16 results about "Spacecraft rendezvous" patented technology

The invention discloses a spacecraft approaching rendezvous guidance method based on limited thrust, and belongs to the technical field of spacecraft guidance. The invention aims to solve the problem that fuel consumption and hardware cost cannot be taken into consideration at the same time in an existing spacecraft short-distance guide section autonomous approaching method. Comprising the steps of calculating a theoretical speed increment of a current task stage; in the execution process of the task stage, the state of the propeller is judged according to a preset period; if the propeller is in the propelling process, the remaining speed increment is calculated by combining the theoretical speed increment and the single-cycle speed increment of the propeller; determining that the propeller is in a non-propelling state until the current residual speed increment is smaller than a set threshold value; the theoretical speed increment is calculated again and compared with a track safety belt threshold value to judge whether the tracking spacecraft exceeds the safety belt range or not, and if yes, the task state is updated, and the propeller is controlled to conduct propelling so as to correct the transfer orbit until the task stage is finished. According to the invention, high-precision guidance control is realized.

The invention discloses a spacecraft rendezvous and docking time synchronization optimization control method based on reinforcement learning, and the method comprises the steps: building a spacecraft six-degree-of-freedom relative motion model which comprises a translation dynamic model based on a Clohessy-Wiltshire equation and an attitude dynamic model based on quaternion; a constrained error space is converted into an unconstrained space through a homeomorphic mapping function, and a tracking error system in a strict feedback form is reconstructed; based on an Actor-Critic reinforcement learning framework, generating an optimal virtual control signal, fusing a time synchronization control item, and forcing synchronous convergence of the position and attitude error in a preset time through a normalized sign function; and performing low-pass filtering estimation on the virtual control derivative by adopting a dynamic surface control technology, generating actual control input meeting the saturation constraint of the actuator in combination with a reinforcement learning framework, and limiting a control input amplitude through a projection operator. According to the method, synchronous convergence of position and attitude errors is realized by combining reinforcement learning and time synchronization control theories.

The invention relates to the technical field of spacecraft control, and provides a spacecraft rendezvous control method based on a disturbance observer and related equipment, and the method comprises the steps: constructing a kinetic model of orbit rendezvous of a target spacecraft at an earth-moon translation point, and designing the disturbance observer based on the kinetic model; defining a cost function based on the system state quantity of the target spacecraft, and constructing an action controller of the target spacecraft according to the cost function; calculating a control interference estimated value of the target spacecraft according to the interference observer, and calculating an optimal control value according to the action controller; and calculating a final control value of the target spacecraft based on the control interference estimated value and the optimal control value, and performing orbit intersection control on the target spacecraft based on the final control value. According to the method, the spacecraft rendezvous control precision can be improved.

A six-degree-of-freedom finite-time space rendezvous control method under multiple constraints belongs to the field of space technology. The implementation method of the present invention is as follows: a six-degree-of-freedom dynamic model is established based on the Lie group SE (3) framework, so that the spacecraft rendezvous process can satisfy the position end constraint while also satisfying the attitude end constraint, fully considering the attitude-orbit coupling problem caused by the thruster installation error in actual engineering, improving the spacecraft rendezvous accuracy, and avoiding the unwinding problem caused by the traditional six-degree-of-freedom dual quaternion representation method. A finite-time sliding mode control strategy is designed so that the rendezvous process can satisfy the mission time window constraint and improve the error convergence speed. In order to solve the potential collision risk in the rendezvous process and meet the safety constraints of the rendezvous mission, an artificial potential function is introduced to achieve collision avoidance in the rendezvous process and improve the safety of the spacecraft rendezvous process. The present invention can shorten the rendezvous time, improve the spacecraft space rendezvous accuracy, and improve the safety of the spacecraft rendezvous process.

The application relates to the technical field of spacecraft control, and provides a spacecraft rendezvous control method based on an interference observer and related equipment, the method comprising the following steps: constructing a dynamic model of target spacecraft rendezvous on a geosynchronous orbit, and designing an interference observer based on the dynamic model; defining a cost function based on system state quantities of the target spacecraft, and constructing an action controller of the target spacecraft according to the cost function; calculating a control interference estimation value of the target spacecraft according to the interference observer, and calculating an optimal control value according to the action controller; calculating a final control value of the target spacecraft based on the control interference estimation value and the optimal control value, and performing orbit rendezvous control on the target spacecraft based on the final control value. The method can improve the precision of spacecraft rendezvous control.

The invention discloses a periodic time-lag sliding-mode control method for a nonlinear spacecraft elliptical orbit intersection system. A nonlinear spacecraft elliptic orbit rendezvous system model subjected to external interference is simplified into a nonlinear periodic system according to reversible linear transformation. On the basis, the nonlinear periodic system is written into a strict nonlinear system matched with interference. A periodic time-lag sliding mode controller with a strict nonlinear system matched with interference is designed by using a periodic time-lag sliding mode control method, and a closed-loop system formed by the strict nonlinear system matched with interference and the controller is stable within a preset true anomaly under the action of the controller. The control design method overcomes the defects that a traditional sliding mode control method of an existing spacecraft elliptic orbit rendezvous system is prone to neglecting control constraints, singularity occurs in preset time control implementation, convergence time depends on an initial value condition, the convergence speed is low, and the traditional sliding mode control method is only suitable for a low-order or single-input system. The problem of spacecraft rendezvous in the preset time is solved.

The invention discloses a spacecraft rendezvous and docking method and device based on step-by-step orbit lifting phase modulation control, and relates to the technical field of space.The spacecraft rendezvous and docking method comprises the steps that on the basis of a preset launching window range, the coplanar phase difference between a target spacecraft corresponding to each launching window and a visiting spacecraft is determined, and based on the determined date, the determined phase error reference, the coplanar phase difference, the phase error reference and the target phase difference, determining a target launching window in a preset launching window range, and based on a step-by-step orbit lifting phase modulation control strategy, launching the visiting spacecraft in the target launching window. According to the invention, the technical problem of low phase modulation control efficiency caused by incapability of accurately determining the launching window in a spacecraft rendezvous and docking task in the related technology is solved.

The application discloses a multi-spacecraft rendezvous orbit interval determination method, and belongs to the technical field of aerospace. The method is as follows: according to the task scale and the carrying capacity of a spacecraft mother ship, the maximum number of sub-spacecraft that can be carried by the spacecraft mother ship is set. The deployment orbit height of the spacecraft mother ship is set, a Lambert problem of the sub-spacecraft released by the spacecraft mother ship and a target spacecraft is established, the sub-spacecraft rendezvous range boundary is solved according to the speed increment capacity of the sub-spacecraft, and a sub-spacecraft rendezvous range database is constructed with the orbit plane angle as input and the sub-spacecraft rendezvous range boundary as output. The minimum rendezvous transfer phase range of all the sub-spacecraft carried by a single spacecraft mother ship is calculated by interpolation using the sub-spacecraft rendezvous range database, and the initial deployment orbit interval covered by the spacecraft mother ship is determined. The process of determining the spacecraft mother ship deployment range is repeated until the corresponding spacecraft mother ship orbit interval of all the target spacecraft is determined.

The invention discloses an anti-saturation control method for a non-cooperative spacecraft rendezvous system with sensor faults. According to the method, in consideration of the deficiency of a non-cooperative spacecraft rendezvous system control method with sensor faults and actuator saturation in the prior art, a second-order all-wheel-drive system model for establishing the non-cooperative spacecraft rendezvous system is proposed, a state observer is designed to estimate a system state vector and interference, an anti-saturation compensator is combined, and the non-cooperative spacecraft rendezvous system control method is established. And the output of the state observer replaces the measured value of the sensor to serve as the input of the controller, an anti-saturation controller is designed, and the non-cooperative spacecraft rendezvous system is controlled. The method is based on a state observer, a sliding mode control theory and an all-wheel-drive system method, the conditions that the system has sensor faults, external disturbance and actuator saturation are considered, and autonomous rendezvous between a tracking spacecraft and a non-cooperative target spacecraft can be achieved.

The application discloses a kind of low-orbit spacecraft's meeting window fast planning method, belong to aerospace field.The application implementation method is: after being applied to the rocket of ground launch reaches the edge of atmosphere, rocket carries out a maneuver, and meets target spacecraft in ascending process.Meeting window for spacecraft meeting target task is obtained.Meeting window fast planning is mainly divided into three steps, first, according to the envelope of rocket shutdown point and pulse size, the ascending reachable range is obtained.Second, the height of this ascending device, the range of the target in the fixed system is used to judge the coarse window.Third, after obtaining the coarse window, the rocket is shot to the target direction, and the shutdown point position in the inertial system is obtained.Based on the point, the Lambert method is used to calculate the velocity increment.If the constraint is satisfied, it is determined that the rocket can meet the target, i.e.the low-orbit spacecraft's meeting window fast planning is realized.The application has the advantages of fast planning speed and strong applicability.

The present application relates to the technical field of spacecraft rendezvous guidance, and is especially a check method and check device for missile-ground communication data based on hash check, comprising: obtaining missile-ground communication data between two communication devices; a binary conversion matrix is installed between the two communication devices, and a key is constructed; the missile-ground communication data is converted based on the binary conversion matrix, and a first parameter is output; the first parameter is converted based on a hash mapping algorithm and the key, and a second parameter is output; the second parameter is compared with a preset key, and if the second parameter is equal to the preset key, the missile-ground communication data is accurate data.

A spacecraft auto-disturbance rejection control method with time delay and actuator saturation includes the following steps: S1: establishing a spacecraft attitude integrated dynamics model with communication time delay and actuator saturation; S2: designing an outer-loop expansion state observer with a time delay predictor; S3: designing an outer-loop auto-disturbance rejection controller; S4: designing an inner-loop expansion state observer; S5: designing an inner-loop auto-disturbance rejection controller with an anti-saturation strategy. The spacecraft auto-disturbance rejection control method with time delay and actuator saturation described in the present invention can not only effectively avoid the adverse effects of external and internal disturbances on the system, but also has a significant inhibitory effect on communication time delay and actuator saturation phenomena, and can ensure that the position and attitude of the spacecraft are accurately controlled within the controllable range of the actuator, providing sufficient guarantee for the smooth completion of tasks such as spacecraft rendezvous and docking, has good control effect, and is widely used in other spacecraft attitude and orbit control systems.

The application discloses a spacecraft rendezvous guidance method based on limited thrust, and belongs to the technical field of spacecraft guidance. The application aims at the problem that the existing autonomous approach method of a spacecraft in a close distance guidance stage cannot simultaneously consider fuel consumption and hardware cost. The method comprises the following steps: calculating a theoretical speed increment of a current task stage; judging a state of a thruster according to a preset period during the execution of the task stage; if the thruster is in a propelling state, calculating a residual speed increment by combining the theoretical speed increment and a single-period speed increment of the thruster; until the current residual speed increment is less than a set threshold value, determining that the thruster is in a non-propelling state; then recalculating the theoretical speed increment, comparing the theoretical speed increment with a track safety belt threshold value, and judging whether a tracking spacecraft exceeds a safety belt range; if yes, updating a task state, and controlling the thruster to perform propelling, so as to correct a transfer orbit, until the task stage is ended. The application realizes high-precision guidance control.

The invention provides an on-orbit spacecraft rendezvous trajectory optimization method based on a Koopman operator, and the method comprises the following steps: building a kinematic model for an on-orbit spacecraft rendezvous task; dynamic constraint convexity is executed on the basis of a Koopman operator; performing no-fly zone process constraint convexity; and constructing a confidence domain and executing iterative calculation. According to the method, global linearization is carried out through a Koopman operator, and through linearization dynamic constraint, the method has higher approximation precision compared with a traditional linearization method; in addition, under the condition that a proper observation function is selected, the sensitivity of optimization solution to the initial conjecture can be further reduced. Therefore, the fitting precision of the original nonlinear system is improved, and the optimization precision and quality of the rendezvous trajectory of the on-orbit spacecraft are improved.

This invention discloses a nonlinear model predictive control method for obstacle avoidance during NRHO rendezvous, belonging to the field of spacecraft rendezvous control technology. Addressing the simultaneous needs of rendezvous mission and dynamic debris avoidance during NRHO rendezvous under a three-body or high-fidelity ephemeris dynamics environment, modeling and control design are conducted within this high-fidelity ephemeris dynamics environment. A three-dimensional mapping method suitable for non-convex obstacle avoidance domains with apertures in three-dimensional space is constructed, mapping the non-convex feasible domain induced by obstacles in the original space to a convex feasible domain in convex space. A nonlinear model predictive control problem with obstacle avoidance constraints is constructed within the mapped space, unifying rendezvous terminal accuracy, control cost, thrust constraints, and obstacle avoidance constraints into a single optimization framework. This achieves unified closed-loop control of rendezvous and docking and dynamic debris avoidance under a three-body or high-fidelity dynamics background.

The application discloses a spacecraft rendezvous and docking method and device based on step-by-step orbit raising phase modulation control, and relates to the technical field of spaceflight. The spacecraft rendezvous and docking method comprises the following steps: determining the coplanar phase difference between a target spacecraft and a visiting spacecraft corresponding to each launch window based on a preset launch window range; determining a phase error reference based on a determination date; determining a target launch window from the preset launch window range based on the coplanar phase difference, the phase error reference and a target phase difference; and launching the visiting spacecraft in the target launch window based on a step-by-step orbit raising phase modulation control strategy. The application solves the technical problem that the launch window cannot be accurately determined in the spacecraft rendezvous and docking task in the related art, resulting in low phase modulation control efficiency.