A method and apparatus for adjusting the inclination of a sun-synchronous orbit

Abstract

This invention provides a method and apparatus for adjusting the inclination of a sun-synchronous orbit. The method includes: determining the positional relationship between an initial orbit and a target orbit, such that the two intersection points of the initial orbit and the target orbit are located above the equator, and using one of the intersection points as the satellite's orbit change point; determining the orbit change angle required for the satellite to adjust from the initial orbital plane to the target orbital plane; comprehensively determining the satellite's single ignition duration, ignition time, and shutdown time based on the satellite's system status and mission objectives; determining the satellite's ignition direction; analytically calculating the velocity increment of the satellite's single pulse ignition, and determining the total impulse required to change the satellite's orbit inclination based on the velocity increment; analytically calculating the total number of satellite ignitions based on the total impulse and the actual impulse generated by each ignition; and controlling the satellite to perform a single ignition orbit change to complete the inclination adjustment based on the satellite's orbit change point, ignition direction, ignition time, and shutdown time.

Description

Technical Field

[0001] This invention relates to the field of spacecraft orbit control technology, and in particular to a method and apparatus for adjusting the inclination of a sun-synchronous orbit. Background Technology

[0002] A sun-synchronous orbit is an orbit whose orbital plane precession angular velocity is the same as the Earth's angular velocity around the Sun. A key characteristic is the definite correlation between its orbital inclination and altitude. Satellites operating in this type of orbit can pass over a specific region at the same local time during each revisit cycle, making them widely used in remote sensing satellite missions such as Earth observation, environmental monitoring, and weather forecasting. To ensure consistent lighting conditions in a specific area and to achieve periodic revisits, these satellites typically need to maintain a fixed descending node local time, which is the local time corresponding to when the satellite crosses the equatorial plane from north to south.

[0003] Ideally, the orbital plane of a sun-synchronous orbit maintains a fixed angle (Beta angle) with the direction of sunlight, ensuring the stability of the local time at the descending node. However, in practical engineering, unavoidable errors during launch and insertion into orbit cause a slight deviation in the initial orbital inclination of the satellite. Although this initial deviation is small, it accumulates over time, causing a continuous drift in the right ascension of the ascending node, which in turn leads to a gradual deviation of the local time at the descending node from the design value. This deviation results in a difference between the actual overpass time and the preset mission time, affecting the continuity and consistency of data acquisition. In severe cases, it can even prevent the satellite from performing its observation mission as planned, reducing the overall effectiveness of the satellite system.

[0004] The drift rate of the right ascension of the ascending node is primarily influenced by both orbital inclination and orbital altitude. For satellites operating in low Earth orbits, frequent or significant adjustments to orbital altitude can disrupt mission execution; therefore, active orbit lowering is typically not used for long-term drift correction during routine orbit maintenance. Thus, how to correct the long-term drift of the right ascension of the ascending node through appropriate orbital control measures without significantly altering orbital altitude becomes a crucial issue in maintaining the characteristics of a sun-synchronous orbit. Summary of the Invention

[0005] To address the aforementioned technical problems, this invention proposes a method and apparatus for adjusting the inclination angle of a sun-synchronous orbit.

[0006] This invention provides a method for adjusting the inclination of a sun-synchronous orbit to correct right ascension drift at the ascending node. The method includes at least the following steps:

[0007] Step 1: Determine the positional relationship between the initial orbit and the target orbit, so that the two intersection points of the initial orbit and the target orbit are located above the equator, and use one of the intersection points as the satellite orbit change point;

[0008] Step 2: Determine the orbital change angle required for the satellite to adjust from the initial orbital plane to the target orbital plane. ;

[0009] Step 3: Determine the duration of a single satellite ignition based on the satellite's system status and mission objectives. And based on the duration of a single ignition The orbit change point determines the satellite's ignition and shutdown times;

[0010] Step 4: Based on the satellite's speed at the orbit change point in its initial orbit. The coordinate system representation, and the velocity of the satellite at the point of orbit change in the target orbit. The coordinate system is used to determine the satellite ignition direction;

[0011] Step 5: Based on the satellite's speed in its initial orbit Speed ​​of travel on the target orbit and satellite orbit change angle The velocity increment of satellite single-pulse ignition was calculated analytically. And according to the speed increment Determine the total impulse required to change the inclination of a satellite's orbit. ;

[0012] Step Six: Based on the total stroke and the actual impulse generated by each ignition The total number of orbits for satellite ignition was calculated analytically. ;

[0013] Step 7: Based on the satellite's orbit change point, ignition direction, ignition time, and shutdown time, control the satellite to perform... Tilt adjustment is completed after the second ignition and track change.

[0014] Furthermore, the method based on the single ignition duration The method for determining the satellite's ignition and shutdown times based on the orbit change point is as follows: with the orbit change point as the center of symmetry, ignition begins when the satellite reaches the first position point before the orbit change point, and ends when the satellite reaches the second position point after the orbit change point; wherein, the time taken for the satellite to travel from the first position point to the orbit change point is equal to the time taken to travel from the orbit change point to the second position point.

[0015] Furthermore, the speed of the satellite at the point of orbit change in its initial orbit... The coordinate system representation, and the velocity of the satellite at the point of orbit change in the target orbit. The method for determining the satellite's ignition direction using coordinate system representation is as follows: Establish the J2000 coordinate system; the velocity of the satellite in its initial orbit passing through the orbit change point... In the J2000 coordinate system, it is expressed as: The speed of a satellite orbiting in the target orbit as it passes the orbit change point In the J2000 coordinate system, it is expressed as: The ignition direction of the satellite from ignition time to shutdown time in the J2000 coordinate system is expressed as:

[0016] .

[0017] Furthermore, the speed of the satellite operating in its initial orbit... Speed ​​of travel on the target orbit and satellite orbit change angle The velocity increment of satellite single-pulse ignition was calculated analytically. The specific method is as follows: According to the formula: The pulse velocity increment required for tilt adjustment is calculated. ;in, Theoretical inclination value corresponding to the target orbital altitude The tilt angle value of the actual satellite telemetry data.

[0018] Furthermore, the statement based on speed increment Determine the total impulse required to change the inclination of a satellite's orbit. The specific method is as follows: through the formula: The total impulse required to change the inclination of the satellite orbit was calculated; among which... For satellite quality.

[0019] Furthermore, the statement based on the total impulse and the actual impulse generated by each ignition The total number of orbits for satellite ignition was calculated analytically. The specific method is as follows: through the formula: The actual impulse generated by each satellite ignition was calculated. ,in, The thrust generated by satellite ignition The duration of a single satellite ignition. To improve satellite ignition efficiency;

[0020] Through the formula: The total number of orbits required for satellite ignition was calculated. .

[0021] Furthermore, the duration of a single satellite ignition... No more than one-third of the satellite's orbital period.

[0022] Furthermore, if the satellite carries a GNSS receiver, and its output data is the position and velocity in the WGS84 coordinate system, then when collecting telemetry data, the position and velocity in the WGS84 coordinate system should first be converted to the position and velocity in the J2000 coordinate system. The specific method is to multiply the position and velocity in the WGS84 coordinate system by the polar motion matrix, time angle matrix, nutation matrix, and precession matrix in sequence to obtain the position and velocity in the J2000 coordinate system.

[0023] In another aspect, the present invention provides a sun-synchronous orbit inclination adjustment device, comprising at least a memory, a processor, and a computer program stored in the memory and executable on the processor; the processor is configured to execute the sun-synchronous orbit inclination adjustment method described in any of the above embodiments based on the computer program stored in the memory.

[0024] Compared with traditional satellite orbit control methods, the sun-synchronous orbit inclination adjustment method and device of the present invention integrates theoretical calculation, engineering implementation and on-orbit verification, thereby achieving high precision, high autonomy and high reliability of orbit control.

[0025] I. This invention, through an innovative calculation method, can quickly and accurately calculate the optimal ignition direction, single ignition duration, and number of ignitions, and perfectly coordinate them with the satellite control system. This not only avoids fuel waste and orbital errors caused by directional deviations, but also simplifies the flight control process, enhances the satellite's ability to autonomously maintain its orbit and maneuver, effectively extends the satellite's on-orbit lifespan, and reduces operating costs.

[0026] Second, after verification, this invention can enable low-orbit satellites to adjust their inclination using the propulsion system, accurately implement orbital phase adjustments, optimize orbital conditions, and improve the efficiency of on-orbit mission execution.

[0027] Upon reading the detailed embodiments and examining the accompanying drawings, those skilled in the art will recognize additional features and advantages. Attached Figure Description

[0028] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0029] Figure 1 This is a flowchart of the sun-synchronous orbit inclination adjustment method according to an embodiment of the present invention.

[0030] Figure 2This is a schematic diagram showing the positional relationship between the initial orbit and the target orbit in an embodiment of the present invention.

[0031] Figure 3 This is a schematic diagram of the satellite orbit change ignition position according to an embodiment of the present invention.

[0032] Figure 4 This is a flowchart illustrating the conversion of the WGS84 coordinate system to the J2000 coordinate system according to an embodiment of the present invention.

[0033] Figure 5 This is a comparison diagram of the right ascension of the ascending node of the standard solar orbit before tilt adjustment and the actual solar orbit.

[0034] Figure 6 This is a comparison diagram of the local time of the descending node of the standard solar orbit and the actual solar orbit before tilt adjustment.

[0035] Figure 7 This is a schematic diagram showing the change in tilt angle after 11 satellite ignition adjustments.

[0036] Figure 8 This is a comparison diagram of the local time of the descending node of the standard solar orbit and the actual solar orbit after tilt adjustment. Detailed Implementation

[0037] The features and exemplary embodiments of various aspects of the present invention will now be described in detail. To make the objectives, technical solutions, and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only configured to explain the present invention and to exemplify the principles of the present invention, and are not configured to limit the present invention. In addition, the structural components in the drawings are not necessarily drawn to scale. For example, the dimensions of some structural components or regions in the drawings may be enlarged for other structural components or regions to aid in the understanding of the embodiments of the present invention.

[0038] The directional terms used in the following description refer to the directions shown in the figures and are not intended to limit the specific structure of the embodiments of the present invention. In the description of the present invention, it should be noted that, unless otherwise stated, the terms "installation," "connection," and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand the specific meaning of the above terms in the present invention according to the specific circumstances.

[0039] Furthermore, the terms "comprising," "including," "having," or any other variations thereof are intended to cover non-exclusive inclusion, such that a structure or component that includes a list of elements includes not only those elements but also other structural elements that are not expressly listed or inherent to the structure or component. Without further limitations, an element defined by the phrase "comprising..." does not exclude the presence of other identical elements in the article or apparatus that includes the element.

[0040] Spatial relation terms such as "below," "under," "under," "low," "above," "on," and "high" are used for descriptive convenience to explain the positioning of one element relative to a second element, indicating that these terms are intended to cover different orientations of the device, in addition to those different from those shown in the figure. Furthermore, phrases such as "one element on / below another element" can indicate that two elements are in direct contact, or that there are other elements between the two elements. In addition, terms such as "first" and "second" are also used to describe individual elements, areas, parts, etc., and should not be considered limiting. Similar terms are used throughout the description to refer to similar elements.

[0041] It will be apparent to those skilled in the art that the present invention can be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention.

[0042] This invention aims to solve the problem of adjusting the inclination of a satellite operating in a sun-synchronous orbit. According to the formula for the rate of change of the right ascension of the ascending node, the drift rate of the right ascension of the ascending node is mainly affected by both the orbital inclination and the orbital altitude.

[0043] The formula for the rate of change of right ascension at the ascending node:

[0044] Where μ is the Earth's gravitational constant, J2 is the second-order perturbation constant, RE is the Earth's radius, a is the semi-major axis of the orbit, and I is the orbital inclination. Therefore, the rate of change of the right ascension of the ascending node is related to both the semi-major axis and the inclination; that is, adjusting the right ascension of the ascending node can adjust either the orbital altitude or the inclination. However, for satellites operating in low Earth orbits, frequent or significant adjustments to orbital altitude can affect mission execution. Therefore, to maximize the utilization of the satellite's lifespan, instead of actively lowering the orbit for long-term drift correction during routine orbit maintenance, adjusting the satellite's orbital inclination can be used. The inclination adjustment cycle is much shorter than the altitude adjustment cycle.

[0045] See Figure 1 This invention provides a method and apparatus for adjusting the inclination of a sun-synchronous orbit, used to correct right ascension drift of the ascending node. The method for adjusting the inclination of a sun-synchronous orbit according to embodiments of this invention includes at least the following steps:

[0046] S10. Determine the positional relationship between the initial orbit and the target orbit, so that the two intersection points of the initial orbit and the target orbit are located above the equator, and use one of the intersection points as the satellite orbit change point;

[0047] S20. Determine the orbital change angle required for the satellite to adjust from the initial orbital plane to the target orbital plane. ;

[0048] S30. Determine the duration of a single ignition of the satellite based on the satellite's system status and mission objectives. And based on the duration of a single ignition The orbit change point determines the satellite's ignition and shutdown times;

[0049] S40. Based on the satellite's speed at the orbit change point in its initial orbit. The coordinate system representation, and the velocity of the satellite at the point of orbit change in the target orbit. The coordinate system is used to determine the satellite ignition direction;

[0050] S50, based on the satellite's speed in its initial orbit. Speed ​​of travel on the target orbit and satellite orbit change angle The velocity increment of satellite single-pulse ignition was calculated analytically. And according to the speed increment Determine the total impulse required to change the inclination of a satellite's orbit. ;

[0051] S60, according to the total stroke and the actual impulse generated by each ignition The total number of orbits for satellite ignition was calculated analytically. ;

[0052] S70, based on the satellite's orbit change point, ignition direction, ignition time, and shutdown time, controls the satellite to perform... Tilt adjustment is completed after the second ignition and track change.

[0053] In this embodiment, a schematic diagram of the change in satellite orbital inclination is shown below. Figure 2 As shown, the initial orbit is orbit (0), and the target orbit is orbit (1). The optimal orbit change point is at the intersection of the two orbits. For a sun-synchronous orbit, the intersection of the initial orbit and the target orbit is located above the equator. If an inclination change is to be made, a single-pulse ignition is required at the intersection of the orbits whose sub-satellite point is located above the equator.

[0054] Specifically, after determining the positional relationship between the initial orbit and the target orbit, the instantaneous orbital state of the satellite at the initial orbit change point is known, including its spatial position vector. and velocity vector By applying a precisely calculated pulsed velocity increment at this moment. The satellite's instantaneous velocity will become And its location The velocity state is considered constant at the instant the pulse is applied. This abrupt change in velocity state will cause the satellite to deviate from its original orbit and enter a state with new initial conditions ( , The transfer trajectory determined by the given information is used to complete the change of the orbital plane.

[0055] The actual duration of a single satellite ignition depends on the mission requirements and the selected thrusters. Generally, to ensure satellite energy balance and meet the needs of on-orbit mission execution, the duration of a single satellite ignition is... It does not exceed one-third of the satellite's orbital period, which shows that it is difficult for a satellite to change its orbit in a single operation.

[0056] S10. Determine the positional relationship between the initial orbit and the target orbit, so that the two intersection points of the initial orbit and the target orbit are located above the equator, and use one of the intersection points as the satellite orbit change point.

[0057] Satellite fuel is a precious resource that needs to be used efficiently. Therefore, during orbital plane changes, orbit control is usually performed near the ascending / descending node to improve fuel utilization efficiency. To simplify mission analysis and flight control procedures, this embodiment selects the satellite's ascending node as the orbit change point.

[0058] It should be noted that in this embodiment, ignition does not begin at the trajectory change point, but rather the center moment of the entire ignition process is located at the trajectory change point.

[0059] S20. Determine the orbital change angle required for the satellite to adjust from the initial orbital plane to the target orbital plane. .

[0060] Generally, the inclination angle corresponding to each different orbital altitude in a sun-synchronous orbit is fixed; that is, each orbital altitude in a sun-synchronous orbit corresponds to a unique inclination angle. Therefore, the inclination angle of the target orbit is known. However, due to errors in rocket launches and satellite insertion, the actual inclination angle of the satellite's orbit is not the theoretical value. The actual orbital inclination angle of the satellite is obtained from telemetry data transmitted from the satellite to the ground. Therefore:

[0061] In this step, Theoretical inclination value corresponding to the target orbital altitude The tilt angle value from actual satellite telemetry data. For ease of understanding, please refer to... Figure 2 The line connecting the two intersections in S10 is called the nodal line. The satellite rotates around this nodal line from the initial orbital plane. After the angle, it reaches the target orbital plane. This step accurately calculates... The angle directly determines the velocity increment and direction of the satellite during orbit change, dominates the fuel consumed to change the orbital plane orientation, and provides a precise theoretical basis and command reference for subsequent orbit control.

[0062] S30. Determine the duration of a single ignition of the satellite based on the satellite's system status and mission objectives. And based on the duration of a single ignition The orbit change point determines the satellite's ignition and shutdown times.

[0063] Specifically, ignition is performed symmetrically around the orbit change point, starting when the satellite reaches its first position before the orbit change point and ending when the satellite reaches its second position after the orbit change point. The time taken for the satellite to travel from the first position point to the orbit change point is equal to the time taken to travel from the orbit change point to the second position point.

[0064] Because single-pulse ignition applies a speed increment at a specific point... While orbit changes are possible, in practice, low-Earth orbit satellites cannot have their velocity increments applied instantaneously at a single point. Therefore, the actual orbit change operation scheme is as follows:

[0065] like Figure 3 As shown, during actual ignition, the node O where the two orbits (the current satellite orbit 0 and the target orbit 1) intersect is taken as the center point of the ignition duration. Ignition begins when the satellite passes point A and ends when it passes point B. The ignition direction is... The direction is shown. That is, points A and B are time-symmetrical points about point O. For example, if the satellite ignites at point A and passes point O after 60 seconds, then the position the satellite reaches after passing point O and continuing for another 60 seconds is point B, where it shuts down. In this embodiment, the satellite ignition starts at point A and ends at point B, lasting 120 seconds, symmetrical about point O. Regardless of whether the satellite uses chemical or electric propulsion, the duration of a single ignition considers the satellite's energy balance and on-orbit mission requirements. That is, the time it takes for the satellite to pass through arc AB does not exceed one-third of the satellite's orbital period.

[0066] S40. Based on the satellite's speed at the orbit change point in its initial orbit. The coordinate system representation, and the velocity of the satellite at the point of orbit change in the target orbit. The coordinate system is used to determine the satellite ignition direction.

[0067] The actual on-orbit velocity of a satellite is generally expressed in the J2000 coordinate system. The J2000 coordinate system uses 12:00 noon on January 1, 2000 (Gregorian calendar) as the relative reference time.

[0068] That is, the J2000.0 geocentric inertial coordinate system, origin. It is the Earth's center of mass.

[0069] The axis points to the vernal equinox at J2000.0 (12:00 on January 1, 2000 AD).

[0070] The plane is the Earth's horizontal equatorial plane at time J2000.0;

[0071] Axis perpendicular to It is flat and points towards the Earth's North Pole;

[0072] Generally, the right-hand rule applies.

[0073] Step S40 specifically involves: establishing the J2000 coordinate system, where the velocity of the satellite in its initial orbit passing through the orbit change point is... In the J2000 coordinate system, it is expressed as: The speed of a satellite orbiting in the target orbit as it passes the orbit change point In the J2000 coordinate system, it is expressed as: The ignition direction of the satellite from ignition time to shutdown time in the J2000 coordinate system is expressed as:

[0074] .

[0075] S50, based on the satellite's speed in its initial orbit. Speed ​​of travel on the target orbit and satellite orbit change angle The velocity increment of satellite single-pulse ignition was calculated analytically. And according to the speed increment Determine the total impulse required to change the inclination of a satellite's orbit. .

[0076] in:

[0077] Based on the satellite's speed in its initial orbit Speed ​​of travel on the target orbit and satellite orbit change angle The velocity increment of satellite single-pulse ignition was calculated analytically. The specific method is as follows: According to the formula: The pulse velocity increment required for tilt adjustment is calculated. .in, Theoretical inclination value corresponding to the target orbital altitude The tilt angle value of the actual satellite telemetry data.

[0078] Based on speed increment Determine the total impulse required to change the inclination of a satellite's orbit. The specific method is as follows: through the formula: The total impulse required to change the inclination of the satellite orbit was calculated, where For satellite quality.

[0079] S60, according to the total stroke and the actual impulse generated by each ignition The total number of orbits for satellite ignition was calculated analytically. The specific method is as follows: through the formula: The actual impulse generated by each satellite ignition was calculated. ,in, The thrust generated by satellite ignition The duration of a single satellite ignition. For satellite ignition efficiency.

[0080] Then, using the formula: The total number of orbits required for satellite ignition was calculated. .

[0081] Among them, the total number of satellite ignition orbits It is a natural number greater than or equal to 1.

[0082] Based on the satellite orbit change point, satellite ignition direction, satellite ignition time, and power-off time calculated in the above steps, the satellite is controlled to perform... Tilt adjustment is completed after the second ignition and track change.

[0083] It should be understood that the inclination adjustment of low-orbit circular orbits can be referenced in this invention, and is not limited to sun-synchronous orbits.

[0084] Furthermore, satellites typically carry GNSS receivers, which directly output position and velocity in the WGS84 coordinate system. Therefore, if a satellite carries a GNSS receiver and outputs position and velocity data in the WGS84 coordinate system, the position and velocity data in the WGS84 coordinate system must first be converted to the position and velocity data in the J2000 coordinate system when acquiring telemetry data.

[0085] WGS84 is a coordinate system established for use by the GPS global positioning and navigation system. It consists of a global geocentric reference frame, an Earth gravity field model, and the WGS84 horizontal plane. The position and velocity output by the GPS receiver adopts this coordinate system.

[0086] In the WGS84 coordinate system, the origin of the coordinate system is... ;

[0087] Axis: Pointing towards the Conventional Earth Pole (CTP) as defined by the BIH (International Time Service) in 1984.0;

[0088] Axis: The intersection of the zero meridian plane of BIH1984.0 and the CTP equator;

[0089] Axis: Generally, the right-hand rule applies.

[0090] The specific method for converting the position and velocity in the WGS84 coordinate system to the J2000 coordinate system is as follows: Multiply the position and velocity in the WGS84 coordinate system successively by the polar motion matrix, time angle matrix, nutation matrix, and precession matrix to obtain the position and velocity in the J2000 coordinate system. For detailed calculation procedures, please refer to [link to calculation details]. Figure 4 .

[0091] To verify the feasibility of the sun-synchronous orbit inclination adjustment method of the present invention, an example will be provided.

[0092] Taking a certain networking satellite launched in 2019 as an example, the satellite's target orbit was a sun-synchronous orbit at an altitude of 500km, with the descending node at 11:00 local time. The inclination of this sun-synchronous orbit at that altitude was 97.42°. Due to launch errors, the inclination was 97° upon insertion, differing from the standard orbit inclination by 0.42°. Four years later, an accident occurred... Figure 5 Changes in. Figure 5 This is a comparison diagram of the right ascension of the ascending node between the standard solar orbit and the actual solar orbit. Based on... Figure 5 It can be seen that over time, the right ascension of the ascending node of the actual orbit gradually deviates from that of the standard orbit.

[0093] Figure 6 This is a comparison diagram of the local time at the descending node of the standard solar orbit and the actual solar orbit. From Figure 6 It can be observed that the local time of the descending node in the standard sun-synchronous orbit remains consistently at 11:00 AM. However, the local time of the descending node in the satellite's actual orbit changes from 11:00 AM to 5:00 AM, indicating a significant deviation in the satellite's actual orbit relative to the standard orbit. Therefore, satellites with errors in their insertion inclination should correct these errors as soon as possible after insertion.

[0094] The revised plan is as follows:

[0095] Assuming the satellite uses chemical propulsion and thrust ( 4N, total satellite weight ( 70kg, and a correction was made 4 days after the satellite was launched into orbit. First, using the formula of this invention: The total velocity increment required to change the tilt angle is calculated to be 55.57 m / s, using the formula: The calculated total impulse is 3890 Ns. If each rail advances for 2 minutes, the formula is: The actual impulse generated by each ignition was calculated to be 354 Ns, and then the total number of rails was used to determine the impulse. The calculation formula was adjusted to require a total of 11 tracks.

[0096] Secondly, determine the location of the ignition center and the ignition time. The sub-satellite point of the ignition center is the equator, and ignition begins 60 seconds before passing the ignition center and ends 60 seconds after passing the ignition center.

[0097] Then according to ,

[0098] The ignition direction in the J2000 coordinate system can be calculated as: [0.0978961 -0.988066 -0.118922].

[0099] See Figure 7 As can be seen intuitively, after implementing the above strategy and making 11 ignition adjustments to the satellite, its inclination angle was adjusted to be almost the same as that of the sun-synchronous orbit.

[0100] See Figure 8 After the tilt adjustment, the local time at the descending node shifted by 1 hour within four years, a significant improvement compared to the 6-hour shift before the adjustment.

[0101] Therefore, it can be seen that after adjusting the sun-synchronous orbit inclination using the method of the present invention, the satellite's orbital offset is reduced, proving the feasibility of the present invention.

[0102] In another aspect, the present invention provides a sun-synchronous orbit inclination adjustment device. The sun-synchronous orbit inclination adjustment device of the present invention includes at least a memory, a processor, and a computer program stored in the memory and executable on the processor. The processor is configured to execute the sun-synchronous orbit inclination adjustment method described in any of the above embodiments based on the computer program stored in the memory.

[0103] The above embodiments can be combined with each other and have corresponding technical effects.

[0104] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method of adjusting the inclination of a sun-synchronous orbit for correcting the drift of the right ascension of the ascending node, characterized in that, At least comprising the following steps: Determine the position relationship between the initial orbit and the target orbit, so that the two intersection points of the initial orbit and the target orbit are located above the equator, and one of the intersection points is taken as the satellite orbit transfer point; Determining the orbital maneuver angle required to adjust a satellite from an initial orbital plane to a target orbital plane ; According to the system state and task target of the satellite, the single firing duration of the satellite is determined , and according to the single firing duration and the orbit change point, the satellite firing time and shutdown time are determined. According to the satellite's velocity at the initial orbit passing through the point of variable orbit and the satellite's velocity at the target orbit passing through the point of variable orbit , and the satellite's velocity at the target orbit passing through the point of variable orbit , the satellite's firing direction is determined. According to the speed of the satellite running on the initial orbit , the speed of running on the target orbit , and the satellite orbit change angle , the speed increment of the satellite single pulse firing is calculated , and the total impulse required for the satellite orbit change inclination is determined according to the speed increment ;​ According to total impulse And actual impulse produced per firing Satellite firing total orbit number is calculated analytically ; Based on the satellite variable orbit point, satellite ignition direction, satellite ignition time and shutdown time, the satellite is controlled to perform After the second ignition variable orbit, the inclination adjustment is completed.

2. The sun-synchronous orbit inclination adjustment method according to claim 1, characterized by, The single-shot firing time length The method for determining the firing time and the shutdown time of the satellite at the variable orbit point is specifically: Symmetrical to the orbit transfer point, start firing at a first position point before the satellite runs to the orbit transfer point, and end firing at a second position point after the satellite runs to the orbit transfer point; Wherein, the time length of the satellite running from the first position point to the orbit transfer point is equal to the time length of the satellite running from the orbit transfer point to the second position point.

3. The sun-synchronous orbit inclination adjustment method according to claim 2, characterized in that, The method for determining the satellite firing direction comprises the following steps of: determining the coordinate system expression of the satellite velocity at the initial orbit passing through the variable orbit point according to the satellite velocity at the initial orbit passing through the variable orbit point , and the coordinate system expression of the satellite velocity at the target orbit passing through the variable orbit point . Establish a J2000 coordinate system; the satellite running on the initial orbit passes the variable orbit point with a speed is expressed in the J2000 coordinate system as the satellite running on the target orbit passes the variable orbit point with a speed is expressed in the J2000 coordinate system as the firing direction of the satellite from the firing moment to the shutdown moment is expressed in the J2000 coordinate system as 。 4. The sun-synchronous orbit inclination adjustment method according to claim 2, characterized by, The method comprises the following steps of: calculating the speed of the satellite running on the initial orbit , the speed of the satellite running on the target orbit , and the satellite orbit transfer angle , and calculating the speed increment of the satellite single-pulse firing by analysis . The incremental pulse velocity required for the tilt adjustment is calculated according to the formula: ; and ​ wherein, theoretical inclination value corresponding to the target orbit height inclination value of the satellite actual telemetry data.

5. The sun-synchronous orbit inclination adjustment method according to claim 4, characterized in that, The method according to the speed increment Determining the total impulse required to change the inclination of a satellite orbit The method specifically comprises: The total impulse required for changing the inclination of the satellite orbit is calculated by the formula: where is the mass of the satellite.

6. The sun-synchronous orbit inclination adjustment method according to claim 5, characterized in that, The total impulse The actual impulse generated by each firing The total impulse of the satellite firing is calculated analytically The method is specifically: The actual impulse generated by each firing of the satellite is calculated by the formula: wherein is the thrust generated by the satellite firing, is the duration of a single firing of the satellite, is the firing efficiency of the satellite;​ The total number of orbits in which the satellite needs to be fired is calculated by the formula: .​ 7. The sun-synchronous orbit inclination adjustment method according to claim 2, characterized by, Satellite single firing duration No more than one third of the satellite orbit period.

8. The sun-synchronous orbit inclination adjustment method according to claim 3, characterized by, If the satellite carries a GNSS receiver, the output data is the position and velocity in the WGS84 coordinate system, then when collecting telemetry data, first convert the position and velocity in the WGS84 coordinate system to the position and velocity in the J2000 coordinate system; The specific method is: multiply the position and velocity in the WGS84 coordinate system by the polar motion matrix, the hour angle matrix, the nutation matrix and the precession matrix in turn to obtain the position and velocity in the J2000 coordinate system.

9. The sun-synchronous orbit inclination adjustment method according to claim 1, wherein, The total number of satellite firings is a natural number greater than or equal to 1.

10. A sun-synchronous orbit inclination adjustment device, characterized by, At least comprising a memory, a processor and a computer program stored in the memory and executable on the processor; the processor is configured to execute the sun-synchronous orbit inclination adjustment method according to any one of claims 1 to 9 based on the computer program stored in the memory.

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