62 results about "Spacecraft rendezvous" patented technology

A closed-loop LRF pointing technology to measure the range of a target satellite from a chaser satellite for rendezvous is provided that includes several component technologies: LOS angle measurements of the target satellite on a visible sensor focal plane and the angles' relationships with the relative position of the target in inertial or LVLH frame, a relative navigation Kalman filter, attitude determination of the visible sensor with gyros, star trackers and a Kalman filter, pointing and rate commands for tracking the target, and an attitude controller. An analytical, steady-state, three-axis, six-state Kalman filter is provided for attitude determination. The system and its component technologies provide improved functionality and precision for relative navigation, attitude determination, pointing, and tracking for rendezvous. Kalman filters are designed specifically for the architecture of the closed-loop system to allow for pointing the laser rangefinder to a target even if a visible sensor, a laser rangefinder, gyros and a star tracker are misaligned and the LOS angle measurements from the visible sensor are interrupted.

The invention relates to a super-close optimized collision-avoidance proximity method for a failure satellite, and belongs to the technical field of spacecraft encounter. An object, namely the failure satellite, is designed into an envelope model in the sphere and ellipsoid combined form to simplify configuration of the object; posture rolling of the object is considered, a relative dynamic model of the object and a tracked star as well as constraints of the path of the tracked star are derived in the dynamic object system; positional nondeterminacy caused by navigation and measurement errors is considered, and the flight forbidden area of the tracked start is further broadened by combining the collision possibility problem; and a safe collision avoidance path is planned on the basis of the Gauss pseudospectral method, and closed-loop feedback control is carried out. According to the invention, spatial limitation in super-close proximity are considered, and requirements for a safe collision free task can be met; and the posture coupling feature of a close proximity spacecraft is emphasized, and whether the distance between the aircrafts satisfies the constraints can be directly determined.

The invention provides a method for determining an orbit transferring strategy of time-limited spacecraft coplanar rendezvous, and belongs to the field of the time-limited spacecraft coplanar rendezvous. The problem that the accuracy of the rendezvous algorithm calculation is affected caused by the poor observation effect at rendezvous time due to the fact that an existing spacecraft rendezvous technology has relatively low requirements on lighting conditions is solved. The method comprises the steps that the theoretical limiting time of the entire rendezvous process between a target spacecraft and a service spacecraft is calculated; then the two speed increments and the used time of orbit transferring in the service spacecraft orbit transferring process are calculated; initial values of the movement time of an initial orbit and initial values of the movement time of a target orbit for the optimized variable service spacecraft are calculated; integral is carried out on the rendezvous process to obtain position vectors of the service spacecraft and the target spacecraft at the rendezvous time and the angle of the position vectors; when the angle reaches 0 degree, at the moment, themovement time of the initial orbit and the movement time of the target orbit are used as the orbit transferring parameters and the orbit transferring strategy. According to the method, the calculationresults are accurate, and the iterative calculation process is simple.

A simulation system for rendezvous and docking of a spacecraft comprises a horizontal support module, a target six-degree-of-freedom air bearing table, a tracking six-degree-of-freedom air bearing table, a master control module, and a monitor and display module. Actual features of weight and quality of a rendezvous and docking spacecraft can be simulated by six-degree-of-freedom air bearing tables, a six-degree-of-freedom test on a last translation closing stage can be carried out depending on actual rendezvous and docking spacecraft-mounted devices such as a docking mechanism. Compared with the prior art, in the simulation system, an indoor GPS unit module can measure positions and attitudes and transmit to the master control module and the monitor and display module in real time, the rendezvous and docking process of a spacecraft can be visually, accurately and real controlled and simulated, and the success rate of a simulation test can be raised. In addition, the simulation system can verify the rationality of a work sequence during the rendezvous and docking process of the spacecraft, the reliability and validity of the rendezvous and docking process of the spacecraft can be improved.

The invention discloses a wireless electric power and signal transmission system for spacecraft rendezvous and docking parallel-operated power supply. One or a plurality groups of wireless electric power transmission units are configured on a spacecraft rendezvous and docking mechanism; each group of transmission units consist of electric power transmitting devices and electric power receiving devices. Wireless transmission of electric power and signals can be achieved by using a technique based on electromagnetic induction principles, under same power transmission conditions, the weight and the size of the system are equivalent to those of a wired electric power transmission mode, the electric power transmission efficiency of the system is greater than 85%, that is, the electric power transmission efficiency is higher than that of the wired transmission mode. By adopting the wireless electric power and signal transmission system, the problems that a power supply mode of a conventional electric connector has high abutting precision requirements, abutting interface design is complex, potential hazard of electric sparks can be caused as contact metal is exposed, and the like, can be solved.

The invention relates to a method for extracting pose measurement feature points based on target feature modeling, in particular to a method for extracting feature points under complex celestial background on the basis of target feature modeling and belongs to the technical field of rendezvous and docking. The method includes: firstly, setting four low-consumption light emitting diodes forming a diamond on a target spacecraft; using a spacecraft trace camera to acquire optical feature point images under the complex celestial background in the approach stage of rendezvous and docking of the spacecraft, and performing adaptive threshold binarization and smooth filtering to obtain a binary image; and searching planar diamond structures in the image to obtain image plane coordinates of the spacecraft optical feature points. Interference of the celestial background can be eliminated effectively, random noise can be eliminated, optical plane coordinates of the optical feature points can be obtained easily, results are stable and reliable, and the method is easy to implement.

The invention provides a spacecraft rendezvous and docking multi-degree-of-freedom semi-physical simulation method and a device thereof. The device comprises a machine body, a cross beam, a vertical beam, a three-shaft turntable, a motion simulation device controller, a GNC system and a measurement system. The machine body is mounted on a ground base. The cross beam is placed on the guide rail of the machine body. The vertical beam is installed on the guide rail of the cross beam. The three-shaft turntable is installed on a vertical beam guide rail. The motion simulation device controller, the GNC system and the measurement system are placed next to a system. Based on a similarity theorem and a scaling ratio criterion, the effective travel range of a simulation device is determined according to an aircraft actual flight range, and assuming that a scaling parameter is k, according to the actual flight condition of the aircraft, assuming that an aircraft true maximum speed, an acceleration, an angular speed and an angular acceleration are v[max], a[max], omega[max], and alpha[max] respectively, the maximum speed, the acceleration, the angular speed and the angular acceleration of the simulation device are determined as k*v[max], k*a[max], omega[max], and alpha[max]. The method and the device have the advantages of a wide application range, a simple structure, and multi degrees of freedom.

The invention discloses a gain switching method of a spacecraft rendezvous system and maximal attraction domain estimation. The method comprises steps that firstly, a spacecraft rendezvous relative movement model is established, secondly, the gain switching method of the spacecraft rendezvous system is proposed, and lastly, the maximal attraction domain estimation of a closed loop system can be given through the gain switching method. According to the method, a spacecraft asymmetric performer saturation control problem is converted into a performer symmetric saturation control problem, controller calculation only requires solving linear matrix inequalities, calculation is simple and easy to carry out, dynamic performance of the closed loop system is improved, and the maximal attraction domain estimation of the closed loop system is further given.

The invention relates to a time-varying feedback finite-time stabilization method of a control limited spacecraft rendezvous control system. The stabilization method comprises the steps of: (1), establishing an orbital dynamics model of the control limited spacecraft rendezvous control system, and obtaining a state-space equation; (2), establishing a parametric Lyapunov equation, analyzing the property thereof, according to the positive definite solution P(gamma) of the parametric Lyapunov equation, designing an explicit linear time-varying feedback control law in a control limited condition,namely, designing a state feedback controller of the control limited spacecraft rendezvous control system; and (3), designing controller parameters by constructing an explicit Lyapunov function and utilizing the property of the parametric Lyapunov equation solution, and ensuring that a tracking spacecraft and a target spacecraft complete a rendezvous task in limited time. By means of the time-varying feedback finite-time stabilization method of the control limited spacecraft rendezvous control system in the invention, limited time stabilization of the spacecraft rendezvous control system in the control limited condition can be realized.

The invention provides a gradient separate zone optimization design method for a spacecraft pulse rendezvous trajectory, belongs to orbit design and optimization technologies in the field of spacecraft orbit dynamics and control, and in particular belongs to a rendezvous trajectory design technology in which a spacecraft rendezvous trajectory global optimization solution is obtained through calculation based on an interval arithmetic mathematical model. For a global issue of the method, a gradient optimization result is only used to determine an interval separation point and update a value of an upper bound of a target function, thereby preserving globalness of an interval optimization algorithm. For a large computation amount in the interval optimization algorithm, a gradient optimization method is introduced, so that the value of the upper bound of the target function can be quickly updated, and a gradient segmentation method can be used; in combination with a symbol segmentation method and an interval optimization strategy, intervals that do not comprise an optimal solution are effectively separated and removed, thereby improving operational efficiency. To solve the problem that the interval optimization algorithm has a high storage requirement, only a limited number of intervals are selected for processing in each iteration, so as to ensure that a storage space overflowing phenomenon does not occur in an operation.

The invention relates to a method for controlling an autonomous rendezvous between spacecrafts under action of impulse thrust, which belongs to the technical field of aerospace, and aims to solve the problem that a traditional method for controlling the rendezvous between the spacecrafts under the action of the impulse thrust is easy to be influenced by a disturbance moment because of adopting anopen-loop control mode. The method for controlling the autonomous rendezvous between the spacecrafts under the action of the impulse thrust comprises the following steps: establishing a kinetic modelof the relative motion between the spacecrafts, and converting a state space model of the relative motion into a discrete motion model; introducing state feedback control rate in the impulse action process; introducing virtual energy functions for a impulse action motion and a free motion; determining three inequalities meeting the requirement that the spacecrafts realize the autonomous rendezvous, satisfying a finite impulse thrust inequality, and converting the inequalities into linear matrix inequalities related with X1,X2 and Y1; and calculating a state feedback gain matrix K according tothe worked-out X1 and Y1 matrixes to obtain a state feedback sampling control law u(k)=Kx(k) of the relative motion between the spacecrafts, which meets the design requirement. The method for controlling the autonomous rendezvous between the spacecrafts under the action of the impulse thrust is suitable for the autonomous rendezvous process of the spacecrafts.

The invention belongs to the technical field of teaching demonstration devices, and discloses a magnetic levitation science education device for demonstrating rendezvous and docking of spacecrafts. The device includes two motors enabling two tracks having different lengths to make circular motions around the center of a circle through transmission gears; electric push rods, installed on the tops of the motors, wherein each electric rod is connected to one of the motors through a transmission gear; and a magnetic levitation device used for adjusting the revolution speed and direction of the motors, wherein the magnetic levitation device is connected to the motors by conductors. Based on a dynamics model, a control mathematical space of relative postures and positions of rendezvousing spacecrafts is constructed and a proportional differential control parameter design method is applied to control of relative positions. The magnetic levitation science education device for demonstrating rendezvous and docking of spacecrafts uses the magnetic levitation technology, the lever principle, the transmission principle and so on, and the theoretical knowledge of spacecraft rendezvous and docking is studied in depth during production. A carefully selected docking presentation mode well reflects the postures and trajectories of rendezvous and docking of spacecrafts.

The invention relates to a system test method for a spacecraft docking flight mission. The method includes the following steps: S1, constructing a test system for testing a spacecraft docking flight mission, and adjusting the test system to a real flight mission state; and S2, testing the test system according to a test project. By really carrying various equipment in an spacecraft rendezvous anddocking mission through an astronaut system, a manned spacecraft system, a target spacecraft system and a measurement and control communication system, interfaces among the astronaut system, the manned spacecraft system, the target spacecraft system and the measurement and control communication system can be verified according to a real on-orbit flight mission, the effectiveness of various systemtests during the docking flight mission can be improved, and the successful implementation of an actual rendezvous and docking flight mission can be further ensured.

The invention provides a rendezvous control method for spacecrafts between orbits at instable libration points of solar-terrestrial systems. The rendezvous control method includes determining rendezvous initial orbits, target orbits and rendezvous process time of the spacecrafts between the orbits at the libration points of the solar-terrestrial systems and building rendezvous kinetic models of the controlled spacecrafts; creating nonlinear optimal control problems on the basis of the built rendezvous kinetic models of the controlled spacecrafts; solving the nonlinear optimal control problems by the aid of symplectic preservation numerical processes within a long finite time to obtain co-state variables; updating control input in current time subintervals by the aid of the obtained co-state variables, kinetically simulating each time subinterval and acquiring state variables of the current spacecrafts by the aid of navigation processes; carrying out progression on the time subintervals, utilizing terminal states of the current spacecrafts as initial states of a next time subinterval, and sequentially repeatedly carrying out procedures from a step 302 to a step 400 until rendezvous tasks of the spacecrafts between the orbits at the instable libration points are completed. The rendezvous control method has the advantage that the high-precision spacecraft rendezvous objectivity and performance can be guaranteed by the rendezvous control method.

The present invention discloses a Tau theory-based spacecraft rendezvous guidance method. The Tau theory-based spacecraft rendezvous guidance method is provided according to material and energy transfer tasks between orbit spacecrafts. With the Tau theory-based spacecraft rendezvous guidance method adopted, a suitable rendezvous trajectory can be designed. According to the method of the invention, a relative motion model in a near-distance rendezvous process is analyzed; the expression form of the rendezvous trajectory is given based on the Tau theory; derivation is performed through using the rendezvous trajectory, so that a speed required in the rendezvous process is obtained; and an accelerated speed required in the rendezvous process is solved through using an inverse dynamics method on the basis of the speed required in the rendezvous process. The method provided by the invention is a bionic method; the instinctive movement of animals in the nature, which has been continuously optimized through thousands of years, is concluded and summarized, and therefore, the method has certain optimality; and the rendezvous trajectory designed by using the method can not only satisfy rendezvous requirements, but also has excellent adaptability.

The invention discloses a spacecraft autonomous rendezvous and docking guidance strategy generation method based on reinforcement learning. The method comprises the following steps: modeling a spacecraft rendezvous and docking process into a Markov decision process model; solving the Markov decision process model by adopting a dynamic programming algorithm to obtain scores of different actions adopted in all states, and generating a decision table; taking all states in the decision table as training data features, taking scores of all states in the decision table under each action as training data labels, and constructing training data; constructing a neural network model, and training the neural network model by adopting the training data to obtain a neural network model serving as approximate representation of the decision table; for a certain state, calculating scores of all actions in the state through the obtained neural network model, and selecting the action with the maximum score as the optimal guidance strategy; and based on the optimal guidance strategy, enabling the spacecraft to perform autonomous rendezvous and docking.

The invention discloses a spacecraft rendezvous and docking trajectory planning event constraint convexity method, and belongs to the technical field of aerospace. The implementation method comprises the following steps of: discretizing an event constraint solution space to obtain a series of spatially discrete grids; calculating a constraint function value at each discrete grid boundary point, and performing interpolation on constraint function data to obtain a linear expression form of a constraint function; constructing an event detection function, and performing relaxation and convex grid judgment constraint on an event constraint function to obtain a convex event constraint function form; constructing a switching function, describing an original event trigger function, and obtaining a convex expression form of the event trigger function; and solving the convex expression form to obtain an optimal spacecraft rendezvous and docking trajectory, thereby improving convergence and efficiency of spacecraft rendezvous and docking trajectory planning with event constraints on the premise of not changing an original problem solution space.

The invention relates to a method for controlling optimal continuous thrust under J2 perturbation by adopting relative navigation information, and relates to the field of spaceflight orbit control. The method aims at solving the problem of rendezvousing of two spacecrafts operating on an elliptical orbit under J2 perturbation, and the continuous thrust optimal orbit control method which achieves rapid calculation and can be used in orbit is provided under the situation that information of the target aircraft is unknown, and the tracking spacecraft only adopts the relative navigation information. The method does not need an optimized initial predicted value, the influences of earth J2 perturbation on the two-satellite approaching process are considered, and the method is suitable for continuous thrust orbit rendezvousing under the elliptical orbit. The method faces the field of orbit rendezvousing, autonomous orbit rendezvousing of the tracking spacecraft to the target spacecraft can beachieved only through relative motion information of the tracking spacecraft and the target spacecraft without the need of support provided by the ground, and thus quite high on-orbit application value is achieved.

The invention discloses an autonomous spacecraft rendezvous control method based on a finite time fast nonsingular terminal sliding mode. The method comprises the following steps: establishing an approximate relative motion kinetic model between a tracking spacecraft and a tracked spacecraft so as to describe an action relationship of real-time control input force and lumped uncertain interferenceacting on the tracking spacecraft on an actual relative position and an actual relative speed of the two spacecrafts; designing a finite time adaptive disturbance observer to estimate lumped uncertain disturbance according to the current actual relative position, the current actual relative speed and the real-time control input force to obtain a disturbance estimation value; and designing a control unit comprising a fast nonsingular terminal sliding mode surface so as to update the real-time control input force according to the interference estimation value, the current actual relative position and the current actual relative speed. Compared with an autonomous spacecraft rendezvous linear sliding mode control method, the control method provided by the invention can realize faster dynamicresponse characteristics and higher steady-state precision.

The invention discloses a carrier rocket task planning method for rapid spacecraft rendezvous, and relates to the technical field of aerospace. The method comprises the following steps of S1, determining a longitude adjustment value of geographic longitudes of ascending nodes of a tracking spacecraft according to a propellant reserved by a carrier rocket; S2, calculating a geographic longitude interval in which the tracking spacecraft can rendezvous with a target spacecraft after the tracking spacecraft enters an orbit according to the longitude adjustment value and a plurality of selectable launching sites of the carrier rocket; S3, determining the geographic longitudes of the theoretical ascending nodes and a selected launching site of the carrier rocket according to the geographic longitude interval and the geographic longitudes of the ascending nodes of all circles of the target spacecraft orbiting the earth in one day; S4, calculating theoretical launching time according to the selected launching site and the geographic longitudes of the theoretical ascending nodes; and S5, launching the carrier rocket at the selected launching site and the theoretical launching time to enablethe tracking spacecraft to enter a rendezvous orbit. According to the carrier rocket task planning method, at least one launching moment can be ensured in one day, so that the tracking spacecraft canrendezvous with the target spacecraft rapidly.

The invention discloses an improved solving method for a Lambert orbital transfer problem based on a Newton iteration thought. The method comprises the steps of 1, calculating a transfer angle theta,an auxiliary variable c and an auxiliary variable s; 2, determining the shape of the track according to the transfer angle and the auxiliary variable in the step 1; 3, when delta t is larger than delta tp, determining that the transfer orbit is an elliptical orbit, and carrying out solving; 4, when delta t is smaller than delta tp, determining that the transfer orbit is a hyperbolic orbit, and carrying out solving; 5, when delta t is equal to delta tp, determining that the transfer orbit is a parabola orbit, and carrying out solving; 6, expressing the Lambert orbital transfer problem analytical equation in the steps 3-5 as a function of a conic curve semi-major axis, and calculating a conic curve semi-drift diameter; 7, calculating orbital transfer point speed and intersection speed; and 8, proving the high efficiency and accuracy of the improved solving method for the Lambert orbital transfer problem based on the Newton iteration thought. The invention aims to improve the solving speed of the Lambert orbital transfer problem in spacecraft rendezvous.