigation coordinate system and a body coordinate system; based on the INS navigation equation, obtaining an INS error model, and carrying out time domain analysis to obtain an analytical solution of time domain error divergence; s2, designing a time-frequency transformation NTFT, proposing a normal phase time-frequency transformation method, and constructing a complete observation equation; s3, substituting a null method into a calculated longitude and latitude result, and accurately extracting frequency, phase and instantaneous amplitude in inertial navigation longitude and latitude; and S4, establishing a resolving framework of the inertial navigation system periodic oscillation suppression method based on time-frequency transformation, and constructing a negative feedback loop through input gyroscope and acceleration information to suppress periodic errors.

The invention discloses a manned lunar exploration-oriented earth-moon emergency return orbit rapid design method, which comprises the following steps of: establishing a high-precision orbit kinetic equation in a geocentric inertial coordinate system by taking the earth as a central celestial body, and describing the motion state of an aircraft in an earth-moon space; determining a general transfer orbit and a free return orbit of the manned lunar exploration task by adopting a B-plane design method in combination with a high-precision orbit kinetic equation of the aircraft; globally searching an initial value range of a design variable under the condition of comprehensively considering reentry point height and reentry angle constraints by taking a general transfer orbit or a free return orbit as a baseline; and on the basis of the initial value range of the design variable, further considering the reentry range, the reentry inclination angle and the coplanar constraint with the drop point, and determining a final design variable and an emergency return orbit by adopting a differential correction optimization algorithm. The method has an important reference value for rapid design and parameter selection of the emergency return orbit in a manned lunar exploration task.

The invention provides an inter-satellite link sun-sun avoidance method for a giant satellite base satellite network. The method comprises the following steps: initializing the giant satellite base satellite network; calculating the date of the inter-satellite link influenced by the sun; calculating the included angle between each inter-satellite link and the sun position vector in the date influenced by the sun; and avoiding the sun. According to the method, on the basis of the uniform distribution characteristic of a Walker constellation, the dates of the same-orbit link and the different-orbit link influenced by the sun-sun are calculated under a J2000 geocentric inertial coordinate system, the dates of the inter-satellite links influenced by the sun-sun are calculated according to the motion law of the sun and the characteristics of the different types of inter-satellite links, and the data of the inter-satellite links influenced by the sun-sun are calculated within the dates influenced by the sun-sun. According to the relative position relation of the satellites, the included angle between the inter-satellite link and the sun is calculated in real time, whether the inter-satellite link is affected by the sun or not is judged, the optical axis direction of the laser terminal does not need to be obtained in real time, the included angle between the inter-satellite link and the sun does not need to be calculated in real time on the date without the sun influence, and invalid calculation expenses are reduced.

The invention provides a centimeter-level broadcast ephemeris method and device based on a GNSS precise ephemeris, and the method comprises the steps: converting a classical Kepler orbit element into a nonsingular orbit element for the near-zero eccentricity rate of a middle earth orbit satellite; performing fast Fourier transform on a satellite orbit, selecting a certain number of additional parameters from the optional additional parameter set according to a spectral analysis result of the Fourier transform, and combining the additional parameters with an original broadcast ephemeris model to construct a centimeter-level broadcast ephemeris model; interpolating the precise ephemeris to obtain dense ephemeris data points, and determining centimeter-level broadcast ephemeris model parameters; calculating an error in a geocentric inertial coordinate system, converting the error into an orbital coordinate system, and determining the precision of a centimeter-level broadcast ephemeris model by using a space signal ranging error; and broadcasting the centimeter-level broadcast ephemeris model for satellite orbit determination and user high-precision positioning. The method provides support for real-time high-precision satellite positioning application.

The invention belongs to the technical field of satellite communication, and discloses a satellite communication beam switching method. The positions of a satellite and an aircraft are determined under a geocentric inertial coordinate system, an elevation angle is calculated, and visible wave beams are screened to form a neighbor cell list. And screening out a candidate beam list from the neighbor cell list based on the received power threshold. And if the list is not empty, constructing a decision matrix containing the receiving power, the signal-to-noise ratio, the number of available channels and the satellite residual service time for each candidate beam. And objectively calculating the weight of each performance index based on a matrix by using a CRITIC method. And by using the weight, a TOPSIS model based on Tanimoto similarity optimization is adopted to calculate the relative close degree of each candidate wave beam. And finally, selecting the candidate beam with the maximum relative closeness as the optimal beam to trigger switching. According to the method, through the objective weight and the optimization model, unnecessary frequent switching is reduced, and the communication stability and the system resource utilization rate are improved.

The invention discloses an orchard crawler differential robot trajectory tracking method based on adaptive incremental control, and relates to the technical field of agricultural robot control. The method comprises the following steps: receiving environmental data collected by a robot sensor, constructing a probability grid map of an orchard environment according to the environmental data, and aligning the probability grid map with a pre-constructed map inertial coordinate system; the course angle error and the distance error between the current position and a target point are calculated based on the pose of the robot, an error vector is formed, and the target point is dynamically planned according to the fruit maturity and reachability priority detected by a visual system and a laser radar point cloud. According to the method, coordinate transformation is not needed, and different from traditional tracking control, a control instruction is generated directly based on pose increment, the course and distance error of a target point is calculated directly based on an inertial coordinate system, and real-time matrix projection operation is avoided.

The invention relates to a flight path tracking control method for an unmanned autonomous submersible vehicle, and belongs to the technical field of unmanned submersible vehicle control. The method comprises the following steps: firstly, establishing a nonlinear dynamic model of the submersible vehicle, analyzing kinematics and dynamic characteristics of the nonlinear dynamic model, considering various underwater forces and moments, and defining an inertial coordinate system and a submersible vehicle body coordinate system to obtain a kinematics equation and a dynamic equation so as to provide a basis for control design; secondly, a non-singular integral terminal sliding mode control method is adopted, a sliding mode surface is designed, the singular problem of traditional sliding mode control is avoided, a sliding mode control law is deduced, system finite time convergence to a balance point is achieved, and an expected state is rapidly tracked. And finally, a brand new and simple adaptive control strategy is provided, an adaptive law is designed according to system errors and sliding mode surface information, a gain matrix is adjusted in real time, and system uncertainty and external interference are compensated. The method has the advantages of being high in robustness, rapid in convergence, simple and easy to implement.

The invention relates to a low-orbit star network-oriented random target tracking estimation method, which comprises the following steps of: establishing a dynamic filter AUKF of multiple targets in a geocentric inertial coordinate system by taking position, speed and thrust acceleration vectors of the targets j (j = 1, 2,..., Nt) in the geocentric inertial coordinate system as state vectors; establishing a maneuvering target space-based observation model based on the measurement equation of the target under the geocentric inertial coordinate system; based on the calculation result of the maneuvering target space-based observation model at the current moment, constructing a maneuvering starting and ending detector, when it is judged that the target starts maneuvering, using a maneuvering dynamics filter AUKF to track and estimate the position, speed and maneuvering acceleration information of the target, and when it is judged that the maneuvering of the target is ended, using the maneuvering dynamics filter AUKF to track and estimate the maneuvering speed of the target. And a standard Kalman filter UKF is switched to carry out state updating on the position and the speed of the target. According to the method, the relay observation and whole-course tracking performance estimation problems of the low-orbit star network on the uncertain maneuvering target are solved, and the method can be used for verification and evaluation of a large-scale star network system detection simulation test.

The invention relates to an ephemeris orbit calculation method suitable for orbit extrapolation on an elliptical orbit satellite. The ephemeris orbit calculation method comprises the following steps: calculating ephemeris orbit base point parameters on the ground; ephemeris orbit base point parameters which are injected from the ground are provided; and performing orbit extrapolation by using the ephemeris orbit base point parameters to calculate the position and the speed of the satellite in the inertial coordinate system. According to the method, orbit extrapolation is carried out on the basis of ephemeris orbit base point parameters injected on the ground, the calculation amount is low, the method is suitable for satellite-borne calculation, the calculation precision is high, the high-precision requirement of real-time orbit forecasting in the elliptic orbit satellite inter-satellite communication period can be met, the calculation capacity constraint of a satellite service system can be met, and the completion degree of an inter-satellite communication task is improved.

The invention provides a correlation method and system for space debris observation data and cataloguing orbit data. The method comprises the following steps: firstly, within a time range of currently processed batch observation data, recurring to an orbit matrix distributed at equal time intervals through an orbit of an orbit recursion cataloguing target, and generating a position interpolation function based on a time sequence and the orbit matrix; and for observation data in a processing time range, firstly unifying the observation data and an observation station position to a J2000.0 inertial coordinate system, and obtaining a multi-target position prediction value corresponding to observation time through an interpolation function. Calculating the difference between the predicted value and the measured value of the observation data; and calculating an associated object and confidence according to the difference. The method is easy to implement, high in efficiency during batch observation data processing and suitable for association of observation data of different observation devices and different types of orbit space debris and catalogued orbit data.

The invention provides a satellite flexible attitude dynamics modeling method considering the change of the mass center of a whole satellite, and the model is used for simulating the on-orbit flexible attitude behavior of the satellite. The method comprises the following steps: establishing an inertial coordinate system Oi, an orbital coordinate system Oo, a star coordinate system Ob and a sailboard coordinate system Of; establishing an attitude motion equation Eq.1 of a whole satellite system, an equation Eq.2 of flexible vibration of the sailboard and an attitude motion equation Eq.3 of the sailboard relative to a central body, wherein the Eq.1, Eq.2 and Eq.3 consider the influence of the mass center change of the whole satellite and the flexible vibration of the large sailboard; and calculating a direction cosine matrix RoI of the centroid orbital coordinate system, a direction cosine matrix RsI of the star coordinate system, an attitude conversion matrix Rso of the star coordinate system relative to the centroid orbital coordinate system and an Euler angle, and taking the Euler angle as attitude control input. According to the satellite attitude dynamics calculation method considering the satellite centroid change and the sailboard elastic deformation in real time, the problem that the instantaneous centroid change cannot be simulated in an existing attitude dynamics algorithm is solved.

The invention relates to a GNSS (Global Navigation Satellite System) auxiliary measurement-based autonomous navigation method for a lunar satellite formation release process, which comprises the following steps of: establishing a kinetic model of a primary satellite and a secondary satellite under a lunar center inertial coordinate system J2000; establishing and fusing an inter-satellite measurement model and a distance measurement model of a main satellite fused GNSS signal to obtain a collaborative observation model; and acquiring orbit prediction information based on the kinetic model, acquiring collaborative theory observation information based on the collaborative observation model, performing fusion processing on the orbit prediction information and the collaborative theory observation information by adopting an extended Kalman filtering algorithm, and estimating orbit states of the primary satellite and the secondary satellites. According to the autonomous navigation method, full-autonomous orbit determination can be realized in orbit without depending on the support of a ground station.

The invention relates to an error self-calibration and compensation method for a liquid floated gyroscope platform type inertial navigation system, which comprises the following steps of: locking a platform body at 13 frame angle positions under the condition of a static base, controlling the platform body to track an inertial coordinate system through a stable loop, and collecting measured values of an accelerometer and a frame angle sensor. And calculating an attitude angle of the table body under an earth coordinate system by using a table body attitude calculation method of an inertial navigation system, and comparing the attitude angle with a table body attitude angle measured by a platform frame angle sensor to obtain an attitude misalignment angle observed quantity. And performing linear fitting on the state misalignment angle sequence of each section of the process of tracking the inertial coordinate system to obtain the comprehensive angular velocity error of the inertial navigation system at the respective calibration position. The total least square algorithm is adopted to solve residual errors of drift error coefficients of all zero-order terms, first-order terms and cross terms, estimated values of measured parameters are corrected, the corrected error coefficient estimated values are used for iterative calculation until an iterative convergence criterion is met, and self-calibration results of all the drift error coefficients are obtained.

The invention provides a state estimation method and device, electronic equipment and a storage medium. According to the specific implementation scheme, the method comprises the following steps: acquiring a state quantity, first data acquired by a first inertial measurement unit arranged on first equipment, and second data acquired by a second inertial measurement unit arranged on second equipment; the state quantity comprises a target parameter, a relative position and a relative attitude between the second equipment and the first equipment, an angular velocity bias and an acceleration bias of the first inertial measurement unit, and an angular velocity bias and an acceleration bias of the second inertial measurement unit; the target parameter comprises at least one of the following two: the relative speed between the second equipment and the first equipment, and a projection result obtained by projecting the relative speed between the second equipment and the first equipment under an inertial coordinate system to a reference coordinate system; obtaining a state estimation equation meeting a predetermined lever arm constraint condition; an estimated state quantity is determined based on the state quantity, the first data, the second data, and a state estimation equation.

The invention relates to an error-free strapdown inertial navigation method based on a launching inertial coordinate system, and relates to the technical field of strapdown inertial navigation. According to the error-free strapdown inertial navigation method based on the launching inertial coordinate system, under the condition that the output angular velocity of a gyroscope and the output specific force of an accelerometer meet a polynomial hypothesis, smooth angular motion or linear motion can always be infinitely approximated by a polynomial according to mathematical knowledge, and the linear motion can be infinitely approximated by the polynomial according to a navigation equation of strapdown inertial navigation under the launching inertial coordinate system. And a strapdown inertial navigation value updating algorithm for the attitude, the speed and the position under the engine inertial system is exported by adopting a Taylor series expansion method. The algorithm does not have any principle error and comprises perfect compensation for non-exchangeable errors of attitude, speed and position updating, and the algorithm precision is far higher than that of a traditional two-subsample algorithm.