105 results about "Inertial coordinate system" patented technology

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 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 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 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 application provides a method for analyzing a pointing angle of an inter-satellite link, comprising: obtaining orbital elements of a satellite and orbital parameters of a target satellite matched with an SGP4 model in a J2000.0 inertial coordinate system; determining a target recursion time t; obtaining positions of the satellite and the target satellite at the time t in the J2000.0 inertial coordinate system; determining a pointing vector from the satellite to the target satellite at the time t in the J2000.0 inertial coordinate system; determining the pointing vector in a satellite orbital coordinate system; determining the pointing vector in a satellite body coordinate system; determining the pointing vector in a satellite antenna coordinate system; determining a pointing angle of the satellite antenna coordinate system at the target recursion time; and determining whether a cycle number reaches a set cycle number, if the cycle number does not reach the set cycle number, continuously obtaining the pointing angle of the satellite antenna coordinate system at a next target recursion time until the pointing angle of the satellite antenna coordinate system at the set cycle number is obtained.

The invention belongs to the technical field of electromagnetic field calculation, and discloses a method and a system for calculating an electromagnetic scattering field of a moving target coated with a lossy dispersive medium. The method comprises the following steps: firstly, carrying out three-dimensional modeling on a target surface by adopting triangular surface elements to form a model file, and importing the model file into a main program to obtain geometric feature data of a target; secondly, carrying out relativistic coordinate transformation on the space-time component by utilizing Lorentz transformation, and mapping a dynamic problem to an inertial coordinate system synchronous with a target; then establishing an equivalent reflectivity model for the lossy dispersive medium coated area by adopting a transmission matrix method; then efficiently solving a transient scattering field in the motion coordinate system by adopting a time domain physical optical method; and finally, reconstructing a calculation result to a laboratory coordinate system through inverse Lorentz transformation, and outputting electromagnetic field distribution in a dynamic scene. Through the method provided by the invention, the electromagnetic scattering field of the moving target coated with the lossy dispersive medium can be quickly calculated.

The multi-detection-view-angle target matching method based on the error direction consistency comprises the steps that multiple targets are detected at different angles through multiple aircrafts, and images of the multiple targets at different detection view angles are obtained; detecting each image by adopting a target recognition algorithm, and calculating a sight angle of a target in a camera coordinate system; acquiring a sight angle under an inertial coordinate system, and acquiring position and attitude angle information of multiple aircrafts; calculating direction vectors of the target under different detection view angles and direction vectors among the multiple aircrafts; constructing a target matching model under multiple detection view angles by using the characteristic that the direction vector of the same target under different detection view angles and the corresponding aircraft are in the same plane and have a unique intersection point; according to the characteristic that the error direction of the current multi-target data frame has consistency, target rough matching under multiple detection view angles is completed, and a matching threshold value is calculated; and performing differential matching with the multi-target rough matching value under each detection view angle based on the matching threshold, and coding the successfully matched target.

The application discloses a kind of based on liquid rocket's interstage separation gap calculation method and system, including establishing the inertial coordinate system of liquid rocket launch, rocket body coordinate system and the coordinate system of storage tank;Obtain the dynamics equation of liquid rocket mass center motion in inertial coordinate system and the dynamics equation of liquid rocket rotation around mass center in rocket body coordinate system;Obtain the displacement H of liquid rocket core first stage and liquid rocket core second stage and attitude angle;Obtain the displacement L1 of liquid rocket core second stage mass center deviating from nozzle center line, the displacement L2 of liquid rocket core second stage mass center deviating from liquid rocket core first stage axis and the distance L3 of nozzle outer edge to nozzle center line;Obtain interstage separation gap as L4, L4=R-L3-(L1-L2), solve the problem that core second stage engine nozzle and interstage section sleeve may collide in the process of large-diameter liquid rocket separation, to improve the precision of separation scheme design, attitude control system design and storage tank anti-slosh structure design.

The invention relates to the field of satellite orbit calculation, in particular to a low-orbit satellite ephemeris analysis method based on a satellite orbit plane coordinate system, which comprises the following steps: acquiring TLE two-line orbit data of a target satellite, geodetic coordinates of a terminal, UTC time and earth rotation average angular velocity; calculating the state of the satellite in the orbital plane; calculating a transformation matrix from a geocentric inertial coordinate system ECI to a satellite orbit plane coordinate system; the LBH coordinates of the terminal are converted into ECEF rectangular coordinates; a rotation matrix from the ECEF to the ECI is constructed, and the ECI system speed of the terminal is calculated; according to the transformation matrix, transforming the ECI coordinate of the terminal into a final satellite orbit plane coordinate system; performing simplified calculation based on a unified calculation framework of a satellite orbit plane coordinate system; and outputting the distance, the relative speed, the Doppler frequency offset and the propagation time delay result which are calculated under the satellite orbit plane coordinate system. And the calculation complexity and the realization difficulty of high-precision orbit forecasting are obviously reduced.

The embodiment of the invention provides a three-dimensional maneuvering target interception method, a storage medium and equipment. The method comprises the following steps: determining a constraint relation between the speed and the maximum field angle of an interception aircraft and the speed and the maximum field angle of an intercepted aircraft; calculating a collision angle control gain; obtaining a relative speed expected value in the inertial coordinate system according to the expected collision angle in the speed coordinate system; constructing a sliding mode vector as an auxiliary variable according to the relative speed expected value, the sight line direction unit vector and the sight line conversion rate information; setting a gain value of auxiliary variable control; a guidance instruction to intercept the aircraft is calculated. The embodiment of the invention provides a three-dimensional space guidance law design method considering collision angle and field angle constraints at the same time, the problem of three-dimensional space interception under multiple complex constraint conditions is solved, and an air defense and anti-missile combat system is assisted to obtain higher combat performance.

A relative navigation sensor field-of-view obstacle identification and avoidance method comprises the following steps: firstly, setting initial attitude control parameters of a tracking aircraft, and calculating a unit vector of a relative navigation sensor field-of-view optical axis in an inertial coordinate system; then, respectively calculating the included angles between the unit vector relative to the visual field optical axis of the navigation sensor and the unit vectors of the earth, the moon and the non-target aircraft; and finally, sequentially judging whether interference elimination operation of the earth, the moon and the non-target aircraft needs to be carried out or not. Through the above steps, the method effectively eliminates the influence of the earth, the moon and the non-target aircraft on the target recognition of the relative navigation sensor, and improves the precision of the target recognition of the relative navigation sensor.

The present application relates to a kind of inertial measurement device platform body axis unbalance moment estimation and offset method, belong to inertial measurement technical field;Inertial coordinate system OXYZ is established based on inertial measurement device;Establish the control model of platform body axis stable loop in inertial measurement device under the influence of unbalance moment;The model of unbalance moment acting on platform body axis is established;According to unbalance moment model, establish unbalance moment parameter observer, estimate the unbalance moment parameter that platform body axis is subjected to by unbalance moment parameter observer, obtain unbalance moment parameter estimated value;According to unbalance moment parameter estimated value, calculate platform body axis compound control moment;Through platform body axis compound control moment, improve the control precision of platform body axis stable loop;The present application has the characteristics of low cost, high precision, strong engineering practicability, is applicable to the high-precision control of platform system platform body axis stable loop under the influence of unbalance moment.

The invention discloses a multi-quad-rotor unmanned aerial vehicle formation collision avoidance fault-tolerant control method, and relates to the technical field of unmanned aerial vehicle control, and the method comprises the steps: constructing a kinetic model and a dual-state constraint model of a multi-quad-rotor unmanned aerial vehicle, and building an inertial coordinate system and a vehicle body coordinate system; according to the invention, the dynamic model and the dual-state constraint model are constructed, unified modeling processing is carried out on the motion behavior and the space safety boundary of the multi-quad-rotor unmanned aerial vehicle, and real-time acquisition and dynamic compensation are carried out on the state and fault information of the unmanned aerial vehicle in combination with an online estimation mechanism of a distributed fault diagnosis observer. The system can still maintain stable operation under the conditions of actuator faults and coupling disturbance; the position error and the formation error of the unmanned aerial vehicle are constrained, the mapping from virtual control to actual control is realized in combination with an attitude coupling compensation mechanism, the collision risk in the formation flight process is effectively avoided, and the trajectory tracking precision is improved.

The invention discloses a step-by-step marginalized point-line visual inertia SLAM (Simultaneous Localization and Mapping) method and system. Firstly, an environment image and IMU data are obtained and preprocessed; point-line features in the image are extracted, line features with the length larger than the minimum line segment length threshold value in all layers of an image pyramid are reserved, and the direction included angle and the center point Euclidean distance of any two reserved line features are calculated; if the direction included angle of the two line features is smaller than or equal to an angle threshold value and the center point Euclidean distance is smaller than or equal to a distance threshold value, combining the two line features into a new line feature; then, aligning the visual inertial coordinate system, and estimating an initial state; and finally, performing local visual inertia optimization and mapping by using a sliding window based on a step-by-step marginalization strategy to obtain a prior error. And meanwhile, loopback detection is carried out, and loopback errors and prior errors are added into the next round of sliding window optimization to carry out tight coupling state estimation. The method has relatively high robustness and consistency, and meanwhile, the real-time performance is ensured.

This invention discloses a method for precise airdrop and rapid in-flight alignment based on inertial navigation / satellite integrated navigation. The method includes: after the navigation device is powered on, establishing a carrier and navigation inertial coordinate system through inertial solidification; periodically acquiring angular rate and specific force data from the inertial measurement unit and position and velocity data from the satellite receiver; sequentially calculating carrier system tracking, specific force integration, navigation system tracking, and acceleration integration; and recursively solving the problem using a Rodrigues parameter multi-vector attitude determination algorithm to calculate the real-time attitude matrix, repeating this process until alignment is complete. This invention simplifies the initial alignment problem without initial values ​​for arbitrary attitudes into an unconstrained linear optimal estimation problem for Rodrigues parameters, utilizing specific force and acceleration integration as measurement information, and fully utilizing all effective data during the alignment process. It requires no initial attitude information and can rapidly complete alignment under complex dynamic base conditions such as airdrop swaying and rotation, outputting attitude results in real time, significantly improving alignment accuracy and speed.

A virtual object display device comprises a display and a display control device configured to perform display control of the display. The display control device includes: a coordinate system calculation unit configured to detect movement and rotation of the virtual object display device in a real world, and use an inertial coordinate system in which a coordinate origin thereof follows the movement of the virtual object display device and an effective field of view of the display is rotated therein in accordance with the rotation of the virtual object display device so as to define an arrangement position of an inertial coordinate system virtual object; and a display control unit configured to, when the inertial coordinate system virtual object is included in the effective field of view of the display, display the inertial coordinate system virtual object within the effective field of view.

The application discloses a spacecraft single-pulse absolute reachable set solving method based on differential algebra technology and belongs to the technical field of on-orbit motion of space spacecrafts; firstly, a geocentric inertial coordinate system, a geocentric geodetic rotating coordinate system and a cylindrical coordinate system are established, and a dynamics model of on-orbit motion of the spacecraft in the cylindrical coordinate system is established; then, the earth shape perturbation, the solar pressure perturbation and the third body gravity perturbation of the sun and the moon are converted into the cylindrical coordinate system for representation, and a dynamics model considering multiple perturbation forces is obtained; finally, the absolute reachable set is solved in the dynamics model considering multiple perturbation forces based on Jet Transport technology; the dynamics model considering multiple perturbation forces is established in the cylindrical coordinate system, the absolute reachable set is solved by using Jet Transport technology based on the dynamics equation, and the absolute reachable set is simulated by using C++ and MATLAB; compared with the traditional Monte Carlo method, the solving method based on Jet Transport technology is high in accuracy and calculation efficiency.

The invention relates to the technical field of integrated navigation, in particular to a positioning method and system based on optical angle measurement and astronomical-inertial integrated navigation. According to the method, firstly, a star image and a navigational satellite image are obtained at the same time by shooting a star map, after recognition and processing, the attitude of a camera relative to a geocentric inertial coordinate system is calculated by utilizing the star, and then the coordinate of the navigational satellite in the coordinate system is determined. And obtaining angle information of the navigation satellite relative to the observer under the earth coordinate system through coordinate transformation. Optical angle measurement information is fused with attitude, speed and position information output by an inertial navigation system, a Kalman filtering state equation and an observation equation are constructed, and finally real-time high-precision calculation and updating of the position of the observer are achieved. According to the method, the accuracy of optical angle measurement, the limited distance reference of celestial navigation and the autonomous continuity of inertial navigation are fused, and the navigation positioning accuracy and reliability under the complex environment and the satellite denial condition are effectively improved.

A low-orbit remote sensing satellite three-axis rapid maneuvering attitude trajectory planning and tracking method relates to the technical field of aerospace, in particular to the technical field of low-orbit remote sensing satellite three-axis rapid maneuvering attitude trajectory planning and tracking, which effectively solves the balance problem of maneuverability and stability in the process of attitude transformation and improves the superiority of anti-interference capability. The method comprises the following steps: establishing a satellite kinematics model and a satellite dynamics model; calculating the quaternion of the desired coordinate system relative to the satellite initial coordinate system at the maneuvering starting moment; completing attitude trajectory planning through an improved tracking differentiator; calculating the total control rate; obtaining the real control rate through an actuator; calculating the quaternion of the satellite body coordinate system relative to the inertial coordinate system and the angular velocity of the satellite body coordinate system relative to the inertial coordinate system; obtaining the real body quaternion and the real body angular velocity through a measuring mechanism; and cyclically planning the trajectory and tracking.

The invention discloses an inter-satellite link fault avoidance method and system based on time quantitative calculation under the sun-sun effect, and mainly solves the problems that the calculation complexity of inter-satellite link faults of a satellite network is high and the link fault duration cannot be directly calculated under the influence of the sun-sun effect in the prior art. According to the implementation scheme, the method comprises the following steps: initializing constellation parameters, establishing a J2000.0 geocentric inertial coordinate system, and determining a satellite position at any moment; the right ascension and declination of the sun are calculated, the included angle between the sunlight and each satellite orbit plane is calculated according to the parameters, the link fault position existing in the orbit is determined, and the link fault range of each orbit is calculated; calculating positions and fault duration of all fault links in each orbit according to the link fault range and the positions of satellites; and the satellites autonomously perform link avoidance according to the link fault range and the fault duration. According to the method, the calculated amount can be effectively reduced, the response speed is improved, the duration of the link fault is directly obtained, and the method can be used for analyzing the inter-satellite link fault condition of the satellite network under the sun-sun effect.

The application discloses a satellite solar cell array rotation angle estimation method, which comprises the following steps: step 1, establishing an earth inertial coordinate system, a satellite body coordinate system and a solar cell array coordinate system, and defining a satellite solar cell array rotation angle; step 2, establishing a state equation of an augmented state composed of a sun vector, the solar cell array rotation angle and a gyro constant drift noise; step 3, calculating an augmented state quantity estimation value by using an extended Kalman filter; and step 4, calculating a satellite solar cell array rotation angle estimation value. The method fuses gyro angular velocity and sun sensor vector data, constructs a nonlinear dynamic model containing a sun vector, a rotation angle and a gyro bias, and designs an EKF algorithm to estimate the solar cell array rotation angle in real time, thereby solving the strong dependence of a traditional scheme on a star sensor and a mechanical sensor, and significantly improving the autonomous recovery capability and energy guarantee reliability of a satellite in an extreme fault scene.

The application discloses an interception guidance method and related device considering speed time variation and target maneuvering, relates to the technical field of missile guidance, and comprises the following steps: acquiring missile maneuvering parameters and target maneuvering parameters; establishing a three-dimensional guidance model under the condition of speed time variation and target maneuvering according to the missile maneuvering parameters and the target maneuvering parameters; constructing a time-varying navigation ratio proportional guidance law based on the three-dimensional guidance model; increasing a vertical acceleration component on the basis of the time-varying navigation ratio proportional guidance law to generate an improved augmented proportional guidance law; decomposing a missile normal acceleration vector in a missile total acceleration vector to a pitch plane and a yaw plane, and generating an attitude control instruction based on the conversion between an inertial coordinate system and a speed coordinate system; and adjusting the attitude control instruction based on the real-time updated missile time-varying navigation ratio, the missile normal acceleration vector and the target maneuvering parameters to intercept the target. The application can effectively improve the interception precision and stability of the missile in a complex maneuvering environment.

This invention provides a time-sliding mode control method and device for quadrotor unmanned aerial vehicles (UAVs) based on actuator failure, relating to the field of UAV flight control technology. Based on the structural characteristics of the X-shaped quadrotor UAV, this invention establishes the body coordinate system and inertial coordinate system of the quadrotor UAV. Considering the influence of partial actuator failure and bias faults, a dynamic model of the quadrotor UAV is constructed that treats uncertain external disturbances and actuator failures as lumped disturbances. An auxiliary equation is introduced to design a predefined time disturbance observer. Non-singular predefined time-sliding mode controllers based on the position subsystem and attitude subsystem are designed to segment the sliding surface. Based on the sliding mode controller and combined with the disturbance observer, coordinated control of the inner and outer loops of the quadrotor UAV is achieved. This invention can effectively cope with unknown disturbances and actuator failures, significantly improving the system's response speed, robustness, and fault tolerance, and is suitable for high-precision flight missions in complex environments.

This invention provides a method for determining rocket guidance gain relationships based on impact analysis, comprising: acquiring rocket flight trajectory data, including the position and velocity of the first-stage rocket in the launch inertial coordinate system and the theoretical recovery landing point position, and establishing a high-precision landing point prediction model; based on the landing point prediction model, determining the influence coefficient of the shutdown point state deviation on the final landing point deviation, the influence coefficient including the position influence coefficient and the velocity influence coefficient; actively compensating for the current position deviation based on the current velocity deviation in the guidance law, and constructing a second-order equation relationship between the position influence coefficient, the velocity influence coefficient, and the position and velocity deviations, based on the principle that the combined influence of the position and velocity deviations on the landing point should tend to zero, thereby obtaining the constraint relationship between the proportional gain P and the differential gain D. This method, oriented towards determining the guidance gain relationship of the recovery landing point, can meet the requirements of high-precision rocket recovery.

The application discloses a kind of solving method of unmanned aerial vehicle equivalent simulation spacecraft kinematics characteristics, steps are as follows: the inertial coordinate system required for constructing equivalent model and satellite orbit coordinate system, relative motion equation in orbit coordinate system is established;According to the equation, the spacecraft orbit kinematics parameters and unmanned aerial vehicle equivalent kinematics parameters are determined;Combining each parameter basic dimension, dimension matrix is listed, and the basic parameter quantity describing physical process is determined according to the rank of matrix;Based on the similarity of model and prototype, the similarity condition is applied to dimensionless number, the equation describing two systems is established, and then the scale factor of each kinematics parameter is derived.The method makes full use of the controllable characteristics of unmanned aerial vehicle flight height, speed and vertical climbing / descending speed, and realizes the simulation of satellite motion process by unmanned aerial vehicle through formulating reasonable space-time scaling rules, combining spacecraft kinematics characteristics and unmanned aerial vehicle flight control technology.

This invention discloses a time-optimal trajectory planning method for flapping-wing aircraft based on time-varying wind field information, relating to the field of flapping-wing aircraft technology. The method includes: establishing a time-varying wind field model, outputting continuous wind speed and direction that vary with time; establishing a particle kinematic model of the flapping-wing aircraft in both inertial and air coordinate systems based on the time-varying wind field model, and achieving the transformation between the two coordinate systems by designing a virtual target point; deriving the optimal trajectory form of the flapping-wing aircraft in the time-varying wind field model based on the Pontryagin minimum principle; establishing the trajectory equation of the flapping-wing aircraft in the air coordinate system based on the particle kinematic model; establishing equality conditions for solving the trajectory parameters based on the geometric characteristics of the flapping-wing aircraft trajectory; and solving the trajectory parameters using Newton's method based on the equality conditions to obtain the flight trajectory in the inertial coordinate system. This invention can improve the trajectory planning capability of flapping-wing aircraft in complex wind field environments.

This invention discloses a thrust vector control modeling method for a multi-engine parallel aircraft, relating to the field of launch vehicle flight control technology. The method includes: for each engine of the multi-engine parallel aircraft, obtaining the rotation angle from the engine's local yaw coordinate system to its orthogonal inertial coordinate system; combining the vertical yaw angle command and the lateral yaw angle command to determine the transformation matrix corresponding to each engine; obtaining the first thrust of each engine; using the transformation matrix to calculate the second thrust of each engine after yaw in the aircraft's body axis system; obtaining the coordinates of the yaw point of each engine in the aircraft's body axis system; and determining the control force and control torque generated by the yaw motion of all engines based on the second thrust and the yaw point coordinates. This invention simplifies the six-degree-of-freedom dynamics modeling and servo command allocation modeling of rockets, improving model readability and versatility.

A UWB-based heterogeneous unmanned aerial vehicle cluster outdoor cooperative positioning method comprises the following steps: grouping an unmanned aerial vehicle cluster, taking a part of unmanned aerial vehicles as master unmanned aerial vehicles, and taking the other unmanned aerial vehicles as slave unmanned aerial vehicles; establishing a flight inertial coordinate system by using one of the main unmanned aerial vehicles, and determining the main positions of the other main unmanned aerial vehicles in the flight inertial coordinate system by using a first positioning method; determining the relative position of the slave unmanned aerial vehicle relative to the master unmanned aerial vehicle in the flight inertial coordinate system by using a second positioning method; inertial motion data of the slave unmanned aerial vehicle is obtained, and a slave position of the slave unmanned aerial vehicle is generated through an improved extended Kalman filtering model based on the relative position and the inertial motion data. According to the UWB-based heterogeneous unmanned aerial vehicle cluster outdoor cooperative positioning method and the multi-unmanned aerial vehicle system provided by the invention, the positioning precision is higher, and the cost is lower.

The application relates to a satellite orbit and satellite attitude determination method and device after satellite-rocket separation. The method comprises the following steps: calculating a first position vector and a first velocity vector of a satellite in an epoch equatorial coordinate system according to orbit elements of a rocket in a second earth-centered equatorial coordinate system and a satellite-rocket separation time; determining a second position vector and a second velocity vector of the satellite in a J2000 earth-centered inertial coordinate system based on the first position vector and the first velocity vector; taking the second position vector and the second velocity vector as orbit initial values, and determining orbit parameters of the satellite before injection into an orbit by using an orbit recursion algorithm; determining an attitude direction cosine matrix of the satellite in the J2000 earth-centered inertial coordinate system at the satellite-rocket separation time based on a launch time; and taking the attitude direction cosine matrix as attitude initial values, and determining attitude parameters of the satellite before injection into the orbit by using a star gyroscope combination filtering method. The scheme can quickly determine the orbit and the attitude of the satellite, and improves the autonomous detection capability of the satellite.