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

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 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 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 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 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 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.