A satellite electric propulsion system and an on-orbit position keeping method in case of failure

Abstract

The present invention provides a satellite electric propulsion system, including: electric thrusters, each of which has a thrust direction pointing to the center of mass of the satellite, wherein: a rectangular area is defined on the back floor of the satellite, the geometric center of the rectangle is located at the intersection of the projection of the satellite center of mass on the back floor, the four sides of the rectangle are respectively perpendicular to the satellite X-axis and Y-axis, and four electric thrusters are symmetrically installed at the four vertices of the rectangle. And a method for maintaining position on orbit in the event of a fault, including the following steps: when near the ascending node, the remaining electric thrusters on the north side are ignited in turn; when near the descending node, the remaining electric thrusters on the south side are ignited in turn; the satellite uses the above method to correct the inclination angle of the north-south position, and the total ignition time of the remaining electric thrusters each day corresponds to the inclination angle correction required for the north-south position maintenance each day. Using the technical solution of the present application, parameters such as the ignition position and ignition duration of the electric thrusters can be automatically obtained, the use efficiency of the propellant is improved, and the in-orbit position maintenance requirements of the GEO satellite can be met.

Description

Technical Field

[0001] The present invention relates to an on-orbit position keeping technology for a geostationary orbit satellite, and in particular to a satellite electric propulsion system and an on-orbit position keeping method in the event of a fault. 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] GEO satellites are generally equipped with four electric thrusters. In normal mode, the four electric thrusters are ignited in turn every day. When one or two electric thrusters fail, the normal mode electric propulsion position keeping strategy cannot effectively maintain the east-west and north-south position. Summary of the invention

[0008] 1. Purpose:

[0009] When the electric thrusters are maintaining position on orbit, one or two failures may occur. At this time, the normal mode electric thruster position keeping strategy cannot effectively maintain the east-west and north-south positions. It is necessary to design an optimized strategy for maintaining the east-west and north-south positions under the failure mode for the remaining electric thrusters.

[0010] This paper proposes an optimization strategy for on-orbit east-west and north-south position keeping in the event of a single or two diagonally installed electric thrusters failing. By optimizing parameters such as the electric thruster ignition position and ignition duration, the efficiency of propellant use is improved, which 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 present invention provides a satellite electric propulsion system, characterized in that:

[0013] There are four electric thrusters, each of which is installed on the back floor of the satellite through a dual-axis vector adjustment mechanism, and its thrust direction points to the center of mass of the satellite, wherein:

[0014] 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 center of mass 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.

[0015] The dual-axis vector adjustment mechanism can point the thrust direction of the electric thruster toward the center of mass of the satellite. At this time, each electric thruster can generate thrust in the north-south and east-west directions at the same time.

[0016] Furthermore, there is a certain distance between the two electric thrusters on the same side.

[0017] The present invention proposes an on-orbit position keeping method in the event of a fault, based on the above-mentioned satellite electric propulsion system, comprising the following steps:

[0018] Every day when the satellite is near the ascending node, the remaining electric thrusters on the north side are ignited in turn, and the midpoint of the accumulated total ignition arc length of the electric thrusters corresponds to the position of the ascending node;

[0019] Every day when the satellite is near the descending node, the remaining electric thrusters on the south side are ignited in turn, and the midpoint of the accumulated total ignition arc length of the electric thrusters corresponds to the position of the descending node;

[0020] The satellite uses the above method to perform inclination correction for north-south position maintenance, and the total ignition time of the electric thrusters remaining each day corresponds to the inclination correction required for north-south position maintenance each day.

[0021] Furthermore, by adjusting the ignition time of each remaining electric thruster, the correction of the fixed-point longitude deviation is simultaneously completed during the north-south position keeping.

[0022] Furthermore, when a single electric thruster fails, after the remaining three electric thrusters are ignited in turn, the fourth and fifth ignitions are carried out using two electric thrusters that are not faulty and are located diagonally opposite to each other. The fifth ignition is carried out immediately after the fourth ignition. The fourth and fifth ignitions have the same duration, generating a net radial component for eccentricity control. The fourth ignition duration and the fourth ignition end position are determined with the goal of eliminating the additional eccentricity generated by the first to third ignitions.

[0023] Furthermore, when two diagonal thrusters fail, the remaining two electric thrusters are used to perform the fourth and fifth ignitions respectively. The fifth ignition is performed immediately after the fourth ignition. The fourth and fifth ignitions have equal durations, generating a net radial component for eccentricity control. The fourth ignition duration and the fourth ignition end position are determined with the goal of eliminating the additional eccentricity generated by the first to third ignitions.

[0024] The present invention proposes an electric thruster ignition parameter optimization method, which is used in the above method and comprises the following steps:

[0025] Taking the moment when all ignition of the electric thrusters at the ascending node and descending node of each day is completed as the starting point, orbit prediction for a period of time is carried out, and the minimum maximum deviation of the satellite relative to the fixed longitude during this period is taken as the optimization goal. The ignition duration of the electric thrusters at the ascending node and descending node of each day is taken as the optimization variable, and the optimization method is used to optimize the ignition duration of the electric thrusters to ensure that the deviation of the satellite relative to the fixed longitude is controlled to the minimum. If the deviation does not meet the east-west position maintenance accuracy requirement, the total ignition time of the ascending node and descending node of each day is appropriately increased, and the optimization calculation is re-performed until the deviation meets the east-west position maintenance accuracy requirement.

[0026] The present invention proposes an electric thruster ignition parameter optimization method, which is used in the above method and comprises the following steps:

[0027] The fourth ignition end time and the fourth ignition duration are used as optimization variables, and the completion time of all ignitions of the ascending node and descending node electric thrusters every day is taken as the starting point. Orbit prediction for a period of time is carried out, and the minimum maximum eccentricity of the satellite during this period is taken as the optimization goal. The fourth ignition end time and its corresponding ignition duration are optimized and calculated using the optimization method, and the satellite's eccentricity is corrected to the minimum every day.

[0028] Furthermore, the optimization method is an exhaustive method or a parameter optimization method.

[0029] 3. Effect

[0030] The present invention proposes a satellite electric propulsion system and an on-orbit position keeping method in the event of a fault, which can automatically obtain parameters such as the ignition position and ignition duration of the electric thruster, improve the use efficiency of the propellant, and meet the on-orbit position keeping requirements of the GEO satellite. The present invention has the following advantages:

[0031] 1) In the event of a single or two diagonal electric thrusters failing, the remaining electric thrusters can simultaneously maintain the GEO satellite's position in both the east-west and north-south directions.

[0032] 2) The electric thrusters are most efficient when the north-south position is maintained at the ascending and descending nodes. The total ignition time of the electric thrusters can be calculated based on the daily inclination correction. In the case of a single electric thruster failure, only the ignition duration of the remaining three electric thrusters needs to be optimized. In the case of a diagonal electric thruster failure, only the ignition duration of the remaining two electric thrusters needs to be optimized. The optimization goal only needs to consider the minimum deviation of the fixed point position.

[0033] 3) The added 4th and 5th ignitions only affect the eccentricity and will not cause changes in the inclination angle and fixed point deviation, making it easier to obtain the optimal solution.

[0034] 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

[0035] Figure 1 is a schematic diagram of the definition of a geostationary orbit satellite coordinate system in an embodiment of the present invention;

[0036] Figure 2 is a schematic diagram of the layout of an electric thruster for a geostationary orbit satellite in an embodiment of the present invention;

[0037] Figure 3 Schematic diagram of a position keeping strategy in the case of a single electric propulsion failure of a geostationary orbit satellite in an embodiment of the present invention;

[0038] Figure 4 2 is a schematic diagram of an inclination change curve during a 28-day position keeping period of a geostationary orbit satellite in an embodiment of the present invention;

[0039] Figure 5 2 is a schematic diagram of an eccentricity variation curve during a 28-day position keeping period of a geostationary orbit satellite in an embodiment of the present invention;

[0040] Figure 6 2 is a schematic diagram of a longitude variation curve during a 28-day position keeping period of a geostationary orbit satellite in an embodiment of the present invention;

[0041] Figure 7 Schematic diagram of the longitude and latitude variation curve of the geostationary orbit satellite during the 28-day position keeping period in an embodiment of the present invention. DETAILED DESCRIPTION

[0042] 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.

[0043] 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.

[0044] First, some coordinate systems and calculation formulas required for the present invention are given.

[0045] 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 .

[0046] The specific implementation process of the present invention is as follows:

[0047] (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.

[0048] (2) Full electric propulsion position holding strategy:

[0049] In case of a single electric thruster failure, the remaining three electric thrusters will be ignited in turn every day. The northern electric thruster near the ascending node will be ignited, and the southern electric thruster near the descending node will be ignited. The electric thrusters at the ascending node and descending node are used to correct the orbital inclination and the east-west longitude drift rate. The total ignition time corresponds to the correction amount of the orbital inclination every day. The 4th and 5th ignitions are added to find the best ignition position in the unignited area within one orbital cycle. The remaining two diagonal electric thrusters will be ignited in turn. The ignition duration of the two electric thrusters is equal, and a net radial component is generated for eccentricity control. If the WN electric thruster fails and only one electric thruster is allowed to ignite at the same time: near the ascending node, t asc is the time of passing the ascending node, at t asc -Δt EN / 2, the EN electric thruster starts to ignite, and the duration is Δt EN ; Near the descending node, t dec is the descending node moment, at t dec -Δt WS At time t, the WS electric thruster starts to ignite, and the duration is Δt WS ; at t dec At time t, the ES electric thruster starts to ignite, and the duration is Δt ES , total ignition time Δt EN +Δt WS +Δt ES Corresponding to the correction of orbital inclination every day. ecc -Δt ECC At this moment, the EN electric thruster ignites, and the duration is Δt ECC , at t ecc At this moment, the WS electric thruster ignites, and the duration is Δt ECC , EN and WS electric thrusters produce a net radial component for eccentricity control.

[0050] When two diagonal electric thrusters fail, the remaining two electric thrusters will ignite in turn every day. The northern electric thruster near the ascending node will ignite, and the southern electric thruster near the descending node will ignite. The electric thrusters at the ascending node and descending node are used to correct the orbital inclination and the east-west longitude drift rate. The total ignition time corresponds to the correction amount of the orbital inclination every day. The 4th and 5th ignitions are added to find the best ignition position in the unignited area within one orbital cycle. The remaining two diagonal electric thrusters will ignite in turn. The ignition duration of the two electric thrusters is equal, and a net radial component is generated for eccentricity control. If the WN and ES electric thrusters fail, and only one electric thruster is allowed to ignite at the same time: near the ascending node, t asc is the time of passing the ascending node, at t asc -Δt EN / 2, the EN electric thruster starts to ignite, and the duration is ΔtEN ; Near the descending node, t dec is the descending node moment, at t dec -Δt WS / 2, the WS electric thruster starts to ignite, and the duration is Δt WS , total ignition time Δt EN +Δt WS Corresponding to the correction of orbital inclination every day. ecc -Δt ECC At this moment, the EN electric thruster ignites, and the duration is Δt ECC , at t ecc At this moment, the WS electric thruster ignites, and the duration is Δt ECC , EN and WS electric thrusters produce a net radial component for eccentricity control.

[0051] (3) Electric thruster ignition parameter optimization strategy:

[0052] The starting point is the time when all the five electric thruster ignitions are completed each day, which is defined as T f , make orbit prediction at ΔT time, count the deviation Δλ=λ-λ0 of the satellite relative to the fixed longitude λ0 during this period and use it as the optimization target. The optimization target function can be defined as Opt=min{max[λ(t)-λ0]}, t f ≤t≤t f +Δt, taking the ignition duration of the ascending node and the descending node electric thruster as the optimization variable x, using the optimization method (such as the simplex method) to optimize the ignition duration of the electric thruster, and determine the ignition duration of the ascending node and the descending node electric thruster corresponding to the minimum optimization function value; taking the 4th and 5th ignition alternation time and ignition duration as the optimization variables, the optimization objective function can be defined as Opt = min{max[e(t)]}, t f ≤t≤t f +Δt, optimize the electric thruster ignition position and duration using an optimization method (such as the simplex method) to determine the electric thruster ignition position and duration corresponding to the minimum optimization function value. According to the above optimization steps, it can be ensured that the deviation and eccentricity of the satellite relative to the fixed point longitude can be minimized every day.

[0053] A specific implementation of the present invention is as follows: Assuming that the satellite fixed-point longitude is 120° east longitude, the east-west and north-south position maintenance 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 angle 0°. The indicators of a single electric thruster are: thrust 80mN, specific impulse 3800s. The installation position of the electric thruster is: 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. Assuming that the WN electric thruster fails, the optimized calculation results of the 365-day electric thruster position maintenance in the east-west and north-south directions are as follows: Figure 4 to Figure 7 As shown in the figure, it can be seen that the east-west and north-south position accuracy requirements are basically met.

[0054] The main requirement for determining the vicinity of the ascending node and the descending node is that the midpoint of the total arc segment should be at the ascending node or the descending node. The ignition time length is determined by the inclination correction amount, which corresponds to the range.

[0055] 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.

[0056] 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 method for maintaining position on track in the event of a fault, characterized in that: Satellite electric propulsion system: The system includes: There are four electric thrusters, each of which is installed on the back floor of the satellite through a dual-axis vector adjustment mechanism, and its thrust direction points to the center of mass of the satellite, wherein: A rectangular area is defined on the back floor of the satellite. The geometric center of the rectangular area 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 comprises the following steps: Every day when the satellite is near the ascending node, the remaining electric thrusters on the north side are ignited in turn, and the midpoint of the accumulated total ignition arc length of the electric thrusters corresponds to the ascending node position; Every day when the satellite is near the descending node, the remaining electric thrusters on the south side are ignited in turn, and the midpoint of the accumulated total ignition arc length of the electric thrusters corresponds to the position of the descending node; The satellite uses the above method to perform inclination correction for north-south position maintenance, and the total ignition time of the electric thrusters remaining each day corresponds to the inclination correction amount required for north-south position maintenance each day; By adjusting the ignition time of each remaining electric thruster, the correction of the fixed point longitude deviation is completed simultaneously during the north-south position keeping; When a single electric thruster fails, after the remaining three electric thrusters are ignited in turn, the fourth and fifth ignitions are performed using two electric thrusters that are not faulty and are located diagonally opposite to each other. The fifth ignition is performed immediately after the fourth ignition. The fourth and fifth ignitions have the same duration, and a net radial component is generated to control the eccentricity. The fourth ignition duration and the fourth ignition end position are determined with the goal of eliminating the additional eccentricity generated by the first to third ignitions. When two diagonal electric thrusters fail, the remaining two electric thrusters are used for the fourth and fifth ignitions respectively. The fifth ignition is carried out immediately after the fourth ignition. The fourth and fifth ignitions have equal durations, generating a net radial component for eccentricity control. The fourth ignition duration and the fourth ignition end position are determined with the goal of eliminating the additional eccentricity generated by the first to third ignitions.

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