169 results about "Orbital control" patented technology

A star sensor attitude-determination method during the autonomous orbital control fault recovery includes: (1) predicting the satellite inertia attitude according to data measured by a gyro; (2) calculating the filtering modified innovation amount according to the inertia attitude of the satellite and the optical axis vector and the lateral axis vector under an inertial coordinate system which are measured and output by the star sensor, and calculating the error of the innovation amount between the front and the back periodicals, which is used for judging the consistency of star sensor data; (3) judging the consistency of the star sensor data; (4) fixing the attitude of the star sensor according to the two vectors; (5) introducing the star sensor which is combined with the gyro for the correction of the satellite attitude under the condition that the star sensor data arranges the initial value of the attitude estimation. The method can improves the reliability of the orbital control fault recovery, saves the time for the fault recovery and ensures that the orbital control is accurately recovered in time.

An active satellite dispenser is preferably attachable to a reusable launch vehicle for deployment of one or more satellites into one or more desired orbits. The active satellite dispenser includes a center mast that releasably receives the satellite(s), a liquid propellant rocket, and an orbital control system on an avionics pallet. In the preferred embodiment, a pressurized gas selectively pressurizes the propellant tanks (which may include fuel and oxidizer tanks), to provide propellant to the rocket. In operation, the launch vehicle releases the satellite dispenser in a first deployment orbit. The active dispenser rocket and orbital control system then transport the active dispenser and satellite(s) into the final deployment orbit. In the preferred embodiment the active dispenser can operate multiple times to place individual satellites in different orbits.

The invention relates to the research field of a microgravity simulating platform of a space robot, in particular to a microgravity air floatation target satellite simulator system with five degrees of freedom. The system consists of three parts, namely a movement simulator, a satellite simulator and a ground control panel. The movement simulator simulates the states of zero gravity, low friction and micro interference of a base satellite; the satellite simulator has a basic single-shaft attitude adjustment and track control function under a condition that a satellite provided by the movement simulator is in a track movement simulating environment, comprises a sensor, a controller and an actuating mechanism (a counteractive flywheel and a cold gas injecting system), and is provided with auxiliary systems such as a power system, a wireless communication assembly, a structural component and the like; and the ground control panel provides a basic function of the wireless data communication between a target satellite and the ground, and realizes the control on the target satellite and the receiving, storage and display of the data.

A propulsion system for the orbital control of a satellite with terrestrial orbit travelling with a speed of displacement along an axis V tangential to the orbit comprises two propulsion assemblies, fixed to the satellite facing one another with respect to the plane of the orbit, each of the propulsion assemblies comprising two propulsion modules; each of the propulsion modules successively comprising: a motorized link for rotation about an axis parallel to the axis V, an offset arm, and a platen supporting a propulsion unit able to deliver a thrust oriented along an axis perpendicular to the axis V; the two propulsion modules of each propulsion assembly being linked to the satellite on either side and substantially at equal distances from a plane perpendicular to the axis V passing through a centre of mass of the satellite.

The invention discloses a satellite autonomous control system for achieving a distributed formation flight. The satellite autonomous control system is composed of six channels, is embedded in satellite controllers, and can generate orbit control commands in advance in an existing satellite management and control system to provide readiness time for engine catalyzing bed heating, attitude maneuver, ground verification and the like. The satellite autonomous control system is characterized in that channel control command sets are stored in RAMs of the satellite controllers, the channel control command sets which are circularly called are stored in a dynamic data flow, and therefore on-satellite resources such as databases and data inquiring are prevented from being occupied. The satellite autonomous control system is embedded in the satellite controllers to be called in the mode of a relatively-independent subprogram, in other words, the management and control system polls the autonomous orbit control progress at each moment. The satellite autonomous control system can serve as supplement of the existing satellite management and control system, an original on-satellite management and control software system is not required to be redesigned, and the satellite autonomous control system can achieve tasks such as formation structure capturing, structure reconstituting and structure keeping of a satellite, and take part in the function that fault satellites in a formation are evacuated.

The invention provides an optimization method for spacecraft iterative guidance, and belongs to the field of space orbit transfer vehicle transfer orbital control. The optimization method comprises the steps of firstly determining an initial boot point and a shutdown point of a main engine; according to the boot point and the shutdown point and a terminal constraint weight factor, using an optimal guidance algorithm to perform iterative guidance simulation calculation; and finally, after a shutdown point condition is satisfied, ending iterative guidance, and obtaining deviation data. When the position deviation of the shutdown point in the X direction exceeds a threshold, the shutdown point is adjusted, so as to reduce the position deviation of the shutdown point in the X direction; and then iterative guidance simulation is performed until the shutdown point satisfies the requirement, and the optimization is over. The optimization method for the spacecraft iterative guidance provided by the invention gets rid of small angle correction hypothesis of traditional iterative guidance, and meanwhile, a terminal constraint under an orbit injection point orbit coordinate system is transformed into an equivalent terminal constraint under a geocentric inertial coordinate system, and appropriate weight adjustment is performed, so that the accuracy of numerical solution and the adaptability of the guidance method are improved, so as to ensure a spacecraft to finally reach the requirement of a task point.

The invention discloses a wireless communication module and a wireless communication satellite using the module, relates to a wireless communication module and a wireless communication satellite, and solves the problems that the conventional satellites have complex structure, poor reliability and poor safety and are not favorable for tests. In the wireless communication module, a radio frequency antenna is connected with a radio frequency signal transmitting / receiving end of a radio frequency chip circuit; a signal end of the radio frequency chip circuit is connected with a signal end of a microcontroller; and an external data acquisition and control signal end of the microcontroller is a baseband data input / output end of the wireless communication module. In the wireless communication satellite, a power supply provides working power for a measurement and control system, a satellite management system, an attitude and orbit control system and an effective load management and control system; and the measurement and control system, the satellite management system, the attitude and orbit control system and the effective load management and control system form a wireless network through four wireless communication modules. The wireless communication module and the wireless communication satellite using the module are applied to the field of high-reliability and high-safety satellite communication.

An ultra-compact aerovortical swirl-enhanced combustion (ASC) system features an aerovortical swirl generator for use in rocket thrusters utilizing hypergolic or non-hypergolic propellants. The ACS thruster can be sized for diameters ranging from about 0.5 to about 2.0 inches, and producing thrust levels of approximately 5 lbf to about 250 lbf. A plurality of helicoid flow channels in the swirl generator introduces swirl into a flow stream of a first propellant within ultra-compact sized rocket thrusters. The ASC system also includes injectors for introducing a second liquid propellant into the swirling flowfield to promote rapid and efficient atomization, mixing and vigorous combustion, which, results in major improvements in combustion and propulsion performance over current rocket thrusters, but in much shorter combustor systems. Hence, the ultra-compact ASC system is a substantial improvement in small bipropellant chemical propulsion thrusters, which can be utilized in-space satellite, spacecraft maneuvering and attitude / orbit control.

The invention provides a satellite attitude track control test system and method. The satellite attitude track control test system comprises a simulator head, a simulator intelligent controller and a simulator remote client side, wherein the simulator head is used for outputting simulation signals in a test and simulating the operating environment of a satellite to be tested; the simulator intelligent controller is connected with the simulator head, controls the operation of the simulator head and collects the work state of the simulator head, and the simulator remote client side is connected with the simulator intelligent controller, controls the operation of the simulator intelligent controller remotely and displays the work state of the simulator head; the simulator remote client side is connected with the simulator intelligent controller through a bus. The satellite attitude track control test system and method can realize a simulator remote control function, carry out remote control and state display through the simulator remote client side, realize unattended operation of a test site, and improve work efficiency.

The invention discloses a spatial non-cooperation target autonomous visual line rendezvous no-model preset performance control method, and the method comprises the following steps: 1), building a non-cooperation target autonomous visual line rendezvous relative orbit motion model in a visual line coordinate system; 2), defining a relative orbit control state quantity, and converting the relative orbit motion model into an affine form; 3), taking the exponential convergence speed, overshoot and steady-state error boundary as the constraints, and building a preset performance boundary function; 4), designing a relative orbit no-model preset performance controller; 5), designing a relative posture no-model preset performance controller. The method achieves the autonomous rendezvous control of a non-cooperation target with spinning and unknown maneuver. The method can achieve the high-precision and high-robustness autonomous rendezvous control of the non-cooperation target without the specific parameters of the model, and is higher in theoretical significance and engineering application value. In addition, the method greatly improves the safety of a rendezvous task of the non-cooperation target.

The present invention relates to a method (50) for controlling the orbit of a satellite (10) in earth orbit, the orbit of the satellite (10) being controlled by commanding, according to a plane of operation, the propulsion means (30, 31), including at least one thruster, and means for moving (20, 21) said propulsion means. According to the invention, said plane of operation includes at least two orbit-control operations, thrust powers of the propulsion means (30, 31) during said two orbit-control operations having respective thrust directions that are not parallel in an inertial frame of reference, each one of said thrust powers being determined so as to simultaneously control the inclination and the position of the orbit of the satellite as well as to form a momentum that is suitable for unloading a device for storing angular momentum of said satellite in a plane orthogonal to the direction of thrust of said thrust power.

A filtration unit for a swimming pool or spa, is disclosed, the filtration unit including a support base (20); a filter housing (30) mounted on top of the support base; and a recess (21) in the support base, the recess adapted to receive a pump (60), at least in part. A multi-port valve (50) is laterally mounted to the filter housing wherein an orbital control arm (54) of the multi-port valve may be moved in a substantially vertical plane to select alternative operations of the filtration unit. The unit may further include a top with space for electronic controls and / or chemicals.

The invention relates to a layout method for a micro attitude and orbit control thruster array. In order to solve the problem that the conventional thruster array is unsuitable for orbit control, the invention provides a regular hexagon-based thruster array and respectively designs a micro-thruster array-based thruster allocation algorithm according to functions to be realized, namely attitude control and orbit control, wherein the array can avoid an eccentric thrust due to the algorithm, and can rapidly and accurately find the needed coordinate position of each thruster by using a distribution rule of the thruster in the array and a related control algorithm. The layout method for the micro attitude and orbit control thruster array has the advantages of improving the utilization rate of the thruster in the array and effectively avoiding ending the service life of the array too early.

The invention discloses a satellite orbit simulation method and device. The method comprises the following steps: acquiring total simulation parameters; according to the total simulation parameters, combining different conditions of numbers of satellites to obtain orbit parameters of satellites to be simulated under an inertial system respectively; when satellite orbit transfer orbital control requirements exist, realizing incremental computation of a speed needed by satellite transfer orbital control through calling an orbit optimization program, so as to calculate a total acceleration of the satellites; and finally, calling an orbit dynamic calculation program to finish satellite orbit simulation. Aiming at the satellites with the different numbers, the satellite orbit simulation is carried out and an applicable range of a simulation program is wide; and furthermore, the satellite transfer orbital control can be realized through calling the orbit optimization program, and the simulation program is open so that development and expansion are facilitated.

The invention discloses a track control method for a space rendezvous ground guide section with limited orbit maneuvering, aiming to realize high-precision track control of the space rendezvous ground guide section under the condition of limited orbit maneuver. The technical solution is to first determine the orbital maneuvering scheme of the tracking spacecraft as five orbital maneuvering schemes, determine the orbital control parameters M1~M5 that need to be calculated; calculate the orbital control parameter M1 after the tracking spacecraft enters orbit, and upload it to the tracking spacecraft. The maneuvering system performs the first orbital maneuver; after the first, second, and third orbital maneuvers are completed, the orbital control parameters M2, M3, and M4 are calculated respectively, and uploaded to the tracking spacecraft, and the orbital maneuvering system performs the second, third, and fourth maneuvers Orbital maneuvering; finally implement the fifth orbital control according to the orbital control parameter M5 (obtained when calculating M4). The invention has less number of track maneuvers, lower requirements on the attitude control system, improves the track control accuracy of the ground guidance section under the condition of limited track maneuvers, and has fast calculation speed and good real-time performance.

The invention provides a satellite on-orbit transient temperature simulation and prediction system which comprises a thermal control module, an orbit and attitude control module, a hardware interface module, a software interface module and a satellite thermal model module. The thermal control module is used for simulating the satellite thermal control management function. The orbit and attitude control module is used for simulating the satellite on-orbit working condition function. The hardware interface module is used for achieving conversion between data of the thermal control module and data of the software interface module and conversion between data of the orbit and attitude control module and the data of the software interface module. The software interface module is used for achieving protocol conversion for data communication with the satellite thermal model module. The satellite thermal model module calculates and analyzes the satellite thermal state according to the current satellite operation state and outputs a result. According to the system, data such as a satellite on-orbit operation orbit, the attitude of a satellite, the device working condition and a heater working condition can be received in real time, the thermal state of satellite on-orbit operation is simulated, temperature distribution in the satellite on-orbit operation process is predicted, and the thermal design function and the active thermal control functions of the satellite are verified.

The invention provides a line vibration measurement-based satellite engine on-orbit thrust real-time calibration method. According to the line vibration measurement-based satellite engine on-orbit thrust real-time calibration method, ground line vibration test data are adopted to correct a response simulation finite element model, and a transfer function between thrusts and line vibration is calibrated; interference elimination is performed on the basis of simulation result data and on-orbit test data, so that line vibration response caused by an engine on-orbit thrust can be obtained; engineon-orbit thrust inversion is carried out through using the transfer function between the thrusts and the line vibration; and the engine on-orbit thrust is calibrated in real time. The method of the invention can be widely applied to the real-time calibration of the on-orbit thrusts of various types of engines / thrusters of various spacecrafts and can facilitate the strategy correction and optimization of orbital control and attitude control which are performed by spacecrafts in an on-orbit manner.

The invention discloses an eccentricity ratio freezing common-rail double-pulse control intermediate orbit determination method. The method is specifically implemented according to the following steps: 1, determining a pre-control satellite orbit at an intermediate moment of first orbital transfer; 2, calculating the position and speed of a satellite when the orbit is not changed at the intermediate moment of the first orbit change; 3, determining a satellite orbit after secondary orbit transfer intermediate time control; 4, confirming the control type, the control quantity and the control phase of a double-pulse track; 5, respectively calculating a track eccentricity ratio variable quantity, a track flat semi-major axis variable quantity, a first track control actual speed increment and asecond track control actual speed increment; 6, calculating a satellite speed vector after the first orbital transfer intermediate time control; 7, determining a double-pulse control middle arc section track. By means of the method, the available working arc section of the orbit in the spacecraft control period can be improved, and certain economic benefits are achieved for on-orbit operation ofthe spacecraft.

The invention discloses a high-integration three-unit cubic satellite capable of maneuvering orbital transfer. The cubic satellite comprises a satellite basic platform, and the satellite basic platform comprises a main force bearing structure, a solar cell array, a satellite computer, a lithium battery pack, a power supply control panel, a UV antenna, a UV communication machine, an attitude measurement and control system, other related accessories and a propulsion and control system; and the control system comprises a double-pulse solid rocket thruster, an ignition control panel of the double-pulse solid rocket thruster, a cold air micro thruster, a pulse modulation module of the cold air micro thruster, a mass moment executing mechanism and a mass center adjusting mechanism. The three-unit cubic satellite has the thrust vector control capability and is high in integration degree, rapid orbital transfer and attitude orbit control of the cubic satellite can be conducted, large-range orbital maneuver of the satellite is achieved, and the application field of the cubic satellite is greatly expanded.

Provided is a system adopting a mobile project robot to drill holes. The system adopting the mobile project robot to drill the holes comprises a BIM tunnel hole drilling model, a CPIII railway control net, and the project robot, wherein the CPIII pathway control network comprises a control point location column. The project robot comprises a pathway mobile platform, a power system, a industrial computer, six axles mechanical hand, a PLC controlling system, a positioning system, a photograph system, a rebar detecting instrument, a real time laser hole depth detector, a electric hammer and a dust-precipitating system are installed on the railway mobile platform. According to the system adopting the mobile project robot to conduct tunnel drilling, the drilling quality is high, the consistency is good, the system is efficient, reliable, safe and fast, and the human cost is saved, at the same working hours, tunnel drilling working distance can approach the several times of the artificial drilling holes.

The invention discloses an intelligent lunar soft landing orbit controller. The orbit controller comprises a simulation module, an optimizing module and an execution module, wherein in the simulation module, the lunar soft landing orbit is divided into a plurality of segments, a thrust direction angle changing curve of the orbit is fitted by aid of thrust direction angles of nodes at segmentation and corresponding moments, and then the whole soft landing orbit is obtained through simulation of a soft landing dynamic model. The optimal thrust direction angles of the nodes and the optimal soft landing terminal moment are obtained through optimizing of the optimizing module. The optimizing module adopts an improved intelligent algorithm, the improved intelligent optimizing method controls the size of subgroups, disturbance factors are added, searching variety is improved, and the optimizing module has very high searching capacity and can search a lunar soft landing orbit requiring the least fuel consumption through optimizing. An optimal soft landing command signal is transmitted to the execution module, and the optimal lunar soft landing control requiring the least fuel consumption is realized finally.

The invention discloses a formation configuration reconstruction optimization using a finite element method. The finite element method is used for dividing the formation reconstruction time into a series of equally spaced intervals, and the problem of solving the reconstruction path is transformed into an optimization problem; therefore, the limitation problem of the traditional reconstruction method is overcome, and the reconstruction of a satellite formation is realized. In the optimization process, the optimal orbit control quantity and the corresponding reconstructed orbit state quantity in each time subinterval are determined according to the numerical iterative method; and the iteration and correction are performed again on the loading of the control acceleration applied by a small thrust engine to the spacecraft in the optional subinterval and the reconstruction orbit state quantity considering the criteria of controlling the total fuel consumption to be minimum, the control opportunity, and the safety factor to avoid collision among spacecrafts.

The invention provides a multi-stage health state acquisition method for a low-orbit remote sensing satellite navigation positioning system. The multi-level health management comprises: equipment-level health management and system-level health management; wherein the equipment-level health management comprises equipment operation health management and equipment function health management; the system-level health management comprises: a navigation positioning system operation health management and navigation positioning system function health management. The implementation body of the system-level health management comprises a satellite-borne computer, an orbit control subsystem and a payload subsystem, wherein information interaction exists between the satellite-borne computer and a navigation and positioning system in a low-orbit remote sensing satellite. The invention provides a multistage health management method for a low-orbit remote sensing satellite navigation positioning system. Compared with single-level health management, the management means are more diverse, the health state is more comprehensively reflected, the limitation of fault handling in single-level health management is solved, and on-orbit autonomous safe operation of the low-orbit remote sensing satellite can be more powerfully guaranteed.

The invention relates to a satellite scanning system and discloses a space scanning bracket with a self-balancing function. The space scanning bracket with the self-balancing function comprises an azimuth scanning bracket (5) which is arranged on an arc-shaped bracket (4), the azimuth scanning bracket is provided with two sliding rails and two sliding blocks, and the arc-shaped bracket (4) is provided with an internal arc and an arc-shaped sliding rail. A loading instrument (7) and a balance weight (1) are arranged on two sides of the azimuth scanning bracket (5) and do oscillatory scanning with speed in the opposite direction on the sliding rails of the azimuth scanning bracket (5) through rollers, the product of rotation inertia and rotation speed of the loading instrument encircling a rotation center and the product of rotation inertia and rotation speed of the balance weight encircling the rotation center are equal in number and opposite in direction. The space scanning bracket with the self-balancing function solves the problems of a complex control system, a high manufacturing cost and unreasonable structure design and the like existing in an existing scanning mechanism and achieves the beneficial effects of relieving the burden of the satellite attitude and the orbital control system, reducing the development cost, prolonging the satellite service life and the like.

The invention discloses an orbit maintaining method for an ultra-low orbit satellite. The method comprises the steps: setting satellite parameters and atmospheric parameters, building an orbit dynamics model according to the satellite parameters and the atmospheric parameters, and enabling the satellite parameters to comprise satellite orbit parameters, satellite basic parameters and initial satellite orbit control parameters; an optimization model is constructed, the initial satellite orbit control parameters are optimized according to the optimization model, the orbit dynamics model, the satellite orbit parameters and the satellite basic parameters, optimized orbit control parameters are obtained, and the optimization model comprises a plurality of optimization constraint conditions; According to the optimized orbit control parameters, an air suction type propulsion system is controlled to collect air and exhaust the air, so that orbit maintenance of the ultra-low orbit satellite is realized. The technical problem that in the prior art, an ultra-low earth orbit satellite is difficult to maintain in orbit for a long time is solved.

The invention discloses a satellite attitude control method based on variable control periods. The satellite attitude control method includes the steps of S1, selecting corresponding control periods and controller parameters according task requirements; S2, using original data collected by a gyro to calculate inertial angular speed according to the control periods; S3, using attitude information output by an attitude sensor to calculate a satellite attitude determined angle; S4, using the inertia angular speed obtained in S2 and the attitude determined angle obtained in S3 to calculate controlmoment; S5, using the control moment to transmit control pulse width to an actuator; S6, returning to S1 to repeat a new round of satellite attitude control to achieve variable-period satellite attitude control. The satellite attitude control method has the advantages that different control periods and controller parameters are selected according to different task requirements, restriction of thecalculation quantity of attitude control tasks and computer hardware conditions is broken through, and high-precision transfer orbital control and high-precision, long-term and reliable earth observation tasks are completed.