Satellite Constellation Electric Propulsion Deorbiting Control Method and System

Abstract

The present invention provides a satellite constellation electric propulsion deorbiting control method and system, which relates to the technical field of spacecraft orbit design and control. The method includes: Step S1: According to the deorbiting mission requirements, select the characteristic point height of the deorbiting orbit, and calculate the semi-major axis and eccentricity of the target orbit; Step S2: Adjust the semi-major axis of the orbit, continuously fire along the tangential direction until the semi-major axis is reduced to the target value; Step S3: Adjust the eccentricity of the orbit, fire throughout the orbit, and fire along the direction perpendicular to the apsidal line until the eccentricity meets the target value; Step S4: Fire at the orbit characteristic point until the perigee height and apogee height meet the deorbiting requirements. The present invention can achieve rapid evaluation and optimization of the deorbiting orbit parameter set; by adopting the strategy of firing in different coordinate systems in stages, the coupling between the semi-major axis and the eccentricity is reduced, which is beneficial to simplifying the complexity of the actual mission deorbiting scheme.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft orbit design and control, and particularly to an electric propulsion deorbiting control method and system for a satellite constellation. Background Art

[0002] The realization of the overall function of the constellation depends on the normal operation of all networked satellites. When a certain satellite fails and is difficult to be completely repaired, it is necessary to implement an end-of-life disposal strategy for it, and make it deorbit and enter the atmosphere to burn up through orbit control, or transfer it to a discarded orbit to prevent the generation of space debris by colliding with other satellites. For low-earth orbit satellites with an orbital altitude of about 1000 km, when deorbiting, it can be placed into a deorbiting orbit with a perigee altitude of about 400 km, and the atmospheric drag is used to pull it into the atmosphere and crash.

[0003] When the perigee altitude of the deorbiting orbit is determined, the deorbiting time shortens as the apogee altitude decreases, and a lower apogee altitude requires the satellite to increase a larger velocity increment. Electric propulsion has the advantages of high specific impulse and long life, but also has the disadvantages of small thrust and long orbit adjustment time. When using electric propulsion for deorbiting, it is urgent to consider the requirement of rapid deorbiting of the faulty satellite to reduce the risk of its collision with other satellites. Therefore, it is necessary to study the rapid deorbiting electric propulsion control algorithm for such satellites.

[0004] At present, some research has been carried out on the orbit transfer of spacecraft. After retrieval, the main ones related to algorithm design are as follows:

[0005] The invention patent with the publication number of CN101767657A discloses a satellite deorbiting device and method based on an electrodynamic tether, which solves the problems of high deorbiting cost of the propellant deorbiting method and low deorbiting efficiency of the atmospheric drag deorbiting method, thereby providing a satellite deorbiting device and method based on an electrodynamic tether. This patent focuses on the design of the deorbiting device using the space electromagnetic environment, while this patent focuses on the electric propulsion orbit control algorithm, with a large difference.

[0006] The invention patent with the publication number of CN103144784B discloses a deorbiting method applicable to navigation GEO satellites, and gives the deorbiting operation and requirements of navigation GEO satellites, ensuring that the navigation GEO satellites can smoothly enter the discarded orbit after the end of their service life. This patent focuses on the deorbiting control and operation of high-orbit satellites, while this patent focuses on the low-orbit orbit control algorithm, with great differences in algorithm design and operation.

[0007] The invention patent with the publication number CN108860662A discloses a method for deorbiting medium-orbit satellites based on solar radiation pressure. Considering that there is no clear principle for dealing with abandoned satellites in the medium-orbit region and the development prospect of solar sail technology, a perturbation force model and a long-term evolution model of abandoned satellites are established. The long-term evolution is analyzed with and without solar sails, and the feasibility of using solar sails to achieve re-entry into the atmosphere is verified. This patent focuses on the solar radiation pressure deorbiting control and operation of medium-orbit satellites, while this patent focuses on the electric propulsion orbit control algorithm for low orbits, and there are significant differences in the control object, model, algorithm design, and operation.

[0008] Generally speaking, the attention paid to the research on the fast deorbiting electric propulsion control algorithm for low-orbit satellite constellations is still insufficient, and the relevant research results are also relatively lacking, which does not match its important application value. Therefore, the deorbiting control of satellite constellations should be further studied in depth to provide an effective orbit control scheme for its mission design. Summary of the Invention

[0009] In view of the deficiencies in the prior art, the present invention provides a method and system for electric propulsion deorbiting control of satellite constellations.

[0010] According to a method and system for electric propulsion deorbiting control of satellite constellations provided by the present invention, the solution is as follows:

[0011] In a first aspect, a method for electric propulsion deorbiting control of satellite constellations is provided, and the method includes:

[0012] Step S1: According to the deorbiting mission requirements, select the characteristic point height of the deorbiting orbit, and calculate the semi-major axis and eccentricity of the target orbit;

[0013] Step S2: Adjust the semi-major axis of the orbit, and continuously fire along the tangential direction until the semi-major axis is reduced to the target value;

[0014] Step S3: Adjust the eccentricity of the orbit, fire throughout the orbit, and fire along the direction perpendicular to the apsidal line until the eccentricity meets the target value;

[0015] Step S4: Fire at the orbit characteristic points until the perigee height and apogee height meet the deorbiting requirements.

[0016] Preferably, the step S1 includes:

[0017] Step S1.1: Generate a perigee / apogee height envelope matrix according to the set deorbiting time;

[0018] Step S1.2: In the envelope matrix traversal, calculate the deorbiting time corresponding to each adopted point;

[0019] Step S1.3: Compare the off-orbit times corresponding to each sampling point, select the perigee / apogee altitudes therefrom according to the principle of minimum fuel consumption, determine the off-orbit orbit parameters, and use the two-body model to calculate the corresponding semi-major axis and eccentricity.

[0020] Preferably, the step S2 includes:

[0021] According to the semi-major axis of the current initial orbit and the semi-major axis of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio;

[0022] In the orbital coordinate system, set the ignition direction to be along the negative X-axis; adopt the strategy of full-orbit ignition, and continuously lower the orbit according to the calculated ignition time until the semi-major axis meets the requirements.

[0023] Preferably, the step S3 includes: According to the eccentricity of the current on-orbit orbit and the eccentricity of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio;

[0024] Establish an apsidal-fixed coordinate system with the origin of the coordinate system being the satellite's center of mass, the X-axis pointing to the perigee mass, the Y-axis perpendicular to the X-axis in the orbital plane and pointing to the flight direction, and the Z-axis forming a right-handed coordinate system with the other two axes;

[0025] In the apsidal-fixed coordinate system, set the ignition direction to be along the Y-axis, adopt the strategy of full-orbit ignition, and continuously lower the orbit according to the calculated ignition time until the eccentricity meets the requirements.

[0026] Preferably, the step S4 includes:

[0027] Calculate the current perigee / apogee altitudes according to the adjustment results of step S2 and step S3;

[0028] Compare with the target characteristic point altitude, calculate the difference in perigee altitude, ignition time and velocity increment, and ignite along the X-axis of the orbital coordinate system at the apogee to adjust the perigee altitude;

[0029] Compare with the target characteristic point altitude, calculate the difference in apogee altitude, ignition time and velocity increment, and ignite along the X-axis of the orbital coordinate system at the perigee to adjust the apogee altitude.

[0030] In a second aspect, a satellite constellation electric propulsion off-orbit control system is provided, and the system includes:

[0031] Module M1: Select the characteristic point altitude of the off-orbit orbit according to the off-orbit mission requirements, and calculate the semi-major axis and eccentricity of the target orbit;

[0032] Module M2: Adjust the semi-major axis of the orbit, continuously ignite along the tangential direction until the semi-major axis is reduced to the target value;

[0033] Module M3: Adjust the eccentricity of the orbit, perform full-orbit ignition, and ignite along the direction of the vertical apsidal line until the eccentricity meets the target value;

[0034] Module M4: Ignite at the orbital characteristic points until the perigee altitude and apogee altitude meet the requirements for orbit departure.

[0035] Preferably, the module M1 includes:

[0036] Module M1.1: Generate the perigee / apogee altitude envelope matrix according to the set orbit departure time;

[0037] Module M1.2: Traverse and adopt in the envelope matrix, and calculate the corresponding orbit departure time for each adopted point;

[0038] Module M1.3: Compare the orbit departure times corresponding to each sampling point, select the perigee / apogee altitude according to the principle of minimum fuel consumption, determine the orbit departure orbit parameters, and use the two-body model to calculate the corresponding semi-major axis and eccentricity.

[0039] Preferably, the module M2 includes:

[0040] According to the semi-major axis of the current initial orbit and the semi-major axis of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio;

[0041] In the orbital coordinate system, set the ignition direction to be along the negative X-axis; adopt the strategy of full-orbit ignition, and continuously lower the orbit according to the calculated ignition time until the semi-major axis meets the requirements.

[0042] Preferably, the module M3 includes:

[0043] According to the eccentricity of the current on-orbit orbit and the eccentricity of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio;

[0044] Establish an apsidal line-fixed coordinate system, with the origin of the coordinate system being the satellite's center of mass, the X-axis pointing to the perigee mass, the Y-axis perpendicular to the X-axis in the orbital plane and pointing to the flight direction, and the Z-axis forming a right-handed coordinate system with the other two axes;

[0045] In the apsidal line-fixed coordinate system, set the ignition direction to be along the Y-axis, adopt the strategy of full-orbit ignition, and continuously lower the orbit according to the calculated ignition time until the eccentricity meets the requirements.

[0046] Preferably, the module M4 includes:

[0047] Calculate the current perigee / apogee altitude according to the adjustment results of module M2 and module M3;

[0048] Compare the heights of the target feature points, calculate the difference in the height of the near point, the ignition time, and the velocity increment, and ignite along the X-axis of the orbital coordinate system at the apogee to adjust the height of the perigee;

[0049] Compare the heights of the target feature points, calculate the difference in the height of the far point, the ignition time, and the velocity increment, and ignite along the X-axis of the orbital coordinate system at the perigee to adjust the height of the apogee.

[0050] Compared with the prior art, the present invention has the following beneficial effects:

[0051] 1. The present invention can quickly calculate, evaluate, and determine the deorbiting orbit parameters according to the deorbiting requirements, fully consider the requirements of rapid deorbiting, realize the rapid evaluation and optimization of the deorbiting orbit parameter set, adopt the strategy of decoupled adjustment of the semi-major axis and eccentricity, provide an effective orbit control method for the deorbiting of satellite constellations, and realize the rapid evaluation and optimization of the deorbiting orbit parameter set;

[0052] 2. By adopting the strategy of staged ignition in different coordinate systems, the present invention effectively reduces the coupling between the semi-major axis and eccentricity, which is beneficial to simplifying the complexity of the deorbiting scheme of actual tasks. BRIEF DESCRIPTION OF THE DRAWINGS

[0053] By reading the following detailed description of non-limiting embodiments with reference to the accompanying drawings, other features, objects, and advantages of the present invention will become more apparent:

[0054] Figure 1 It is a flowchart of the method of the present invention;

[0055] Figure 2 It is the time history of the semi-major axis in the specific implementation manner;

[0056] Figure 3 It is the time history of the eccentricity in the specific implementation manner;

[0057] Figure 4 It is the time history of the apogee height in the specific implementation manner;

[0058] Figure 5 It is the time history of the perigee height in the specific implementation manner. DETAILED DESCRIPTION OF THE INVENTION

[0059] The following describes the present invention in detail with reference to specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but do not limit the present invention in any form. It should be noted that those of ordinary skill in the art can make several changes and improvements without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0060] The embodiment of the present invention provides a satellite constellation electric propulsion deorbiting control method, referring toFigure 1 As shown in Figure 1 , the method specifically includes:

[0061] Step S1: According to the requirements of the deorbiting mission, select the characteristic point height of the deorbiting orbit, and calculate the semi-major axis and eccentricity of the target orbit;

[0062] In this step, first, according to the set deorbiting time, generate the perigee / apogee height envelope matrix, and then traverse and adopt in the envelope matrix to calculate the deorbiting time corresponding to each adopted point; then compare the deorbiting times corresponding to each sampling point, and according to the principle of the least fuel consumption, select the perigee / apogee height from them to determine the deorbiting orbit parameters, and use the two-body model to calculate the corresponding semi-major axis and eccentricity.

[0063] Step S2: Adjust the semi-major axis of the orbit, and continuously fire along the tangential direction until the semi-major axis is reduced to the target value;

[0064] In this step, first, according to the semi-major axis of the current initial orbit and the semi-major axis of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio;

[0065] Secondly, in the orbital coordinate system, set the ignition direction to be along the negative X-axis;

[0066] Finally, adopt the strategy of full-orbit ignition, and continuously lower the orbit according to the calculated ignition time until the semi-major axis meets the requirements.

[0067] Step S3: Adjust the eccentricity of the orbit, and perform full-orbit ignition, and fire along the direction perpendicular to the line of apsides until the eccentricity meets the target value;

[0068] Specifically, in this step, according to the eccentricity of the current on-orbit orbit and the eccentricity of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio;

[0069] Establish a line-of-apsides fixed coordinate system, with the origin of the coordinate system as the satellite's center of mass, the X-axis pointing to the perigee mass, the Y-axis perpendicular to the X-axis in the orbital plane and pointing to the flight direction, and the Z-axis forming a right-handed coordinate system with the other two axes;

[0070] In the line-of-apsides fixed coordinate system, set the ignition direction along the Y-axis, adopt the strategy of full-orbit ignition, and continuously lower the orbit according to the calculated ignition time until the eccentricity meets the requirements.

[0071] Step S4: Fire at the orbital characteristic points until the perigee height and apogee height meet the deorbiting requirements.

[0072] According to the adjustment results of Step S2 and Step S3, calculate the current perigee / apogee height;

[0073] Compare the heights of the target feature points, calculate the difference in the height of the perigee, the ignition time, and the velocity increment, and ignite along the X-axis of the orbital coordinate system at the apogee to adjust the height of the perigee.

[0074] Compare the heights of the target feature points, calculate the difference in the height of the apogee, the ignition time, and the velocity increment, and ignite along the X-axis of the orbital coordinate system at the perigee to adjust the height of the apogee.

[0075] The following is the numerical simulation verification of the electric propulsion deorbiting control method for satellite constellations:

[0076] According to the following calculation conditions, perform numerical simulation calculations, and the results are referred to Figure 2 and Figure 3 As shown, in the simulation calculation, the semi-major axis of the working orbit is set to 7378.137 km, the eccentricity is 0.001, and the orbital inclination is 60°; at the same time, according to the calculation conditions, the results are also referred to Figure 4 and Figure 5 As shown, the perigee height and apogee height of the deorbiting orbit are 400 km and 900 km respectively, the corresponding semi-major axis is 7028.137 km, and the eccentricity is 0.0355713. The satellite thrust-to-weight ratio is set to 0.0001 m / s 2 , and the specific impulse is 1600 s.

[0077] The embodiment of the present invention provides an electric propulsion deorbiting control method for satellite constellations, which realizes the rapid evaluation and optimization of the deorbiting orbit parameter set; by adopting the strategy of staged ignition in different coordinate systems, the coupling of the semi-major axis and eccentricity is effectively reduced, which is beneficial to simplifying the complexity of the actual mission deorbiting scheme.

[0078] Those skilled in the art know that in addition to implementing the system and its various devices, modules, and units provided by the present invention in the form of pure computer-readable program code, the method steps can be logically programmed to enable the system and its various devices, modules, and units provided by the present invention to be implemented in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers to achieve the same function. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered as a hardware component, and the devices, modules, and units included therein for implementing various functions can also be regarded as the structure within the hardware component; the devices, modules, and units for implementing various functions can also be regarded as both software modules for implementing the method and the structure within the hardware component.

[0079] The specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. Without conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A method for controlling the orbit departure of a satellite constellation using electric propulsion, characterized in that, Including: Step S1: According to the requirements of the deorbiting mission, select the characteristic point height of the deorbiting orbit, and calculate the semi-major axis and eccentricity of the target orbit; Step S2: Adjust the semi-major axis of the orbit, continuously fire along the tangent direction until the semi-major axis is reduced to the target value; Step S3: Adjust the eccentricity of the orbit, fire throughout the orbit, and fire along the direction perpendicular to the apsidal line until the eccentricity meets the target value; Step S4: Fire at the characteristic points of the orbit until the perigee height and apogee height meet the deorbiting requirements; The said Step S1 includes: Step S1.1: Generate the perigee / apogee height envelope matrix according to the set deorbiting time; Step S1.2: Traverse and adopt in the envelope matrix, and calculate the deorbiting time corresponding to each adopted point; Step S1.3: Compare the deorbiting times corresponding to each sampling point, and select the perigee / apogee height according to the principle of the most fuel-saving, determine the deorbiting orbit parameters, and use the two-body model to calculate the corresponding semi-major axis and eccentricity.

2. The method for controlling the orbit departure of a satellite constellation using electric propulsion according to claim 1, characterized in that, The said Step S2 includes: According to the semi-major axis of the current initial orbit and the semi-major axis of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio; In the orbital coordinate system, set the ignition direction to be along the negative X-axis; adopt the strategy of firing throughout the orbit, and continuously lower the orbit according to the calculated ignition time until the semi-major axis meets the requirements.

3. The method for controlling the orbit departure of a satellite constellation using electric propulsion according to claim 1, characterized in that, The said Step S3 includes: According to the eccentricity of the current on-orbit orbit and the eccentricity of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio; Establish an apsidal line-fixed coordinate system, with the origin of the coordinate system as the satellite's center of mass, the X-axis pointing to the perigee mass, the Y-axis perpendicular to the X-axis in the orbital plane and pointing to the flight direction, and the Z-axis forming a right-handed coordinate system with the other two axes; In the apsidal line-fixed coordinate system, set the ignition direction along the Y-axis, adopt the strategy of firing throughout the orbit, and continuously lower the orbit according to the calculated ignition time until the eccentricity meets the requirements.

4. The method for controlling the orbit departure of a satellite constellation using electric propulsion according to claim 1, characterized in that, The said Step S4 includes: According to the adjustment results of Step S2 and Step S3, calculate the current perigee / apogee height; Compare with the target characteristic point height, calculate the difference in perigee height, ignition time and velocity increment, and fire along the X-axis of the orbital coordinate system at the apogee to adjust the perigee height; Compare with the target characteristic point height, calculate the difference in apogee height, ignition time and velocity increment, and fire along the X-axis of the orbital coordinate system at the perigee to adjust the apogee height.

5. A satellite constellation electric propulsion orbit departure control system, characterized in that, Including: Module M1: According to the requirements of the deorbiting mission, select the characteristic point height of the deorbiting orbit, and calculate the semi-major axis and eccentricity of the target orbit; Module M2: Adjust the semi-major axis of the orbit, continuously fire along the tangent direction until the semi-major axis is reduced to the target value; Module M3: Adjust the eccentricity of the orbit, fire throughout the orbit, and fire along the direction perpendicular to the apsidal line until the eccentricity meets the target value; Module M4: Fire at the characteristic points of the orbit until the perigee height and apogee height meet the deorbiting requirements; The said Module M1 includes: Module M1.1: Generate the perigee / apogee height envelope matrix according to the set deorbiting time; Module M1.2: Traverse and adopt in the envelope matrix, and calculate the deorbiting time corresponding to each adopted point; Module M1.3: Compare the off-orbit times corresponding to each sampling point, select the perigee / apogee altitudes according to the principle of minimum fuel consumption, determine the off-orbit orbit parameters, and use the two-body model to calculate the corresponding semi-major axis and eccentricity.

6. The satellite constellation electric propulsion orbit departure control system according to claim 5, characterized in that, The said module M2 includes: According to the semi-major axis of the current initial orbit and the semi-major axis of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio; In the orbital coordinate system, set the ignition direction to be along the negative X-axis; adopt the strategy of full-orbit ignition, and continuously lower the orbit until the semi-major axis meets the requirements according to the calculated ignition time.

7. The satellite constellation electric propulsion deorbiting control system according to claim 5, characterized in that, The said module M3 includes: According to the eccentricity of the current on-orbit orbit and the eccentricity of the target orbit, calculate the corresponding velocity increment and ignition time according to the two-body model and the satellite thrust-to-weight ratio; Establish an apsidal-fixed coordinate system with the origin at the satellite's center of mass, the X-axis pointing to the perigee mass, the Y-axis perpendicular to the X-axis in the orbital plane and pointing to the flight direction, and the Z-axis forming a right-handed coordinate system with the other two axes; In the apsidal-fixed coordinate system, set the ignition direction along the Y-axis, adopt the strategy of full-orbit ignition, and continuously lower the orbit until the eccentricity meets the requirements according to the calculated ignition time.

8. The satellite constellation electric propulsion deorbiting control system according to claim 5, characterized in that, The said module M4 includes: Calculate the current perigee / apogee altitudes according to the adjustment results of module M2 and module M3; Compare with the target feature point altitude, calculate the difference in perigee altitude, ignition time and velocity increment, and ignite along the X-axis of the orbital coordinate system at the apogee to adjust the perigee altitude; Compare with the target feature point altitude, calculate the difference in apogee altitude, ignition time and velocity increment, and ignite along the X-axis of the orbital coordinate system at the perigee to adjust the apogee altitude.

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