31results about How to "Extended life on orbit" patented technology

The invention discloses an automatic management system for a lithium ion storage battery pack power supply sub system for a satellite. The automatic management system comprises an in-orbit management module, an emergency protection module and a lithium ion storage battery pack in-orbit different stage management mode method; the in-orbit management module is used for preventing the lithium ion storage battery pack from being over-charged; the emergency protection module is used for preventing the lithium ion storage battery pack from being over-discharged; and the use and management mode method for the lithium ion storage battery pack in-orbit different stages is used for prolonging the in-orbit running service life of the lithium ion storage battery pack. The in-orbit management and control for the satellite lithium ion storage battery pack power supply sub system is realized, so that the reliability of the system is greatly improved fundamentally.

The invention discloses a method and system for on-orbit correction of a track control strategy based on thruster deviation estimation. The method comprises the following steps: respectively acquiringthe position information of the center of mass of a whole satellite, a first orbit control thruster and a second orbit control thruster relative to a layout system; acquiring thrust vector directionsof the first orbit control thruster and the second orbit control thruster; carrying out on-orbit identification, and determining moments generated by the first orbit control thruster and the second orbit control thruster according to an on-orbit identification result; constructing an interference torque evaluation function; solving the interference torque evaluation function, determining the minimum value of the interference torque evaluation function, and calculating to obtain jet pulse width; and performing orbit control according to the calculated jet pulse width. According to the invention, an orbit control thrust distribution strategy is updated on orbit by adopting an energy optimization principle, and the minimization of orbit control interference is realized.

The invention discloses a method for determining mounting parameters of orbit control engines of satellites. The method includes determining a mounting position of each engine by the aid of included angles alpha<E>, beta <E> and gamma <E> among a normal line O<m>X<m> of an accurate measurement mirror of the engine and an axis O<sc>X<sc>, an axis O<sc>Y<sc> and an axis O<sc>Z<sc> of a mechanical coordinate system of the corresponding satellite and coordinates x<E>, y<E> and z<E> of the center O<m> of the accurate measurement mirror of the engine on the axis O<sc>X<sc>, the axis O<sc>Y<sc> and the axis O<sc>Z<sc> of the mechanical coordinate system of the satellite. The method has the advantages that change conditions of mass centers of the satellites and thrust vectoring thermal marking results of the engines are comprehensively considered, so that the mounting parameters of the engines can be optimally determined, disturbance torque generated in orbital transfer periods of the satellites can be reduced, satellite propellants can be saved, and the in-orbit lives of the satellites can be prolonged; thrust of each engine points to the average mass center of the corresponding satellite in the corresponding orbit transfer period, so that counterweights for the satellites can be reduced or can be omitted, and the effective load weight of each satellite can be increased.

The invention provides a satellite orbit control engine installation optimization method based on a genetic algorithm. The method comprises the following steps that (1) an objective function of the engine installation optimization is determined according to the requirement of a control system on the satellite layout; (2) according to the engine geometrical parameters and the installation requirements, a constraint function of the engine installation optimization is determined; (3) an AMTLB genetic algorithm toolkit is used for carrying out engine installation parameter optimization. The satellite orbit control engine installation optimization method has the advantages that the interference moment in the orbital transfer period is further reduced, so the satellite fuel is saved, and the satellite on-orbit life is prolonged.

The invention provides a spacecraft orbit maneuver motor installing parameter optimizing method based on a Runge-Kutta algorithm. According to the method, firstly, each coordinate system is defined; then, a mathematical model between undetermined coefficients and a vector from a theoretical circular center A of a motor installing flange to a theoretical circular center C of the motor installing flange is built; a vector from the circle center A to a target point T, a vector from O<s> to the circle center A, a motor thrust vector F<EB>, an acting point position vector, a vector from the circle center A to the target point T, a vector from the circle center A to the motor thrust vector acting point P, a motor thrust vector and a vector from the motor thrust vector acting point P to the target point T are sequentially calculated; and after the undetermined coefficients are solved, the installing angle and the installing position are obtained. The method has the advantages that the optimum design mathematical model is built; and on the premise that a solution meets the constraint condition, the optimum values of motor installing parameters meet the requirements of the spacecraft orbit maneuver motor installing parameter optimization to the maximum degree.

The invention discloses a low-thrust long-term position keeping method for a geostationary orbit satellite. The method comprises the following steps establishing an in-plane and out-plane average orbital motion model of a satellite orbit through spherical coordinates; giving a long-period movement rule of the geostationary orbit satellite in the plane and out of the plane through a phase plane analysis method; on the basis, selecting a fixed-point position keeping window of the geostationary orbit satellite, and acquiring a periodic motion trail, namely a drifting section trail, of the satellite in an uncontrolled state in the fixed-point window; and designing a small thrust control law to obtain a motion trail, namely a thrust section trail, of the satellite in a controlled state so thatthe thrust section trail and the drift section trail form a closed-loop trail together, and small-thrust long-term position maintenance of the geostationary orbit satellite is completed.

The invention discloses a water-based propulsion system and method applied to a high orbit satellite. The water-based propulsion system comprises a power source, a water tank, a static water supply proton exchange membrane electrolysis tank, a hydrogen and oxygen management system, a hecto-Newton scale gaseous hydrogen gaseous oxygen track control engine, and a Newton scale gaseous hydrogen gaseous oxygen attitude control engine. According to the water-based propulsion system and method applied to the high orbit satellite, the problems that hydrogen and oxygen gases are difficult to be managedon a long-term basis in the space environment and that the hydrogen and oxygen gases with an equivalent mixing ratio are difficult to be efficiently utilized are solved, and engineering practicability of a hydrogen and oxygen propulsion system in a space flight platform is improved.

The invention discloses an overcharge protection device applied to an S4R-type circuit. The overcharge protection device comprises an S4R charge module, a discharge adjusting module, a balance module and a lower computer module, wherein the S4R charge module carries out voltage control of a whole lithium battery pack, maximal voltage control of single lithium batteries and overcharge current control; meanwhile, the charge states of the lithium batteries are monitored and controlled; control parameters can be changed through the lower computer module; a BOOST mode is adopted by the discharge adjusting module; meanwhile, the discharge adjusting module comprises a busbar voltage overvoltage and undervoltage protection circuit and an output current limiting protection circuit; the balance module provides sampling and power-balancing loops and logic control interfaces of the single lithium batteries and prevents overvoltage of the single lithium batteries; logic control of the balance module is achieved through the lower computer module; the lower computer module comprises a lithium battery temperature control module and a lithium battery charge and discharge management module; and an input signal required by a lower computer is provided by an analogue collection module.

The present invention provides a spacecraft orbit maneuver motor installation parameter optimization method. The method comprises the steps of: firstly, defining each coordinate system, and establishing a relation among a motor thrust vector, an acting point position vector and an undetermined coefficient; then calculating a relation between a disturbance torque value during each ignition; and with an optimization goal of making a maximum and a minimum of a disturbance torque component during each ignition equal in numerical value but opposite in sign, solving the undetermined coefficient, so as to obtain an installation angle of a spacecraft orbit maneuver motor and an installation position of a motor nozzle. According to the method provided by the present invention, an optimally designed mathematical equation is constructed, and the demand for optimization of the installation parameter of the spacecraft orbit maneuver motor is satisfied to the largest extent.

The invention provides a modular attitude control unit for a satellite. The modular attitude control unit comprises a shell, a momentum wheel and a magnetic torquer, wherein the surface of the shell comprises at least one standard interface; the modular attitude control unit is connected to other functional modules of the satellite through the standard interface; the momentum wheel and the magnetic torquer are arranged in the shell; the momentum wheel is used for controlling the attitude of the satellite to maintain the angular momentum of the system; and the magnetic torquer is used for performing attitude control or momentum wheel unloading management on the satellite.

The invention provides a space laser structure. The space laser structure comprises a laser shell, a cover plate, supporting legs and a laser assembly, wherein the laser shell and the cover plate form a sealing structure; the supporting legs are arranged around the laser shell and are used for fixing the laser shell and a satellite main body; the laser shell and the cover plate are sealed through an O-shaped ring; and the laser shell and the laser assembly are provided with a protective layer, the laser assembly comprises components and an isolation assembly used for improving mechanical stability, part of the components are directly fixed on the inner wall of the laser shell, and the other part of the components are fixed with the laser shell through the isolation assembly. The space laser structure in a compact space is provided by theinvention,reduces the impact on the structural deformation of a resonant cavity from a space environment and the internal pollution of a laser housing through employing a high-strength and stable-performance material, the isolation of the resonant cavity, the power supply of a bus bar, and the double-O-ring sealing, and ensures the long service life and high reliability of the space laser.

The present invention provides an on-orbit performance adjustment and compensation method and system of a spacecraft propulsion system, the method comprising: step 1: setting multiple pressure sensors and multiple parallel orbit control self-locking valves on the orbit control pipeline; step 2: Through the stand-alone liquid flow debugging of the rail-controlled self-locking valve, the flow resistance value of a single rail-controlled self-locking valve is obtained respectively; Step 3: Make the ground adjustment value of the inlet pressure of the oxidant path and the fuel path equal when the rail-controlled engine is working; Step 4: The rail-controlled self-locking valve When debugging the liquid flow of a single-machine control engine, ensure that the accuracy of the engine mixing ratio meets the requirements; Step 5: During the orbital flight of the propulsion system, obtain the actual measured value of the engine inlet pressure during the ignition of the orbit-controlled engine; Step 6: Calculate the inlet pressure of the oxidizer circuit and Compared with the actual measured value of the inlet pressure of the fuel line, the on-orbit performance adjustment and compensation strategy is implemented as needed. The invention can actively adjust and compensate the mixing ratio of the orbit control engine within a certain range, and improve the utilization rate of the propellant.

The invention discloses a virtual cathode arc propeller adopting dielectric barrier discharge. The virtual cathode arc propeller adopting dielectric barrier discharge comprises a metal sleeve, a sealing element, an intake port, a cathode, and an anode with a spray port, wherein the cathode is a virtual cathode; the virtual cathode is a metal cylinder with a length; and a glass dielectric is fixedto one end of the virtual cathode in a sintering manner. According to the virtual cathode arc propeller adopting dielectric barrier discharge, the glass dielectric is fixed to the top end of the cylindrical virtual cathode in the sintering manner, so that an ablation effect of the arc propeller on the cathode and the anode in a working process can be avoided, the on-orbit life of the propeller isgreatly prolonged, and the practicability of the arc propeller is improved.

The invention discloses an orbit control strategy on-orbit correction method and system based on thruster deviation estimation. The method includes: separately obtaining the positions of the center of mass of the entire satellite, the first orbital thruster and the second orbital thruster relative to the layout system information; and, obtain the thrust vector directions of the first orbital control thruster and the second orbital control thruster; perform on-orbit identification, and determine the generation of the first orbital control thruster and the second orbital control thruster according to the on-orbit identification results The torque; construct the disturbance torque evaluation function; solve the disturbance torque evaluation function, determine the minimum value of the disturbance torque evaluation function, and calculate the jet pulse width; perform orbit control according to the calculated jet pulse width. The invention adopts the energy optimal principle to update the orbit control thrust distribution strategy on the orbit, and realizes the minimization of the orbit control interference.

The invention discloses a water-based propulsion system and method applied to high-orbit satellites, wherein the system includes a power supply, a water tank, a static water supply proton exchange membrane electrolytic cell, a hydrogen-oxygen management system, and a 100-new level gas-hydrogen-oxygen rail control engine, cow-level gas hydrogen oxygen attitude control engine. The invention solves the problem that the hydrogen-oxygen gas is difficult to manage for a long time in the space environment and the problem that the equivalent mixing ratio of the hydrogen-oxygen gas is difficult to efficiently utilize, and improves the engineering practicability of the hydrogen-oxygen propulsion system on the space flight platform.

The embodiment of the invention provides a grid flicker safety protection method of an ion electric propulsion system. The method comprises the following steps that (1), the current grid current Ib of the ion electric propulsion system is monitored; and (2), if the current grid current Ib is greater than a first set threshold value, it is judged that the current grid current Ib is an overcurrent signal, when the duration of the overcurrent signal is less than or equal to the time t0, the step (1) is executed, when the duration of the overcurrent signal is greater than the time t1, the operation of adjusting the grid voltage and the grid current to be 0 in the step (3) is executed, and when the duration of the overcurrent signal is greater than the time t0 and less than or equal to the time t1, the judgment and the operation in the step (4) are executed. By utilizing the technical scheme provided by the embodiment, safety protection can be provided for the ion electric propulsion system on the premise of avoiding frequent and repeated starting, and on-orbit safe operation of a satellite is guaranteed.

The invention discloses a dielectric barrier discharge virtual cathode arc thruster, which includes a metal sleeve, a sealing member, an air inlet, a cathode, an anode with a nozzle, the cathode is a virtual cathode, the virtual cathode is a metal cylinder with a certain length, and the virtual cathode One end is fixed with a glass medium by sintering. The invention sinters and fixes the glass medium on the top of the virtual cathode of the cylinder, which can avoid the ablation effect of the arc propeller on the cathode and anode during the working process, greatly prolong the on-orbit life of the propeller, and improve the practicability of the arc propeller.

The invention discloses a separable satellite propulsion system configuration. A propulsion system sends a satellite into a higher orbit and is separable from the satellite after the satellite entersan orbit, so that the problems of 'dead weight' of the satellite, reduction of the in-orbit position guarantee efficiency of the satellite and consumption of precious fuel resources of the satellite are solved. The invention provides the separable satellite propulsion system configuration which can be used for sending a satellite to a specified orbit. The separable satellite propulsion system configuration comprises a force bearing cylinder, a connecting and separating device and a boosting control structure; the force bearing cylinder is used for bearing the satellite; the connecting and separating device is arranged at the top end of the force bearing cylinder for connecting the satellite; and the boosting control structure is connected with the connecting and separating device and is used for propelling the satellite borne by the force bearing cylinder to a specified orbit and separating from the connecting and separating device to separate from the satellite after propelling the satellite to the specified orbit.

The invention provides a pluggable satellite thermal control system, which includes an operating robot, a heat insulation component and a heat dissipation component. The operating robot is connected to the surface of the satellite main body for clamping and installing the heat insulation component and the heat dissipation component; the heat insulation component is arranged on the sun receiving surface of the satellite main body, and is connected to the surface of the satellite main body through a standard interface; and the heat dissipation component is arranged on the The back side of the satellite body is connected to the surface of the satellite body through a standard interface.

The invention provides a control method for constant-current constant-voltage charging of a lithium storage battery. The method includes the steps that work parameters such as voltage, currents and temperature are collected to judge whether a power source system is in the working state of charging the storage battery or not; if the power source system is in the charging state, whether the working temperature of the storage battery is normal or not is judged, and if the working temperature exceeds the limit, the upper limit of charging currents of a constant-current stage is revised as an allowable value at the current temperature; according to monomer voltage of the storage battery, whether the storage battery works in the constant-current stage or a constant-voltage stage is judged; if the storage battery works in the constant-current stage, control quantity is given according to the upper limit of the charging currents of the previous constant-current stage; if the storage battery works in the constant-voltage stage, control quantity needing to be given currently is calculated; when the lithium storage battery is fully charged, the charging currents are set to be zero to stop charging. The constant-current constant-voltage control method is more flexible than a traditional method, and the on-orbit service life of the storage battery is prolonged easily.

The invention discloses a modular solar sail assembled and reconfigurable space-based debris removal platform. The modular solar sail assembled and reconfigurable space-based debris removal platform can change the status of a single-scale debris removal mission, and cover a plurality of removing missions such as large-scale space debris and abandoned satellites over ten centimeters in a high orbit, small satellites such as low-orbit failure cubic satellites and a large number of low-orbit millimeter-class micro-space debris through flexible on-orbit reconfiguration. According to specific debris removal tasks, the required sail area is determined, the number of cell modules is determined by the area, a certain number of the cell modules are released from the platform for on-orbit connectingreconfiguration and perform a corresponding task after on-orbit unfolding.

The invention discloses a method for determining mounting parameters of orbit control engines of satellites. The method includes determining a mounting position of each engine by the aid of included angles alpha<E>, beta <E> and gamma <E> among a normal line O<m>X<m> of an accurate measurement mirror of the engine and an axis O<sc>X<sc>, an axis O<sc>Y<sc> and an axis O<sc>Z<sc> of a mechanical coordinate system of the corresponding satellite and coordinates x<E>, y<E> and z<E> of the center O<m> of the accurate measurement mirror of the engine on the axis O<sc>X<sc>, the axis O<sc>Y<sc> and the axis O<sc>Z<sc> of the mechanical coordinate system of the satellite. The method has the advantages that change conditions of mass centers of the satellites and thrust vectoring thermal marking results of the engines are comprehensively considered, so that the mounting parameters of the engines can be optimally determined, disturbance torque generated in orbital transfer periods of the satellites can be reduced, satellite propellants can be saved, and the in-orbit lives of the satellites can be prolonged; thrust of each engine points to the average mass center of the corresponding satellite in the corresponding orbit transfer period, so that counterweights for the satellites can be reduced or can be omitted, and the effective load weight of each satellite can be increased.

Because an agile satellite can rapidly maneuver along a rolling axis, a pitching axis and a yawing axis, imaging of a strip target, such as an irregular strip target, in any direction of the earth surface can be realized. The irregular strip target is a long-line target composed of a series of continuous points, and the points are connected through straight lines. For an irregular strip target, inorder to reduce the attitude adjustment frequency among satellite imaging tasks, the irregular strip target can be divided into a plurality of rectangular strips for observation by utilizing the characteristic that a camera has a certain imaging breadth. Based on the technical scheme, the invention provides the strip decomposition method for the imaging task of the irregular strip target. Throughverification, the irregular strip decomposition method provided by the invention is feasible, and can effectively reduce the attitude adjustment frequency among satellite imaging tasks.

The invention provides an on-orbit performance adjusting and compensating method and system for a spacecraft propulsion system. The method comprises the steps of 1, arranging a plurality of pressure sensors and a plurality of parallel on-orbit control self-locking valves on an on-orbit control pipeline; 2, obtaining the flow resistance value of each single rail control self-locking valve through single-machine liquid flow debugging of the rail control self-locking valve; 3, enabling the ground adjusting values of the inlet pressures of the oxidant path and the fuel path to be equal when the orbit control engine works; 4, ensuring that the mixing ratio precision of the engine meets the requirement during single-machine liquid flow debugging of the orbit control engine; 5, in the on-orbit flight process of the propulsion system, obtaining the measured value of the engine inlet pressure in the ignition working process of the orbit control engine; and 6, comparing the measured values of the oxidant path inlet pressure and the fuel path inlet pressure, and executing an on-orbit performance adjustment compensation strategy as required. Active adjustment compensation control within a certain range can be conducted on the mixing ratio of the orbit control engine, and the propellant utilization rate is increased.

The invention provides a pluggable satellite thermal control system. The pluggable satellite thermal control system comprises an operation robot, a heat insulation assembly and a heat dissipation assembly. The operation robot is connected with the surface of a satellite body and used for clamping and installing the heat insulation assembly and the heat dissipation assembly. The heat insulation assembly is arranged on the sunny side of the satellite body and connected with the surface of the satellite body through a standard interface; and the heat dissipation assembly is arranged on the nightside of the satellite body and is connected with the surface of the satellite body through a standard interface.

The method for optimizing the installation parameters of spacecraft orbit-changing engines based on the Runge-Kutta algorithm firstly defines each coordinate system, and then establishes the mathematical model between the undetermined coefficients and the theoretical circle center A of the engine installation flange to the theoretical circle center C of the engine installation flange, Calculate the vector from the center A to the target point T, the vector from OS to the center A, the engine thrust vector FEB and the position vector of the action point, the vector from the center A to the target point T, the vector from the center A to the engine thrust vector action point P, and the engine thrust vector Thrust vector and the vector from the engine thrust vector action point P to the target point T, after solving the undetermined coefficient, obtain the installation angle and installation position, the present invention constructs the mathematical model of optimal design, and solves the engine under the premise of satisfying the constraint conditions. The optimal value of the installation parameters satisfies the requirements of the optimization of the installation parameters of the spacecraft orbit change engine to the greatest extent.

The invention discloses a powder rotating detonation space engine. The powder rotating detonation space engine comprises a rotating detonation combustion chamber, a spray pipe, a powder fuel supply device, an oxidizing agent supply device and an ignition device; the rotating detonation combustion chamber comprises a left disc outer shell, a left disc inner shell, a right disc inner shell and a right disc outer shell which are sequentially and coaxially arranged from left to right; a sealed disc-shaped combustion cavity is formed between the left disc inner shell and the right disc inner shell; a fuel premixing cavity is formed between the left disc outer shell and a left disc; an oxidizing agent premixing cavity is formed between the right disc outer shell and the right disc inner shell; the powder fuel supply device comprises a plurality of powder storage tanks and a supercharging device; and each powder storage tank can be used for adding micron-sized powder fuel into the fuel premixing cavity under the supercharging action of the supercharging device. According to the powder rotating detonation space engine, micron-scale powder fuel detonation combustion is adopted for achieving propulsion, the powder rotating detonation space engine can adapt to severe external environments, and the axial size of the detonation combustion chamber can be greatly reduced so that the powder rotating detonation space engine can adapt to installation of spacecrafts, such as satellites and the like.