68 results about "Spacecraft formation" patented technology

The invention relates to a spacecraft fault tolerance attitude cooperation tracking control method based on a normalized neural network, and belongs to the technical field of spacecraft formation flight. According to the method, an attitude motion model of a single spacecraft is built, errors are defined, a control law is designed for the model, a sliding mode function is designed, the derivative of the sliding mode function is solved, an error model is obtained, a control law based on an input normalized neural network is designed, and the states of the spacecrafts are cooperative and consistent; each spacecraft calculates the required control torque according to attitude information of itself and the adjacent spacecraft, the calculated control torque is acted on the corresponding spacecraft via an execution mechanism of each spacecraft, the angular velocity is solved via an obtained attitude dynamic equation, a unit quaternion attitude tracks an expected attitude via the attitude dynamic equation, and consistent attitude of the spacecraft formation is finally realized. According to the method, estimation errors of non-linear function approximation are reduced, the calculating time is reduced, and the convergence rate and the control precision of the system are increased.

The invention discloses a decoupling control method for relative orbits and attitudes of formation satellites, relates to the technical field of the control of the orbits and attitudes of a spacecraft formation, and solves the problems of large satellite calculation amount and low orbit solving efficiency caused by high control dimension of the formation satellites due to serious coupling of the relative orbits and the attitudes of the formation satellites. The method gives two decoupling conditions at first, so that the control of the relative orbits and the attitudes can be designed independently; a thrust vector mobility decoupling constraint condition is introduced to the initialization control of the relative orbits according to satellite attitude mobility constraints indirectly; and during the optimal thrust vector attitude tracking, the possible orientation in space, which meets solar avoidance constraints, of a star sensor optical axis (1) is sought by using a geometric method, and the optimal attitude quaternion and attitude angular velocity are calculated finally by using a double-vector attitude determination algorithm. The method provides important reference value for the control of the orbits and the attitudes of the spacecraft formation.

The invention discloses a spacecraft formation flying relative position cooperative control method based on event triggering. The spacecraft formation flying relative position cooperative control method based on event triggering comprises the steps of: establishing a relative position error dynamics model of a spacecraft formation flying process by considering that an external disturbance condition exists; on the basis of a graph theory, performing communication condition description on a spacecraft formation flying system; by considering that communication resources are limited, designing a virtual speed controller based on an event triggering strategy; and designing actual control input based on the virtual speed controller. By means of the method, the formation configuration is kept same while overall position maneuvering of formation spacecrafts is realized; and furthermore, the method has the advantages of being high in anti-interference capability, effectively reducing the communication capacity between the formation spacecrafts and the like, and is suitable for spacecraft formation flying relative position cooperative control.

The invention discloses an autonomous relative navigation method for multi-information fusion formation spacecrafts, belonging to the field of aviation and space and aiming at solving the problem that current relative navigation in a manner of GNSS plus inter-satellite link leads to no navigation redundancy information so that navigation precision is low. The method according to the invention comprises the steps: absolute navigation information of nodes of a spacecraft is obtained by utilizing a pulsar navigation system and a GNSS satellite navigation system; time synchronous information and distance information in relation to the nodes of other spacecrafts are obtained via a laser link system or a microwave link system and are used for the relative navigation of formation flying of the spacecrafts; the absolute navigation information completes inter-satellite information interaction with the nodes of other spacecrafts via the laser link system or the microwave link system in order to acquire inter-satellite information interaction data of the nodes of other spacecrafts; and in combination with the inter-satellite information interaction data of the nodes of other spacecrafts, the relative navigation information is acquired by processing the absolute navigation information via a house-keeping management system and is used for the relative navigation of formation flying of the spacecrafts.

The invention discloses a radiation open loop tethered satellite formation constant speed spinning deployment control method, relates to a dynamics construction and control strategy design in which the tethered satellite formation is spun and deployed at a constant speed into a radiation open loop form and belongs to the field of spacecraft formation control. In view of the radiation open loop tethered satellite formation running on a circular Kepler track in an earth center gravitational field, in consideration of multi-body system movement characteristics and gravity gradient moment effects, system spinning deployment dynamics is built via a Lagrange equation; on the basis, a closed loop control strategy is designed, the main satellite is spun and deployed at a constant speed, compensation in view of the gravity gradient moment is carried out, damping is added to the tail section of the deployment to restrain system oscillation; the dynamics and the control method are simple, effective and easy to realize; and the whole process of controlling the tethered satellite system to deploy at a constant speed is smooth, safe and high in reliability.

The invention discloses a control method for communication and calculation of a space vehicle formation flight relative position under constraint. The method comprises the following steps of buildinga space vehicle formation flight process relative position error dynamic model; on the basis of a directed graph theory, building an space vehicle formation flight system model; considering the modeluncertainty, designing an adaptive relative position cooperated controller which does not need the relative speed information; considering the communication and calculation constraint; designing the event triggering proper time communication strategy. The method has the advantages that the whole position maneuverability by the space vehicle formation flight system is ensured; meanwhile, the formation structure keeps unchanged; the advantages of high robustness, effective reduction of communication quantity between formation space vehicles and calculation quantity of spaceborne computers and the like are realized; the method is applicable to the cooperated control of the space vehicle formation flight relative position.

The invention discloses a PIO-based round orbital space vehicle formation reconfiguration type anti-collision path planning method, and belongs to the technical field of space vehicle attitude control. According to the algorithm, firstly, a spaceflight short-distance relative motion linear C-W equation is used for dynamics modeling; then, a space vehicle formation discrete dynamic equation is usedfor calculating the total fuel consumption of the space vehicle formation; an infinite series method is used for calculating the collision probability between the space vehicles; a corresponding fitness function is built through an exterior penalty function method; then, a self-adaptive geomagnetic factor coefficient is introduced; the original global search and local search capability of the deterrent algorithm is balanced; the flight optimization path considering the fuel consumption and avoiding collision is finally obtained. The algorithm is simple; the formation fuel consumption and collision avoidance factors between the space vehicles are considered under the condition of given initial and final space vehicle formation; the optimization of the space vehicle formation flight path isrealized.

The invention provides a spacecraft formation relative position control method based on a finite-time distributed type speed observer, and belongs to the field of spacecraft formation relative position control. Firstly, according to a spacecraft relative position dynamics model, the graph theory is used for representing the communication topological relationship between formation members; secondly, the finite-time distributed type speed observer is designed; finally, based on velocity observation information, a potential function method is adopted, and the control method which can make the spacecraft formation members reach their desired positions while avoiding collisions during flight is designed. The method achieves the finite-time estimation of the spacecraft speed, a controller can beindependently designed, and the method is suitable for the relative position control of spacecrafts with missing velocity information. The finite-time distributed type velocity observer is adapted totimely and accurately provide velocity estimation information, the relative position of the spacecraft is ensured to reach expected value, and meanwhile the occurrence of collision is avoided.

The invention relates to the technical field of spacecraft formation reconfiguration, and provides a libration point spacecraft formation reconfiguration method based on a self-adaption agent model. The method comprises the steps that 100, a sun-terrestrial system libration point nearby spacecraft formation controlled kinetic equation is built; 200, according to the spacecraft formation reconfiguration mission requirement, an optimization target is selected to obtain the optimization problem of spacecraft formation reconfiguration; 300, the self-adaption agent model of spacecraft formation reconfiguration is built, and the optimal reconfiguration track of a spacecraft formation is obtained through the self-adaption agent model; 400, the value range of a variable in the self-adaption agent model is narrowed, the step 300 is repeated in sequence, and the validity of the self-adaption agent model is verified. On the premise of the high computational efficiency, the formation reconfiguration result very close to the actual optimal solution can be obtained.

The invention discloses an input-limited finite time attitude cooperative tracking fault-tolerant control method, and belongs to the technical field of multi-spacecraft formation flight. In order to reduce communication paths and avoid resource waste, a communication topological structure which has less communication traffic, comprises a directed spanning tree taking a navigator as a root node andonly has a part of followers capable of obtaining navigator information is adopted among spacecraft formation members. In order to enable a following spacecraft which cannot obtain pilot informationto obtain reference information, a distributed finite time state observer is designed to observe the pilot information. Considering communication bandwidth limitation, actuator fault and saturation, system uncertainty and external interference, based on a finite time algorithm, a redundant fault-tolerant algorithm, an adaptive algorithm and a saturation function, a finite time attitude cooperativetracking fault-tolerant control strategy is perfected, and the robustness and practicability of the control system are further improved.

The embodiment of the invention discloses a method and device for determining a relative orbit of a spacecraft formation. The method comprises: measuring a first motion parameter of a first spacecraftin a spacecraft formation flying around a target object relative to a reference object; according to the first motion parameter, determining a first orbit number of the first spacecraft flying aroundthe target object; measuring a second motion parameter of a second spacecraft in the spacecraft formation relative to the first spacecraft; according to the first orbit number and the second motion parameter, determining a second orbit number of the second spacecraft flying around the target object; and determining an average relative orbital number of movement of the second spacecraft relative to the first spacecraft by combining the first orbit number and the second orbit number.

A spacecraft formation position cooperative control method aims at a master-slave spacecraft formation position cooperation system. An event triggering mechanism is designe, an aperiodic control method based on the event triggering mechanism is formed, the frequency of executing the control method by an airborne microprocessor is reduced, the computational complexity is reduced, aiming at the conditions of external disturbance and state time delay in an actual system, an H-infinity robust performance index is constructed, and parameters of a controller and the triggering mechanism are determined by combining an LMI method, so that adverse effects of external disturbance and time delay are inhibited, robust bounded stability of the system under non-periodic control is ensured, and rapid andaccurate relative position control is realized.

The invention relates to an input saturation self-adaptive gesture collaborative tracking control method and belongs to the technical field of multi-spacecraft formation flight. The method gains the advantages that a bipower algorithm enables a formation system to quickly, stably and dynamically adjust a function optimization control gain to further reduce the saturation incidence, a saturation function controls input saturation amplitude limiting, and an adaptive law suppresses interference and compensation inertia time-varying uncertainty. The invention provides the input saturation self-adaptive gesture collaborative tracking control method which enables spacecraft formation members to quickly complete gesture collaborative tracking. According to the method, influences of input saturation, interference, inertia time varying and the like are comprehensively considered, an input saturation collaborative tracking control strategy is improved, a collaborative tracking error system can be rapidly stabilized, and the robustness and practicability of a control system are further improved.

A spacecraft formation flying method in a weak-stability boundary area of the invention comprises the following steps: (1) solving out an earth-moon low-energy transfer orbit of a main spacecraft based on weak stability boundary properties under a sun-earth-moon-spacecraft four-body model; (2) deriving a relative dynamic equation of an accompanying spacecraft under the sun-earth-moon-spacecraft four-body model; (3) designing a Hamiltonian structure shape maintenance controller; (4) solving out a bounded relative flying orbit of the accompanying spacecraft; and (5) planning a main spacecraft and accompanying spacecraft formation flying orbit based on the earth-moon low-energy transfer orbit of the main spacecraft obtained in step (1). The precision of a formation flying basic dynamic model is improved. The factors of gravity environment change are taken into consideration in formation control. Moreover, a formation flying control method based on a Hamiltonian structure shape maintenance controller is put forward to generate a bounded relative flying orbit in the earth-moon low-energy transfer process.

The invention discloses an anti-collision six-degree-of-freedom cooperative control method for spacecraft formation, and the method comprises the steps: S1, determining a complete mathematic model ofsix-degree-of-freedom movement of spacecraft master-slave formation, and converting a kinetic model of a relative position into a kinetic model related to the distance between spacecrafts; and S2, determining a spacecraft formation anti-collision six-degree-of-freedom preset performance function, performing error model conversion, and determining a six-degree-of-freedom cooperative robust controllaw of the sliding mode variable structure. According to the invention, the sliding mode variable structure control technology and the preset performance control are combined, the controller design utilizes the advantages of the preset performance function, the transient state performance and the steady state performance of the system are both considered, and the anti-collision requirement of thesystem is met; through sliding mode variable structure control, it is ensured that the system has certain robust interference suppression performance under the action of external disturbance; six-degree-of-freedom high-precision cooperative control of the relative attitude and the relative position of the spacecraft considering external interference and anti-collision constraints is realized.

The invention discloses a distributed calculating method for spacecraft dynamics, and belongs to the technical field of control and simulation. The method comprises the steps: employing a distributed calculating strategy, enabling the dynamics simulation operation of a plurality of spacecraft to be distributed to different simulation computers, enabling each simulation computer to just carry out the dynamics resolving for one spacecraft, obtaining the dynamics data of other spacecraft through the data exchange among different simulation computers, and completing the combined simulation testing task of the plurality of spacecraft. The method irons out the defects that a conventional centralized-type simulation technology is not sufficient in calculation capability of simulation computers during the centralized simulation of the dynamics of a plurality of objects and needs to carry out the repeated modeling, and improves the simulation calculation efficiency.

A kind of aircraft formation (F) relative to the spacecraft formation (F) and foreign aircraft (AC0) collision management equipment, the aircraft fleet (F) including leading aircraft (AC1) and at least one UAV (AC2), the equipment includes: formation flight management unit, all aircraft aircraft in the fleet are equipped with formation flight management unit formation flight management unit and is configured for formation flying spacecraft formation in all vehicle management; and the management module and the control module, which is used to determine and application for Spacecraft Formation (F) of all aircraft (AC1, AC2) the coordinated avoidance maneuver, the coordinated avoidance maneuver is determined so as to avoid collision and intrusion aircraft (AC0), at the same time keep the formation Flight. An anti-collision method for an aircraft formation (F) relative to an aircraft formation (F) aircraft (AC0) is also provided.

The invention relates to a spacecraft formation finite time orbit tracking control method considering communication limitation. The method comprises the steps that a spacecraft formation relative orbit model is established; a hysteresis quantizer model is established, and a hysteresis quantizer is adopted to quantize control input signals of slave spacecrafts in the spacecraft formation relative orbit model; in order to facilitate the design of the controller, an intermediate variable is defined; a virtual controller is designed based on a backstepping method and a power integration technology; and according to the designed intermediate variable and the virtual controller, a spacecraft formation finite time tracking controller considering communication limitation is obtained. The controller can ensure that a formation spacecraft realizes tracking control of an expected state in finite time, and the problems of singularity and discontinuity when a finite time tracking controller is designed by adopting a sliding mode control method traditionally can be avoided by adopting a backstepping method and a power integral technology; in addition, the communication load of the control system is effectively reduced through the introduction of a hysteresis quantizer in the design of the controller.

The invention provides a flexible spacecraft finite time attitude cooperative control method for an executing mechanism fault. The flexible spacecraft finite time attitude cooperative control method comprises the following steps: step 1, establishing a non-unwinding flexible spacecraft attitude motion mathematical model based on a rotation matrix; step 2, acquiring attitude information and a reference trajectory of each member in the flexible spacecraft formation, and establishing a relative attitude kinetic equation; step 3, describing a communication topological structure among members of the formation system based on an undirected graph, and designing a fast terminal sliding mode variable in combination with a lumped error; and step 4, designing a self-adaptive fault-tolerant controller in consideration of unknown external disturbance and difficulty in observing flexible dynamic and time-varying inertial parameters. According to the method, real-time estimation and compensation are carried out on uncertain dynamic states based on adaptive control, and the practicability of a spacecraft attitude cooperative fault-tolerant control strategy is remarkably enhanced. According to the method, especially the application of the minimum learning parameter algorithm, the requirement of a control system for computing resources is reduced, so that the purpose of reducing hardware loads is achieved, and the method has very high engineering significance.