73 results about "Orbital maneuver" patented technology

The invention provides a method and a system for acquiring two rows of elements by a numerical orbit recursion device. The method comprises the following steps of: inputting a two-row element template; determining orbit time reference points corresponding to the two rows of elements; according to a set numerical value recursion device model, a numerical value recursion device obtains a position value sequence of the track in the set space coordinate system within a period of time; converting the position value sequence obtained by the numerical recursion device into a true equatorial mean equinox coordinate system; obtaining two rows of elements by using a single-point method through the two rows of element templates; taking the two rows of elements obtained by the single-point method as an initial orbit, and solving an orbit parameter estimation value in the two rows of elements under an optimal criterion by utilizing a least square method; calculating an orbit residual error sequence corresponding to an orbit parameter estimation value in the two rows of elements; and fitting the residual error by using a polynomial model to obtain a compensation item. The method is easy to implement, the numerical orbit recursion device can be effectively applied to be converted into two rows of elements, and the method can be compatible with time periods with orbit maneuvering.

The invention provides a space target orbit maneuver identification method and system based on adaptive LSM, and the method comprises the steps: obtaining the orbit data of a space target, calculating the six orbit elements based on the orbit data, and coding the six orbit elements into a pulse sequence; and inputting the pulse sequence into an improved adaptive LSM for detection to obtain an orbit maneuver behavior identification result of the space target. According to the method provided by the invention, six orbital elements are calculated through the speed and the position of the orbital, then the orbital parameters are input into an improved LSM, and the improved LSM adopts an adaptive threshold neuron as a calculation unit, so that the threshold can be dynamically adjusted according to the current pulse response and attenuated along with time; meanwhile, an improved particle swarm optimization algorithm is combined to carry out optimization setting on an initial value of a neuron, so that the initial performance of the model is improved; and a learnable neuron connection weight based on an attention mechanism is introduced, so that the accuracy and reliability of orbit prediction are improved.

The invention relates to a spacecraft maneuver detection method based on TLE data and kernel density estimation. The invention relates to the technical field of spacecraft maneuver detection. The method comprises the following steps: acquiring and inputting satellite historical TLE data; gaussian filtering denoising is carried out on the input data; based on the de-noised data, propagation forecasting is carried out through an SGP4 model; constructing a joint residual sequence, and carrying out error distribution modeling and sample scoring; determining a maneuvering threshold value, and judging to obtain abnormal data; and aggregating abnormal data and carrying out maneuvering identification. The method can still realize accurate and stable track maneuver automatic detection under the constraint conditions of unstable TLE data, complex error distribution, no external auxiliary data and the like. The method has good adaptability, robustness and interpretability, and is a track anomaly detection technical scheme which is complete in structure, rigorous in logic and capable of achieving engineering landing.

The invention relates to an autonomous measurement and control orbit determination method and system for a high-orbit satellite. The method comprises the following steps: acquiring satellite-borne GNSS broadcast ephemeris data and single-frequency pseudo-range observation data of a high-orbit satellite; based on the GNSS broadcast ephemeris data and the single-frequency pseudo-range observation data, coarse orbit data of the high-orbit satellite are generated through a dynamic pseudo-range orbit determination algorithm; performing outlier elimination preprocessing on the coarse orbit data to obtain orbit data after outlier elimination; performing smooth fitting and equal-interval sampling on the orbit data after outlier elimination to obtain smooth orbit data; determining an orbit determination arc section according to the scheduling time and the orbit maneuvering time; and using the motion equation and the state transition matrix of the high-orbit satellite to process least square iteration in batches to obtain optimal orbit data, and executing orbit forecasting. The method provided by the invention does not depend on information transmission of ground measurement and control resources, realizes on-satellite autonomous orbit determination, reduces the complexity of data fusion calculation, and is small in calculation amount.

Method, electronic device, system, and computer-readable medium embodiments are disclosed. Some embodiments include a signal processing workflow incorporating a graphical user interface for displaying orbital information for satellites and other spacecraft. In some embodiments, a generative adversarial network (GAN) is employed for evaluating satellite orbital positions, for predicting future orbital movements, for detecting orbital maneuvers of a satellite, and for analyzing such maneuvers for potential nefarious intent.

The invention relates to the technical field of constellation satellite maintenance regulation and control, and discloses an intelligent regulation and control system and method for low-orbit Walker constellation configuration maintenance, and the method comprises the steps: obtaining the change trend of an in-plane phase angle and an inter-plane distance of an orbit of a satellite in real time through a high-precision orbit prediction model; inputting the in-plane phase angle and the inter-plane spacing into a pre-trained decision model, judging whether the parameters exceed a preset configuration maintenance threshold, and positioning an unstable satellite; if the parameters exceed the limit, the model outputs the optimal orbit maneuvering parameters of the satellite based on dual-target constraints of minimum fuel consumption and shortest service interruption; a phase maintenance demand and an on-satellite observation task are fused, an optimal control scheme is generated, and constellation configuration is maintained in combination with task planning and implementation; continuously monitoring the satellite constellation configuration state, and maintaining the stable operation of the constellation. According to the invention, automatic maintenance of Walker constellation configuration is realized, and the control precision and efficiency are improved.

The present invention relates to a method for calculating a long-distance rendezvous orbit maneuver based on multiple laps of latitude arguments, comprising: constructing a long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude arguments; performing a perturbation orbit multi-pulse maneuver calculation based on multiple laps of latitude arguments for given maneuvering parameters to obtain the relative position and velocity of the long-distance rendezvous terminal; using the multi-lap maneuvering point latitude arguments and maneuvering pulses as design variables, and the terminal's relative position and velocity as constraints, employing a sequential quadratic programming algorithm to stably solve the long-distance rendezvous orbit maneuver scheme based on multiple laps of latitude arguments, thereby obtaining the maneuvering pulse design variable values, the number of laps, and the maneuvering point latitude arguments for the given number of laps. Accordingly, the present invention proposes a long-distance rendezvous orbit maneuver calculation system. The present invention overcomes the convergence instability problem caused by the fixed number of laps in traditional long-distance rendezvous orbit maneuver calculation technology, and achieves stable calculation of long-distance rendezvous orbit maneuvers by performing a cross-lap search using multiple laps of latitude arguments.

The invention discloses a method and a system for planning an orbital task of a maneuvering satellite, and belongs to the field of spaceflight simulation analysis. The method comprises the following steps: creating a planning task and setting basic parameters of a maneuvering satellite; setting basic parameters of a target satellite and a general control center position point; maneuvering analysis is carried out with maneuvering satellite parameters, target satellite parameters and the position point of the master control center as constraints, and a task time window meeting the constraints is obtained; after the task time window is determined, an orbital maneuvering multi-strategy algorithm is called for calculation to obtain a maneuvering mode and a maneuvering parameter list of the maneuvering satellite arriving at the task point; and finally, entering a planning event processing module, sorting events according to time, judging whether impact events exist or not, and outputting a planning task success / failure state. The general control center evaluates the fuel remaining amount in real time based on a fuel formula and supports a user to dynamically formulate and issue an orbital transfer scheme, and integrated planning of maneuvering satellite task window calculation, maneuvering strategy selection and safety check is achieved.

The application discloses a Beidou satellite maneuvering orbit recovery method, which comprises the following steps: S1, giving initial satellite state information; S2, detecting whether satellite orbit maneuvering occurs for each ephemeris based on current satellite state information; if yes, entering step S4; if no, entering step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by adopting combined short-arc orbit determination and grouping filtering method, and realizing maneuvering orbit recovery. The combined short-arc orbit determination and grouping filtering method realizes quick recovery of the Beidou satellite maneuvering orbit, so as to guarantee the navigation, positioning and timing service performance of the Beidou system in the case that the Beidou satellite occurs orbit maneuvering. The method can not only effectively improve the orbit determination precision in the case that the satellite occurs orbit maneuvering, but also greatly improve the reliability of the global service capability of the Beidou No. 3 system.

The invention discloses a satellite orbit forecasting method and device considering maneuvering acceleration, and belongs to the field of communication navigation remote sensing, and the method comprises the steps: carrying out the initial value calculation of a satellite orbit before orbit maneuvering, and obtaining an initial value calculation result; the initial value calculation result comprises six Kepler numbers at an initial moment and a sunlight pressure parameter; taking the orbit maneuvering acceleration as a random pulse, and introducing the random pulse and the initial value resolving result into a motion equation to obtain a simplified motion equation considering the maneuvering acceleration; wherein the starting moment, the ending moment and the numerical value of the orbit maneuvering acceleration are known parameters; the starting and ending moments determine the width of the random pulse; and performing numerical integration on the simplified motion equation under the constraint of a plurality of perturbative force models to obtain a satellite orbit forecasting result subjected to simplified dynamic constraint at any moment. According to the method, the track forecasting precision and the calculation load can be effectively balanced, and efficient and high-precision track forecasting is realized.

The invention discloses a maneuvering GEO satellite searching method based on optical orbit measurement data, which comprises the following steps: when an optical telescope does not observe a target satellite at an original running position during sky area patrol traversal, judging that the target satellite possibly has orbit maneuver, after a target is lost, carrying out initial orbit determination on unsuccessfully matched data searched by the optical telescope, and carrying out initial orbit determination on the data; suspected targets are screened based on orbit characteristics, combined orbit determination is performed on the screened data, after an initial orbit is determined, the optical telescope is continuously guided to search and accumulate data, and finally, the orbit of the maneuvering target satellite is determined. According to the method, the maneuvering target satellite (GEO satellite) can be quickly found in the optical telescope patrolling search data, the operation orbit of the maneuvering target satellite is mastered again, cataloguing management of the maneuvering target satellite is recovered, and the problem that the GEO satellite cannot be quickly found after large-amplitude orbit maneuvering is solved.

The present invention discloses a method and system for determining real-time precise orbits and clock offsets of LEO in orbit based on PPP-B2b enhancement, belonging to the technical field of high-precision autonomous orbit determination and clock offset estimation for low-earth orbit satellites. The present invention corrects the orbits and clock offsets of BDS-3 / GPS satellites by using the enhancement information broadcast by the B2b signal of Beidou-3. Under the support of regional BDS-3 / GPS enhancement information, when there are new BDS-3 / GPS tracking satellites, before measurement update, the epoch-to-epoch clock offset variation obtained by using the carrier phase time difference method is used as two orbit determination filtering clock offset parameters for constraint, and the receiver clock offset and undifferenced floating ambiguity parameters are separated in orbit in real time; finally, the precise orbit of LEO and the undifferenced floating ambiguity parameters output by orbit determination filtering are substituted into the phase observation equation to calculate the precise clock offset of the BDS-3 / GPS receiver. The present invention is applicable to the situation where there is only regional GNSS enhancement information support and is less affected by orbit maneuvers.

A method for detecting relative orbital maneuvers of spacecraft based on cluster analysis includes: reading the TLE historical data of the master star of the spacecraft formation and calculating the instantaneous orbital elements, using the SGP4 model for recursion to obtain the orbital state at the moment when the latitude angle is 0; reading the TLE historical data of the slave star of the spacecraft formation and calculating the instantaneous orbital elements, using the SGP4 model to recursively calculate the orbital state at the same time as the master star data, and realizing the time synchronization of the dual-star data; calculating the relative distance of the spacecraft formation, and performing data denoising and normalization to obtain a data set; using the silhouette coefficient to determine the number of clusters, that is, the number of orbital maneuver types; using the fuzzy C-means algorithm for cluster analysis, obtaining a clustering result based on the relative distance of the spacecraft orbit, and realizing orbital maneuver detection. The method of the present invention can effectively avoid the influence of human subjective cognition and experience on the judgment of orbital maneuvers, meet the timeliness requirements of orbital maneuver detection, and significantly improve the accuracy of orbital control maneuver detection.

The invention provides a low-orbit constellation configuration maintenance method using perturbation force and adjacent position relative station deviation, and belongs to the technical field of low-orbit constellation configuration maintenance and control. The method comprises the following steps: acquiring orbit parameters of each satellite and adjacent satellites thereof in a low-orbit satellite constellation in a space, and calculating a relative drift distance and an allowable drift range of each satellite; on the basis of a low-orbit satellite constellation dynamic model and the relative drift distance, calculating theoretical retention time of each satellite in an allowable drift range, and further determining a constellation control sequence and an optimal control moment of each satellite; and when the satellite reaches the optimal control moment, determining the control quantity of the satellite orbit parameters and the propulsion system parameters of the satellite during orbit maneuver so as to maintain the constellation configuration. According to the method, the relative standing position relation of the adjacent satellites serves as control input, autonomous stable maintenance of the constellation relative configuration is achieved by controlling the orbit height and the inclination angle of the satellites, and unnecessary maintenance control and fuel consumption are avoided.

The embodiment of the present application discloses a kind of super low orbit satellite orbit height autonomous maintenance method.In a specific embodiment, the method comprises: the position and speed of satellite-borne GNSS receiver output as input, calculate the actual value of satellite's intersection period when satellite is in ascending node or descending node, according to the relationship between intersection period actual value and intersection period target value, execute orbit maneuver, to maintain orbit height.A kind of super low orbit satellite orbit height autonomous maintenance method of the present application, with the position and speed of satellite-borne GNSS receiver output as input, is realized by combination of electric propulsion and chemical propulsion, can comprehensively utilize the advantages of less electric propulsion working medium consumption and large chemical propulsion thrust, to realize orbit height on-orbit autonomous orbit maintenance.

The invention provides a common-view time transfer method based on multiple GEO satellites, and the method comprises the steps: selecting a plurality of GEO satellites in the visual range of two tracking stations, carrying out the common-view time transfer through employing each GEO satellite, and obtaining the common-view time transfer result of each GEO satellite; preprocessing the common-view time transfer result of each GEO satellite to obtain a preprocessed common-view time transfer result; traversing each GEO satellite, analyzing the stability of the link of the current GEO satellite, and meanwhile, referring to a broadcast ephemeris to judge whether the current GEO satellite is in an orbit maneuvering state; if not in the track maneuvering state, allocating weights to the track maneuvering state based on a static self-adaptive weight allocation method, otherwise, allocating weights to the track maneuvering state based on a dynamic self-adaptive weight allocation method; and based on the weight distributed for each GEO satellite, performing weighted average processing on each preprocessed common-view time transfer result, and finally obtaining a common-view time transfer result of the multiple GEO satellites.

The invention provides a control method and device for a deep space spacecraft, belongs to the technical field of cosmic navigation, and can improve the control precision of the deep space spacecraft, thereby better reducing the fuel consumption of the deep space spacecraft during the long-term execution of a scientific exploration task. According to the method, orbit deviation between an orbit where the deep space spacecraft runs and a nominal orbit is divided into a divergent component and a non-divergent component of the deep space spacecraft, and a maneuvering response curve, indicating the non-divergent component, of the deep space spacecraft is obtained through analysis according to orbit change after orbit maneuvering of the deep space spacecraft. And determining a final value of the orbit maneuver of the deep space spacecraft capable of inhibiting the divergent component and the non-divergent component by combining the initial value of the orbit maneuver of the divergent component of the deep space spacecraft and the maneuver response curve of the deep space spacecraft.

The present invention discloses a method for close observation of a service spacecraft's multi-orbit maneuvering strategy based on J2 perturbation compensation, which solves the problem of how to quickly improve the accuracy of orbit maneuvering and belongs to the field of close observation of a service spacecraft's multi-orbit maneuvering strategy. The method includes: constructing a motion equation under J2 perturbation compensation in the target orbit coordinate system; constructing a multi-orbit maneuvering strategy model of the service spacecraft through the relative motion propagation relation and the motion equation; inputting the position at the termination moment that meets the distance constraint at the termination moment into the model to output the single-velocity increment, and retaining the single-velocity increments that meet both the illumination angle constraint and the single-velocity increment constraint at the termination moment to obtain a velocity increment solution set; constructing a circular region centered on the space target with a radius that meets the distance constraint at the termination moment; marking the task as successful when the service spacecraft reaches the circular region at the specified moment, and obtaining the multi-orbit maneuvering strategy of the service spacecraft for close observation. The present invention realizes precise and rapid orbit maneuvering.

The invention discloses an orbital maneuver compensation method based on constant acceleration and zero sum velocity pulse constraints, and relates to the technical field of satellite orbit calculation, comprising the steps of acquiring GNSS observation data, satellite attitude data, ephemeris data and exogenous data, performing orbit fitting to obtain a rough dynamic orbit, performing cycle slip detection and gross error elimination on the GNSS observation data, and performing orbit maneuver compensation on the rough dynamic orbit. Obtaining a clean observation value required by subsequent parameter estimation, adding a zero sum speed pulse constraint equation on the basis of an original observation error equation by using the clean observation value, a satellite state, a state transition matrix and a parameter sensitive matrix, constructing an augmented observation error equation, and iteratively performing parameter estimation and orbit integration until convergence, thereby obtaining a final observation value. According to the method, the orbit maneuvering compensation method based on the constant acceleration and the zero sum velocity pulse constraint is adopted, and the consistency between the dynamic orbit and the kinematic orbit and the KBR checking precision of a dynamic baseline are improved.

The invention discloses a space traffic information sensing method based on a star sensor network. The space traffic information sensing method comprises a space end and a ground end. At a space end, a plurality of artificial earth satellites form a regional or global network, and each satellite is provided with a star sensor, so that a star sensor network is formed, and the star sensors dynamically shoot space images in real time; an onboard computer of a star sensor is used for processing an original image according to the processes of image acquisition, feature extraction / centroid positioning, star map matching, target recognition / astronomical attitude determination and target orientation calculation, and space target measurement data is obtained. Space target measurement data are downloaded to a data processing center at a ground end, and the data processing center analyzes the space target measurement data according to processes of segmental arc association, maneuvering detection, cataloguing library updating, collision early warning and the like, and generates data products including a space target track, maneuvering information, a collision probability, a collision avoidance strategy and the like. And wide-area and rapid space traffic information perception is realized.

The application discloses a low-fuel-consumption low-orbit satellite orbit recovery method, and belongs to the technical field of satellite orbit recovery.The application aims at the problem that the existing orbit maneuver method needs frequent orbit adjustment, which is not conducive to the long-term operation of the satellite in orbit with low fuel consumption.The application comprises the following steps: establishing an orbit dynamics equation considering the earth oblateness perturbation; using a semi-analytical method to solve the initial orbit elements of a low-fuel-consumption transition orbit according to the current orbit and the target orbit of the satellite; simultaneously, combining the orbit dynamics equation, calculating the drift time required for the satellite to drift on the transition orbit by using the earth oblateness perturbation without fuel consumption and the final orbit elements of the transition orbit at the end of the transition; and based on the Pontryagin maximum principle, calculating the optimal control rate sequence of the thruster fuel of the satellite from the current orbit to the transition orbit and from the transition orbit to the target orbit, and controlling the thruster to realize the orbit recovery of the satellite.The application is used for satellite orbit recovery.

A solar sail includes a bus and a plurality of separate movable vanes coupled to the bus. Each movable vane includes a reflective surface for generating solar radiation pressure and propel the solar sail in space. Each vane may be movable relative to the bus in a fully deployed configuration such that an amount of thrust generated by solar radiation pressure on each vane is controllable.

The present invention provides a method and device for converting a pulsed velocity increment into finite thrust orbit control parameters, and the present invention relates to the field of orbital maneuvers. It includes: obtaining relevant parameters of satellite orbital maneuvers in pulsed mode, defining an inertial coordinate system, defining a spherical coordinate system, and the relevant parameters include satellite dry weight, fuel weight before maneuver, fuel weight after maneuver, orbital pulsed maneuver time, inertial coordinate system velocity increment V, and satellite and celestial body positions; calculating the thrust direction in the spherical coordinate system based on the relevant parameters; estimating the thrust magnitudes at the start and shutdown of the thruster based on the relevant parameters, calculating the on-time of the thruster with finite thrust as the initial value for iterative shooting; using the on-time of the finite thrust as the initial value, through continuous iterative shooting, finally generating a finite thrust velocity direction and on-time equivalent to the pulsed velocity increment. The present invention improves the accuracy of orbit control.

The invention provides an angle-measurement-only multi-pulse self-adaptive closed-loop rendezvous guidance method based on second-order cone programming, and belongs to the relative navigation and rendezvous guidance technology of spacecrafts. According to the method, firstly, an angle-only-measurement relative navigation dynamical model and a measurement model based on a CW equation are constructed under LVLH coordinates, system state observability analysis is carried out, the orbit maneuvering condition enabling the state to be observable is obtained, and the condition is used for constructing an optimal guidance index. Secondly, analyzing and modeling the rendezvous guidance error source, and deducing to obtain a relative orbit guidance error covariance model through error covariance analysis; thirdly, on the basis of a second-order cone programming method, guidance fuel consumption and system state observability conditions are fused, and a multi-target weighted cost function of optimal rendezvous guidance is jointly constructed; meanwhile, based on the established guidance error covariance model, a cost function weight index system with a self-adaptive adjustment function is designed, and real-time self-adaptive adjustment of a guidance strategy along with change of system performance is achieved. The method solves the problems of high fuel consumption, poor flexibility and the like caused by separate design of the relative navigation and the rendezvous guidance model in the traditional only-sight-line measurement rendezvous guidance method, and has the ability of adaptively adjusting and optimizing the rendezvous fuel consumption and the relative navigation system observability.

The present invention discloses a method and system for controlling orbital maneuver divergence during the gravity satellite center of mass calibration process, relating to the field of spacecraft control technology. The method comprises: configuring the output of a magnetic torquer so that the output of the magnetic torquer is converted from an angular acceleration signal to an initial square wave signal or an initial sine wave signal; setting the phase of the square wave signal or the initial sine wave signal at the initial moment to 1 / 4 of a cycle to obtain a square wave signal or a sine wave signal; and performing center of mass calibration on the gravity satellite based on the method signal or the sine wave signal. The present invention solves the problem of orbital maneuver divergence during the gravity satellite center of mass calibration process and prevents the gravity satellite from losing control during the center of mass calibration process.

The invention provides a satellite launching / recycling system capable of moving back and forth freely and being loaded on orbit, and relates to the technical field of spaceflight. The system comprises a spacecraft which can be in on-orbit butt joint with a preset space station; the spacecraft comprises a three-section cabin body which comprises a front-section cabin body, a middle-section cabin body and a rear-section cabin body; a head fairing capable of being opened and closed is arranged at the end part of the front section cabin; a butt joint opening is formed in the head fairing. The top facing part of the middle cabin body is configured to be a cabin door capable of being opened and closed; an expandable folding solar wing and at least one satellite parking device are arranged on the inner side of the cabin door; a satellite parking point is formed after the cabin door is opened; a loading space used for loading a satellite to be launched and / or a recycled satellite, a mechanical arm device capable of grabbing the satellite and a satellite on-orbit launching device are arranged in the middle cabin body. The spacecraft is used as a carrier for satellite launching / recycling and / or a carrier for implementing orbital maneuver and returning to the ground in orbit. Satellite launching and recycling can be achieved with low cost and high efficiency.

This invention relates to a ground flight control and operation support system for a very low Earth orbit (UEO) satellite constellation. It incorporates predicted space environment information, satellite attitude change information, and orbit control information into the high-precision satellite orbit prediction process, effectively ensuring higher orbit prediction accuracy for the UEO constellation and maximizing the assistance to ground control stations during satellite transits in capturing and tracking the satellite. The satellite attitude and orbit control system simulation module performs real-time simulations of the impact of attitude control and orbit control on orbital maneuvers, thereby achieving higher orbit control accuracy and providing better assurance for the precise execution of missions by the UEO constellation. Automatic scheduling of Earth remote sensing and data transmission tasks, along with ground system fault identification and handling, and collision warning, provide more intelligent support for the operation of the UEO constellation.

The invention discloses a multi-satellite pursuit game method based on a survival type differential game, and relates to a spaceflight pursuit satellite optimal method.The method comprises the steps that firstly, a particle swarm algorithm and a Newton iteration algorithm are combined, and a complex two-point boundary value problem is converted into a parameter optimization problem with constraints; and then solving a countermeasure saddle point corresponding to the game countermeasure for chasing the satellite. On the premise of considering the accuracy of the algorithm, shortening the calculation cost and ignoring the attitude problem of the satellite, the model is simplified, and the thrust direction angle of the satellite is used as the control quantity to realize the orbital maneuver of the satellite. Assuming that the satellites are all under the condition of continuous small thrust, taking terminal end time as a payment function by three game parties, taking a minimum function as a target by chasing the satellites, and contrarily taking the minimum function as the target by escaping the satellites. Simulation results show that the confrontation game of the pursuit satellites in the orbital plane is relatively intense, and the pursuit satellites with higher thrust acceleration have more advantages in confrontation.

A thruster (100) for orbital maneuver, comprising: a diffusion channel (111) having an axis of symmetry (S); a combustion chamber (110); a throat tube (112) interposed between the combustion chamber (110) and the diffusion channel (111); an injection plate (115) facing the combustion chamber (110); a plurality of first injection channels (125) in fluid communication with the combustion chamber (110), each first injection channel comprising an end portion (126) open in the combustion chamber (110); at least one second injection channel (135) in fluid communication with the combustion chamber (110), including an end portion (136) open in the combustion chamber (110); wherein an end portion (126) of the first injection channel (125) is located in the annular region (127) and extends along a respective injection direction (d1); wherein an end portion (136) of the at least one second injection channel (135) is located radially between the axis of symmetry (S) and an end portion (126) of the first injection channel (125); wherein the injection direction (d1) of the end portion (126) of each first injection channel (125) has an axial component directed towards the combustion chamber (110) and a radial component directed towards the axis of symmetry (S).

The present invention relates to an on-orbit estimation method for spacecraft mass characteristic parameters based on thrusters, including: spacecraft actuator configuration, establishment of a dynamic model, establishment of a kinematic model, establishment of an attitude measurement component model, on-orbit estimation of the three-axis moment of inertia based on thrusters, on-orbit estimation of the center of mass in the steady state mode and the orbital maneuver mode, and on-orbit estimation of the spacecraft mass. The present invention solves the problems of on-orbit estimation of the spacecraft moment of inertia, center of mass, and mass, and lays a solid technical foundation for the on-orbit redesign or correction of attitude and orbit control algorithms that rely on mass characteristic parameters. Through simulation analysis, this method is reasonable and effective, with a simple algorithm, small computational load, low fuel consumption during implementation, and is easy to implement in engineering. At the same time, a flexible model is added to the simulation of this method, so it is also applicable to flexible spacecraft.