A satellite electric thrust system and position keeping method

Abstract

This invention provides a satellite electric propulsion system, comprising: four electric thrusters, each mounted on the satellite's backplane, with its thrust direction pointing towards the satellite's center of mass. A rectangular region is defined on the satellite's backplane, with the geometric center of the rectangle located at the intersection of the satellite's center of mass and the projection onto the backplane. The four sides of the rectangle are perpendicular to the satellite's X-axis and Y-axis, respectively. The four electric thrusters are symmetrically mounted at the four vertices of the rectangle. A position-keeping method is provided, specifically: the four electric thrusters fire in turn; each day when the satellite is near its ascending node, the north-side thrusters fire in turn; each day when the satellite is near its descending node, the south-side thrusters fire in turn; the satellite uses the above method to perform inclination correction for north-south position-keeping, with the total firing time corresponding to the daily inclination correction required for north-south position-keeping. This solution improves propellant utilization efficiency and can meet the on-orbit position-keeping requirements of GEO satellites.

Description

Technical Field

[0001] The present invention relates to an on-orbit position keeping technology for geostationary orbit satellites, and in particular to a satellite electric thrust system and a position keeping method. Background Art

[0002] During the in-orbit operation phase, geostationary orbit (GEO) satellites need to regularly maintain their positions in the east-west and north-south directions, control the longitude and latitude of the satellite's sub-satellite point within the specified range, ensure that the satellite payload is pointed to the ground, and avoid electronic interference and physical collisions with adjacent satellites.

[0003] GEO satellites generally use chemical thrusters to maintain east-west position and north-south position, respectively. East-west position maintenance is generally performed once every 7 days, and north-south position maintenance is generally performed once every 14 days. With the rapid development of electric propulsion technology in space applications, the use of electric thrusters for position maintenance has been widely used on GEO satellites.

[0004] The thrust of chemical thrusters is larger, generally reaching 10N, but the specific impulse is lower, generally around 300s; the thrust of electric thrusters is smaller, generally less than 100mN, and the specific impulse is higher, generally around 3000s.

[0005] Since the specific impulse of electric thrusters is more than 10 times that of chemical thrusters, the fuel efficiency of electric thrusters for on-orbit position maintenance is greatly improved. Therefore, electric thrusters for position maintenance are the mainstream configuration of current GEO satellites.

[0006] The vast majority of GEO satellites use electric thrusters to maintain north-south position, while chemical thrusters are still used for east-west position maintenance, such as the A2100 platform, E3000 platform, SpaceBus4000 platform, etc.

[0007] In the prior art, different thrusters are used for north-south position keeping and east-west position keeping, respectively, and the propulsion efficiency and the use efficiency of the propellant are low. Summary of the invention

[0008] 1. Purpose:

[0009] In order to improve the propellant usage efficiency of GEO satellites using electric thrusters to maintain their on-orbit position, the electric thrusters are required to simultaneously maintain their position in both the east-west and north-south directions.

[0010] This patent proposes a satellite electric thrust system and position keeping method, which optimizes the parameters such as the ignition position and ignition duration of the electric thruster, improves the efficiency of the propellant, and can meet the on-orbit position keeping requirements of GEO satellites. This patent is applicable to geosynchronous orbit satellites that use electric thrusters for position keeping.

[0011] 2. Technical solution:

[0012] The technical solution of the present invention is:

[0013] The present invention provides a satellite electric propulsion system, comprising:

[0014] There are four electric thrusters, each of which is installed on the back floor of the satellite, and its thrust direction points to the center of mass of the satellite, wherein:

[0015] Two electric thrusters are arranged on each side of the satellite back floor, and are installed symmetrically with respect to the X-axis of the satellite back floor;

[0016] The two electric thrusters on the same side are installed symmetrically relative to the Y axis of the satellite back floor;

[0017] The satellite back floor is rectangular, and its two midlines are respectively the X-axis and the Y-axis.

[0018] Furthermore, there is a certain distance between the two electric thrusters on the same side.

[0019] The present invention proposes a position keeping method for the above-mentioned satellite electric thrust system, specifically comprising:

[0020] The four electric thrusters are ignited in turn;

[0021] Every day when the satellite is near the ascending node, the electric thrusters on the north side are ignited in turn;

[0022] Every day when the satellite is near the descending node, the south electric thrusters are ignited in turn;

[0023] The satellite uses the above method to make inclination corrections for north-south position maintenance, and the total ignition time corresponds to the inclination correction required for north-south position maintenance every day.

[0024] Furthermore, by adjusting the ignition timing of the four electric thrusters, east-west position keeping is performed simultaneously during the north-south position keeping.

[0025] The present invention provides an electric thruster ignition parameter optimization method, which is used in the above method and is characterized by:

[0026] Taking the moment when all ignition of the four electric thrusters is completed every day as the starting point, orbit prediction for a period of time is carried out, and the maximum deviation and maximum eccentricity of the satellite's relative longitude during the period are taken as optimization targets, and the ignition duration of the four electric thrusters are taken as optimization variables. The optimization method is used to optimize the ignition duration of the electric thrusters every day, which can ensure that the east-west position is maintained every day and the deviation and eccentricity of the satellite's relative longitude can be controlled to the minimum.

[0027] Furthermore, the optimization method is the simplex method.

[0028] 3. Effect

[0029] This method adjusts the layout of electric thrusters, formulates a strategy for electric thrusters to simultaneously maintain both east-west and north-south positions, designs an optimization strategy for electric thruster ignition parameters, and can automatically obtain parameters such as electric thruster ignition position and ignition duration, thereby improving the efficiency of propellant use and meeting the on-orbit position maintenance requirements of GEO satellites. This optimization strategy has the following advantages:

[0030] 1) It can simultaneously maintain the east-west and north-south positions of GEO satellites, thus improving the efficiency of propellant use.

[0031] 2) The electric thrusters work most efficiently when the north-south position is maintained at the ascending and descending nodes. The total ignition time of the four electric thrusters can be calculated based on the daily inclination correction. Only the ignition duration of the four electric thrusters needs to be optimized, which reduces the dimension of the optimization problem, helps shorten the optimization solution time, and improves the feasibility of the project.

[0032] The ignition duration parameters of the electric thrusters can be directly obtained through optimization, which facilitates the autonomous satellite on-orbit position keeping operation. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 A schematic diagram of the definition of a geostationary orbit satellite coordinate system in an embodiment of the present invention;

[0034] Figure 2 It is a schematic diagram of the layout of the electric thruster of the geostationary orbit satellite in an embodiment of the present invention;

[0035] Figure 3 Schematic diagram of the all-electric propulsion position keeping strategy for a geostationary orbit satellite in an embodiment of the present invention;

[0036] Figure 4 The inclination change curve of the geostationary orbit satellite during the 28-day position keeping period in the embodiment of the present invention;

[0037] Figure 5 The eccentricity variation curve of the geostationary orbit satellite during the 28-day position keeping period in the embodiment of the present invention;

[0038] Figure 6 The longitude change curve of the geostationary orbit satellite during the 28-day position keeping period in the embodiment of the present invention;

[0039] Figure 7 1 is a curve showing the change of longitude and latitude during the 28-day position keeping period of a geostationary orbit satellite in an embodiment of the present invention. DETAILED DESCRIPTION

[0040] In order to make the technical solutions and advantages of the present invention clearer, the technical solutions of the embodiments of the present invention will be fully described below in conjunction with the drawings in the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0041] Embodiments of the present invention are described in detail below, and examples of the embodiments are shown in the accompanying drawings, wherein the same or similar reference numerals throughout represent the same or similar elements or elements having the same or similar functions. The embodiments described below with reference to the accompanying drawings are exemplary and are only used to explain the present invention, and cannot be understood as limiting the present invention.

[0042] First, some coordinate systems and calculation formulas required for the present invention are given.

[0043] For the convenience of description, it is necessary to establish a satellite mechanical coordinate system, which is defined as follows: the coordinate origin O is located in the mechanical separation plane between the lower end frame of the satellite and the launch vehicle, and coincides with the theoretical center of the docking ring interface; the OX axis (roll axis) passes through the origin O, is parallel to the theoretical normal direction of the satellite's east plate, and its positive direction is consistent with the outer normal direction of the east plate; the OZ axis (yaw axis) passes through the origin O, is parallel to the satellite axis, and its positive direction points to the ground; the OY axis (pitch axis) forms a right-handed rectangular coordinate system with the OX and OY axes, and its positive direction is consistent with the outer normal direction of the south plate. The payload module is located on top of the propulsion service module, see Figure 1 .

[0044] The specific implementation process of the present invention is as follows:

[0045] (1) Electric thruster layout: Four electric thrusters are used. Each electric thruster is installed on the back floor of the satellite through a dual-axis vector adjustment mechanism. The WN and EN electric thrusters are located on the north side, the WS and ES electric thrusters are located on the south side, the WN and WS are located on the west side, and the EN and ES are located on the east side. The four electric thrusters are installed symmetrically about the OX axis and the OY axis. The east and west electric thrusters are kept at a certain distance, see Figure 2 The dual-axis vector adjustment mechanism can direct the thrust direction of the electric thruster toward the satellite's center of mass. At this time, each electric thruster can generate thrust in the north-south direction and the east-west direction at the same time.

[0046] (2) All-electric thrust position keeping strategy: Every day, the four electric thrusters are ignited in turn. The WN and EN electric thrusters on the north side near the ascending node are ignited in turn, and the WS and ES electric thrusters on the south side near the descending node are ignited in turn. This can achieve the inclination correction for north-south position keeping. The total ignition time corresponds to the inclination correction required for north-south position keeping every day. By adjusting the ignition time of the four electric thrusters, the east-west position keeping can be completed simultaneously during the north-south position keeping, see Figure 3Specifically:

[0047] (3) Electric thruster ignition parameter optimization strategy: The starting point is the time when all ignitions of the four electric thrusters are completed every day, which is defined as T f , make orbit prediction for ΔT time, and calculate the relative longitude λ of the satellite during this period 0 Deviation Δλ=λ-λ 0 And the eccentricity e is counted and used as the optimization target. When longitude deviation control is performed, the optimization objective function can be defined as Opt = min{max[λ(t)-λ 0 ]},t f ≤t≤t f +Δt, when eccentricity control is performed, the optimization objective function can be defined as

[0048] Opt=min{max[e(t)]},t f ≤t≤t f +Δt, the ignition duration of the four electric thrusters is taken as the optimization variable x = [Δt WN Δt EN Δt WS Δt ES ], using optimization methods (such as the simplex method) to optimize the calculation of the electric thruster ignition duration for the next day, it can be ensured that the deviation of the satellite's relative longitude and eccentricity relative to the fixed point can be controlled to the minimum every day.

[0049] In one embodiment of the present invention, it is assumed that the satellite fixed longitude is 120° east longitude, the east-west and north-south position keeping accuracy requirements are both ±0.05°, and the initial orbit parameters are: semi-major axis 42167.2km, eccentricity 0.0001, inclination 0.01°, ascending node right ascension 90°, perigee 0°. The indicators of a single electric thruster are: thrust 80mN, specific impulse 3800s. The installation positions of the electric thrusters are: ES electric thruster is (500,1600,150)mm, EN electric thruster is (500,-1600,150)mm, WS electric thruster is (-500,1600,150)mm, and WN electric thruster is (-500,-1600,150)mm. The satellite center of mass height is 2000mm. The 28-day electric thruster in-orbit east-west and north-south position keeping strategy after optimized calculation is as follows Figures 4 to 7 As shown in the figure, it can be seen that the east-west and north-south position accuracy requirements are basically met.

[0050] In the description of this specification, reference to the terms "one embodiment" and "example" means that the specific features, structures or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic representation of the above terms does not necessarily refer to the combination in a suitable manner in the relative embodiment or example.

[0051] It must be pointed out that the description of the above embodiments is not intended to be limiting but only to help understand the core idea of ​​the present invention. For ordinary technicians in this technical field, any improvements to the present invention and alternatives equivalent to the product, without departing from the principles of the present invention, also fall within the scope of protection of the claims of the present invention.

Claims

1. A position keeping method, characterized in that: Based on satellite electric propulsion system, The system includes: There are four electric thrusters, each of which is installed on the back floor of the satellite, and its thrust direction points to the center of mass of the satellite, wherein: A rectangular area is defined on the satellite back floor. The geometric center of the rectangle is located at the intersection of the projection of the satellite's centroid on the back floor. The four sides of the rectangle are perpendicular to the satellite's X-axis and Y-axis respectively. Four electric thrusters are symmetrically installed at the four vertices of the rectangle. There is a certain distance between the two electric thrusters on the same side; The method is as follows: The four electric thrusters are ignited in turn; Every day when the satellite is near the ascending node, the electric thrusters on the north side are ignited in turn; Every day when the satellite is near the descending node, the south electric thrusters are ignited in turn; The satellite uses the above method to make inclination corrections for north-south position maintenance, and the total ignition time corresponds to the inclination correction required for north-south position maintenance every day; By adjusting the ignition timing of the four electric thrusters, the east-west position keeping is performed simultaneously during the north-south position keeping; The step of optimizing the ignition parameters of the electric thruster is also included, and the step comprises: Taking the moment when all ignition of the four electric thrusters is completed every day as the starting point, orbit prediction for a period of time is carried out, and the maximum deviation and maximum eccentricity of the satellite relative to the fixed longitude during the period are taken as optimization targets, and the ignition duration of the four electric thrusters are taken as optimization variables. The optimization method is used to optimize the ignition duration of the electric thrusters every day, which can ensure that the east-west position is maintained every day, and the deviation and eccentricity of the satellite relative to the fixed longitude can be controlled to a minimum.

2. The method according to claim 1, characterized in that The optimization method is the simplex method.

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