127 results about "Space vehicle control" patented technology

The present invention provides an emulation system having a control system that allows the testing of a satellite control system with all of its hardware in place, i.e. fully integrated. The emulation methodology is applicable to the case of either a rigid spacecraft or a flexible spacecraft, provided that the spacecraft's sensors and actuators are stowed to the rigid part of spacecraft in the case of a flexible spacecraft. Practically, the latter condition is not restrictive, as the actuators and sensors are usually placed rigidly in the satellite bus, while the satellite solar panels constitute the flexible elements. The control system is used to tune the mass properties and dynamic behaviour of a rigid ground-spacecraft in a 1-G environment to those of a flight-spacecraft in 0-G. A six-axis force/moment sensor is placed at an interface of the ground-spacecraft and a manipulator. Signals received from the force/moment sensor, and in some cases signals relating to the position and velocity of manipulator joints, are received into the control system.

A propellant depot (40, 150) includes a utility box (42, 42′) that has space flight equipment. A propellant cartridge adaptor (95) is coupled to the utility box (42, 42′) and to an exchangeable propellant cartridge system (41). The propellant depot (40, 150) also includes a docking adaptor (44) for coupling to an approaching spacecraft (24). A controller (66) controls the transfer of propellant from within the exchangeable propellant cartridge system (41) to the spacecraft (24). A method of providing propellant to a spacecraft in space includes launching an orbital propellant depot (40, 150) into space. The spacecraft is docked to the orbital propellant depot (40, 150) in space. Propellant is transferred to the spacecraft. The spacecraft is separated from the orbital propellant depot (40, 150).

The invention belongs to the field of spacecraft control and relates to a fuzzy singular perturbation modeling and robust attitude control method for complex flexible spacecraft, namely a robust combined control method for fusing static output feedback control and output integration. The method comprises the following steps of: establishing an uncertain continuous fuzzy singular perturbation model and a standard discrete fuzzy singular perturbation model according to a dynamic model and a kinematic model of the spacecraft in combination with fuzzy logic and singular perturbation technology; and designing a robust controller combined by a static output feedback controller and an output integrator by a spectral norm and linear matrix inequality (LMI) method and resolving a group of LMIs which are unrelated to a perturbation parameter so as to obtain a controller parameter and solve an ill-conditioned problem caused by the perturbation parameter and the problem of difficulty in selection of an initial value in an LMI resolving static output feedback controller gain method. Through the method, flexible vibration and external interference can be overcome effectively, and control effects such as high response speed, high attitude control accuracy, high anti-jamming capability and high robust performance are achieved.

The invention relates to a method for controlling an index time-varying slide mode of flexible spacecraft characteristic shaft attitude maneuver, and belongs to the technical field of spacecraft control. The method comprises the steps that firstly, a system dynamically equivalent model, a dynamic model and a flexible vibration model are established under a spacecraft system, then, the vibration frequency and the damping ratio parameter of a closed loop system with the index time-varying slide mode control law are calculated, and a single-shaft multi-modality filtering input shaping device with a characteristic shaft as a rotary shaft is designed according to the designing method of the single-shaft input shaping device to restrain flexible vibration in three-shaft motion. Meanwhile, a state observer is designed to estimate flexible modal information in real time, and the method for controlling an output feedback index time-varying slide mode is formed. At last, saturability analysis is conducted on control torque so as to satisfy the physical saturation constraint of the control torque. By means of the method, the application range of existing input shaping is expanded, the input shaping technology is expanded from single-shaft maneuver to three-shaft maneuver, the self-robustness of filter input shaping is enhanced, and the purpose that the attitude maneuver path of the spacecraft is the shortest is achieved.

The invention belongs to the field of spacecraft control, and relates to a fault mode thrust allocation method for a geostationary orbit satellite electric thruster. The method comprises the steps of firstly, determining a thrust allocation input condition, including ignition position constraint, ignition velocity increment constraint and orbit control demand; secondly, allocating and optimizing position keeping thrust, establishing an electric thruster pointing model, selecting a thruster, defining ignition parameters, and optimizing and calculating the ignition parameters; and finally, calculating an angular momentum unloading deflecting vector. Aiming at the problem of high fuel consumption in a traditional definitive solution mode, a nonlinear hybrid constraint optimization model is established, and key parameters such as time length of each thrust ignition arc, ignition position and the like are solved by taking fuel optimization as a target, so that the method can realize fuel optimization under engineering constraints.

A method and apparatus for estimating attitude sensor bias in a satellite system uses attitude sensors and a spacecraft control processor. Attitude sensors provide output signals, which may contain bias. The present invention interprets the signals, determines the bias present in the signals, and generates an output signal to offset the bias in the signals from the attitude sensors, thereby leading to more accurate positioning of the satellite employing the present invention.

Provided is a Flight Gear general three-dimensional scene data displaying method based on a field programmable gate array (FPGA). The method includes that an FPGA chip is adopted, coding of Verilog hardware description language (HDL) language is conducted on the FPGA chip, and a data transmission protocol 1 is self-defined so as to enable the FPGA chip to finish receiving of serial data of specific frame format and with the baud rate of 115200 bps, analysis of data of a self-definition communication protocol 1 is finished in the FPGA, and the data is processed through corresponding algorithms. Then, the processed data is coded and packaged through another self-definition communication protocol 2 and transmitted to a Simulink project operated on a personal computer (PC) with the baud rate as 115200 bps, a series of data processing is conducted in the Simulink. Finally, corresponding gesture data is transmitted to Flight Gear software through a user datagram protocol (UDP) network transmission module to be displayed in real time in a three-dimensional (3D) mode. The Flight Gear general three-dimensional scene data displaying method has the advantage that design and simulation of an aerospace vehicle controller, simulation of an unmanned aerial vehicle controller, simulation of guided missile control, simulation of vehicle and ship controllers and 3D visual reproduction of actual aircraft test flight data and the like can be well achieved, and a use range is wide.

An emulation system includes a central time source generating a time reference and an emulated spacecraft control processor which contains an embedded processor that provides an emulated input/output interface to communicate simulated spacecraft data. The embedded processor processes the simulated spacecraft data and contains a real time clock engine having a real-time clock period. The system further has a first simulation that processes attitude control system data from the emulated spacecraft control processor to simulate an attitude control system of the spacecraft in real-time. The first simulation engine operative to produce sensor data for input to the emulated spacecraft control processor based on the simulated system dynamics and adjusts the real time clock period in response to the time reference.

The present invention provides a rigid-flexible liquid coupling system attitude controller and maneuvering path combination optimization method. The method comprises the following steps of: establishing a dynamical model of a liquid-filled flexible spacecraft; obtaining an angular acceleration curve, an angular speed curve and obtained multi-section angular position curve of the liquid-filled flexible spacecraft, and planning the attitude maneuvering path of the liquid-filled flexible spacecraft; calculating the expression of each section of the curve in the multi-section curve; employing a PDcontrol system to perform attitude control of the liquid-filled flexible spacecraft; and performing combined optimization of the parameters of the liquid-filled flexible spacecraft controller and themaneuvering path. The method reduces the motivation of the attitude maneuvering for the flexible accessory vibration and liquid shaking so as to achieve large-angle rapid maneuvering and rapid and stable control for the liquid-filled flexible spacecraft.

The invention relates to a vibration and loading combined device of a steering engine room, and belongs to the technical field of spacecraft controllers. The combined device is capable of control surface loading and vibration clamping of a spacecraft. The device is composed of a vibration tool and a torsional moment loading tool. The vibration tool as a clamp can fix the steering engine room to avibration table by switching. Four elastic rod assemblies are mounted on the vibration tool, and elastic torsional moment loads are applied to four steering engines in a stern room simultaneously viatorsinoal deformation of elastic rods. The vibration and loading combined device is characterized by being multifunctional, low in cost, easy to dismount, simple in structure, smooth in load output, consistent in elastic coefficients, and the like.

The invention discloses an executing agency normalized reachable set peak-based control allocation method, which belongs to the field of spacecraft control, and aims at solving the problem that the allocation control method of the traditional redundancy executing agency configuration scheme cannot achieve large allocating space, strong real-time calculating capability and small memory space occupation at the same time. The method of the invention comprises the following steps: 1, judging whether an executing agency has fault information, wherein if an executing agency has fault information, a step 2 is performed, or a step 3 is performed; 2, calculating and updating the reachable information of the executing agent offline; 3, normalizing an expected control quantity given by a system, and forming an expected unit torque point by intersecting with a reachable information unit ball; 4, determining n peaks of a normalized reachable enveloping surface adjacent to the expected unit torque point on said reachable information unit ball; 5, checking one by one to determine a reachable enveloping surface intersected with a ray in the direction of the expected unit torque; and 6, finishing calculating the control quantity of each executing agency according to the control quantity corresponding to the peaks of the reachable enveloping surface.

The invention relates to a sampling control method for a spacecraft, more particularly to a sampling control method for relative motion of spacecrafts. According to the current sampling control method for a spacecraft relative motion, a processing period and a deviation of a digital controller are neglected, so that an accuracy and security of a spacecraft track are influenced; however, the above-mentioned problems can be solved with utilization of the sampling control method provided in the invention. The method comprises the following steps that: step A, a dynamical model on a spacecraft relative motion is established; step B, sampling is carried out on relative states of two spacecrafts; step C, an M matrix and an N matrix are constructed by utilizing an upper boundary line and a lowerboundary line of a sector area described in the step B; step D, a corresponded state feedback control law is obtained; step E, two positive definite symmetric matrixes P and Q are introduced and a following lyapunov function is defined; step F, an intersection process is obtained and is completed, wherein a trust satisfies a formula upper boundary constraint condition (3); and step G, a feasible solution is obtained by utilizing an LMT of MATLAB software. According to the invention, the sampling control method can be applied in design of a spacecraft controller.

The invention provides a diagnostic determination method for a spacecraft control system under influence of noise. First, a multivariate probability distribution statistic model of the spacecraft control system is built through a standard model and an equivalent space method; then a detectability index is obtained; and fault detectability is determined; finally, an isolability index is obtained, and the fault isolability is determined. According to the diagnostic determination method, influence of process noise, observation noise and other interference factors on diagnostic performance are fully taken into consideration, fault diagnosis algorithms are not required to be designed, and the fault detectability and the isolability can be determined according to system information including kinetics, kinematics, a controller model and the like of the spacecraft control system and the configuration and installation condition of a sensor and an actuator, the fault diagnosis process of the spacecraft control system is simplified, meanwhile, a theoretical basis can be provided for design of a diagnosis algorithm, the design method of the spacecraft control system is optimized, and the controllability of the spacecraft control system is improved.

A system for counteracting a disturbance in a spacecraft includes a biasing apparatus that is coupled to the spacecraft and a spacecraft controller within the spacecraft. The disturbance has a known sign, magnitude and time. The biasing apparatus controls the biasing apparatus to place the spacecraft in a first dynamic state or position as a function of the sign, magnitude, and time of the disturbance. The controller also controls the spacecraft to a second dynamic state as a function of the known sign, magnitude, and time so that the spacecraft is oriented in a position other than the desired orientation so that after the disturbance the spacecraft is oriented in the desired orientation in response to the disturbance. The biasing apparatus may comprise a momentum wheel and the disturbance may comprise thrusting firing used for controlling momentum dumping.

The invention discloses a magnetic propelling device for a spacecraft. The structural design is simple, the propelling device adopts a strip-shaped coil structure, and a higher total coil current can be generated through multi-circle wire-wrap connection of a low current wire; the direction and size of pushing force can be adjusted along with the direction of the coil and the size of the coil current, the direction of the pushing force is vertical to the direction of the coil current, and the spacecraft adopting the magnetic propelling device disclosed by the invention is simpler to control. Compared with a magnetic moment propelling method, the pushing force in a propelling method of the magnetic propelling device for the spacecraft is higher under the environments of equal current intensity and earth magnetic field; moreover, a plurality of propelling devices can be arranged in an overlapped mode for combined use, so that high pushing force propelling of the spacecraft is realized; when the propelling device works, a magnetic field is generated around, space charged particles are deflected under the effect of the magnetic field when getting close to the spacecraft, and the device has certain shielding and protecting effects on the space charged particles for the spacecraft; moreover, the magnetic moment sizes of the propelling devices counteract with each other, and no extract magnetic moment effect is generated for the spacecraft.

The invention provides a flexible spacecraft adaptive attitude control law based on modified Rodrigo parameters, which includes the following steps: S1, establishing a kinematics equation and a dynamics equation of a flexible spacecraft based on modified Rodrigo parameters; and S2, using an adaptive back-stepping method to design an adaptive attitude controller based on a modal observer. The beneficial effects of the invention are as follows: the problem that a flexible spacecraft is suppressed by interference signals is solved, and attitude control is realized; and through computation using modified Rodrigo parameters (MRPs), the amount of computation can be reduced by 20-40% compared with computation using quarternions, and the real-time performance of control on a flexible spacecraft can be better enhanced.

The invention relates to the field of complex spacecraft control technologies and simulation verification, and aims to perform real-time simulation verification and achieve the purpose of verifying the real-time performance, safety and reliability of a designed control algorithm. Therefore, the technical scheme adopted by the invention is a spacecraft attitude stability control real-time simulation verification and three-dimensional demonstration method, which comprises the following steps: 1, finite time control algorithm design under input constraints: designing a finite time additional system dynamic state, and designing a controller by combining backstepping control on the basis to realize finite time attitude stability control of a spacecraft; 2, structural design and construction ofa spacecraft real-time simulation platform: performing structural design and construction work of the spacecraft real-time simulation platform according to the control algorithm verification requirement and the spacecraft control structure proposed in the step 1; and 3, design and realization of spacecraft master control software. The method is mainly applied to complex spacecraft control design occasions.

The invention discloses a spacecraft obstacle avoidance control method based on ellipsoid description. The method is used for achieving autonomous obstacle avoidance of a target spacecraft and a tracking spacecraft and comprises the steps: establishing a coordinate system; establishing a relative kinetic equation; determining the shortest distance between an obstacle and the tracking spacecraft; establishing an artificial potential function; calculating an attraction control force; calculating a repulsion control force and calculating a total control force. According to the spacecraft obstacleavoidance control method based on ellipsoid description, the ellipsoid is adopted to describe the appearances of the spacecraft and the obstacle, and the modeling precision can be improved so as to improve the spacecraft control precision; meanwhile, a repulsion potential function is designed by applying a Sigmoid function to generate an obstacle avoidance control force; the attraction potentialfunction is designed by using the state-dependent Riccati equation to generate the attraction control force, and the integrated design of the corresponding controller is completed by using the terminal sliding mode control theory, so that the rapid obstacle avoidance control of the spacecraft can be realized; the control precision is high, the fuel consumption rate is low, and the method can be suitable for the on-orbit real-time operation of the spacecraft.