An ultra-low orbit satellite orbit height autonomous maintenance method

Abstract

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

Description

Technical Field

[0001] This invention relates to the field of satellite orbital dynamics and control. More specifically, it relates to a method for autonomously maintaining the orbital altitude of an ultra-low Earth orbit satellite. Background Technology

[0002] Currently, ultra-low orbit (ULO) generally refers to orbits with an altitude of 150km-300km. Lowering satellites from the traditional 500km orbit to ULE while maintaining the same resolution performance can significantly reduce the weight and cost of optical and SAR remote sensing payloads, as well as launch costs. Furthermore, the reduced spatial radiation levels in ULE pave the way for more industrial products to be deployed in space. Additionally, the closer proximity to the ground reduces path loss, allowing for the use of smaller communication terminals.

[0003] Ultra-low Earth orbit (ULE) remains an untapped area, largely unexplored due to its inherent technological challenges. For example, under the same windward surface, the aerodynamic drag in ULE is more than 100 times that of a 500km orbit. Atmospheric density in ULE is dynamically fluctuating across a wide range due to factors such as solar radiation intensity, geomagnetic activity index, season, and diurnal variation, posing a severe challenge to ULE satellite altitude maintenance technology. The high atmospheric density and wide fluctuations caused by solar activity in ULE mean that traditional ground-based tracking and control stations are insufficient for maintaining satellite altitude. This typically requires orbital maneuver commands to be issued only within specific timeframes and regions that meet tracking and control constraints. This can disrupt the frequent application of altitude maintenance commands, impacting the satellite's on-orbit lifespan. Therefore, autonomous altitude maintenance strategies without ground intervention are crucial for ULE satellites.

[0004] To achieve high-precision orbit altitude maintenance, if the semi-major axis average orbit feature (referred to as the semi-major axis horizontal orbit value) is used as the input for orbit altitude maintenance, the semi-major axis horizontal orbit value needs to be determined with the highest possible accuracy. Using long-term GNSS data with short intervals for precise orbit determination can improve the accuracy of the semi-major axis horizontal orbit value determination, but the computational load is large, and it is difficult to implement autonomously on the satellite. Summary of the Invention

[0005] The purpose of this invention is to provide a method for autonomously maintaining the orbital altitude of ultra-low orbit satellites, so as to solve at least one of the problems existing in the prior art.

[0006] To achieve the above objectives, the present invention adopts the following technical solution:

[0007] The first aspect of this invention provides a method for autonomously maintaining the orbital altitude of an ultra-low Earth orbit satellite, comprising:

[0008] Using the position and velocity output from the satellite's onboard GNSS receiver as input, the actual value of the satellite's node period is calculated when the satellite is at its ascending or descending node.

[0009] Track maneuvers are performed based on the relationship between the actual value of the intersection period and the target value of the intersection period in order to maintain the track altitude.

[0010] Optionally, the step of performing orbital maneuvers based on the relationship between the actual value of the intersection period and the target value of the intersection period includes...

[0011] When the actual value of the intersection period is higher than the target value of the intersection period, no orbital maneuvering is applied in the next orbital period;

[0012] When the actual value of the intersection period is lower than the target value of the intersection period, and the difference between the target value of the intersection period and the actual value of the intersection period is less than the chemical propulsion start threshold, electric propulsion is used to maintain the orbital altitude and the start-up time is fixed in the next orbital period.

[0013] When the actual value of the intersection period is lower than the target value of the intersection period, and the difference between the target value of the intersection period and the actual value of the intersection period is greater than the chemical propulsion start threshold, chemical propulsion will be used to maintain the orbital altitude in the next orbital period, and the start-up time will be calculated in orbit.

[0014] Optionally, using the position and velocity output from the satellite's onboard GNSS receiver as input, the actual value of the satellite's node period is calculated when the satellite is at its ascending or descending node.

[0015] When the satellite passes through the ascending node for the first time, the first moment when the satellite passes through the ascending node is calculated based on the time before and after the satellite passes through the ascending node and the Z-axis component of the position vector in the J2000 coordinate system.

[0016] When the satellite passes through the ascending node for the second time, the second moment when the satellite passes through the ascending node is calculated based on the time before and after the sampled satellite passes through the ascending node and the Z-axis component of the position vector in the J2000 coordinate system.

[0017] Calculate the time difference between two consecutive passages of the satellite through the ascending node to obtain the actual value of the node period.

[0018] Optionally, in the first moment

[0019] Second moment

[0020] The actual value of the intersection period is T = t2 - t1.

[0021] Among them, t1 - The closest moment before the satellite's first passage through the ascending node, z1 - (z1 - <0) represents the z-axis component of the satellite's position in the J2000 coordinate system at the closest moment before the satellite's first passage through the ascending node, t1 +To determine the closest point after the satellite's first passage through the ascending node, z1 + (z1 + >0) represents the z-axis component of the satellite's position in the J2000 coordinate system at the closest moment after the satellite's first passage through the ascending node; t2 - This is the closest point before the satellite's second passage through the ascending node, z2 - (z2 - <0) represents the z-axis component of the satellite's position in the J2000 coordinate system at the closest moment before the satellite's second passage through the ascending node, t2 + To determine the closest point after the satellite's second passage through the rising node, z2 + (z2 + >0) represents the z-axis component of the satellite position in the J2000 coordinate system at the closest moment after the satellite passes through the ascending node for the second time.

[0022] Optionally, electric propulsion is used to maintain orbital altitude during the next orbital cycle, with a fixed operating duration.

[0023] Calculate the satellite's orbital features based on its position and velocity at the most recent time it passed through the ascending node, and obtain the perigee argument.

[0024] Based on the obtained perigee angle, calculate the start and end times of the electric propulsion operation;

[0025] The calculation of the start and end times of electric propulsion based on the obtained perigee angle includes...

[0026] (1) When the perigee is in the Southern Hemisphere, the start time t of electric propulsion. i and the end time t f They are respectively

[0027]

[0028]

[0029] (2) When the perigee is in the Northern Hemisphere, the start time t of electric propulsion. i and the end time t f They are respectively

[0030]

[0031]

[0032] Where, Δt ep The preset fixed power-on duration, n J It considers the average orbital angular velocity of the J2 term perturbation;

[0033]

[0034] In the formula, μ is the Earth's gravitational constant, μ = 3.986 × 10⁻⁶. 14 m 3 / s 2 ; a is the semi-major axis of the satellite orbit; the semi-major diameter of the satellite orbit p = a(1-e 2 e is the satellite orbital eccentricity; i is the satellite orbital inclination.

[0035] Optionally, chemical propulsion may be used to maintain orbital altitude during the next orbital period, including

[0036] Two chemical thrusts are performed to maintain orbital altitude, with a half-orbital period between the two chemical thrusts.

[0037] Optionally, the on-orbit computing power-on duration

[0038]

[0039] In the formula, m is the satellite mass, and μ is the Earth's gravitational constant μ = 3.986 × 10⁻⁶. 14 m 3 / s 2 F is the magnitude of chemical propulsion thrust, and T is the actual value of the intersection period. t The target value for the intersection period.

[0040] Optionally, the first chemical propulsion start time t is executed. 1i and the end time t 1f They are respectively

[0041]

[0042]

[0043] The second chemical propulsion operation begins at time t. 2i and the end time t 2f They are respectively

[0044]

[0045]

[0046] Where n J It takes into account the average satellite orbital angular velocity perturbed by the J2 term.

[0047] A second aspect of the present invention provides a computer device, including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the method provided in the first aspect of the present invention.

[0048] A third aspect of the present invention provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the method provided in the first aspect of the present invention.

[0049] The beneficial effects of this invention are as follows:

[0050] The present invention discloses an autonomous orbital altitude maintenance method for ultra-low orbit satellites, which uses the position and velocity output from a GNSS receiver as input and employs a combination of electric propulsion and chemical propulsion. It comprehensively utilizes the advantages of low propellant consumption of electric propulsion and high thrust of chemical propulsion to achieve autonomous orbital altitude maintenance in orbit. Attached Figure Description

[0051] The specific embodiments of the present invention will be described in further detail below with reference to the accompanying drawings.

[0052] Figure 1 An exemplary method flowchart is shown, in which an embodiment of the present invention can be applied.

[0053] Figure 2 This shows the time history of the satellite's orbital altitude during the maintenance process.

[0054] Figure 3 This shows the timeline of satellite fuel consumption during orbital altitude maintenance. Detailed Implementation

[0055] To more clearly illustrate the present invention, the following description is in conjunction with embodiments and appendices. Figure 1-3 The present invention will be further described below. Similar components in the accompanying drawings are indicated by the same reference numerals. Those skilled in the art should understand that the following detailed description is illustrative rather than restrictive and should not be construed as limiting the scope of protection of the present invention.

[0056] A method for autonomously maintaining the orbital altitude of an ultra-low Earth orbit satellite includes:

[0057] Using the position and velocity output from the satellite's onboard GNSS receiver as input, the actual value of the satellite's node period is calculated when the satellite is at its ascending or descending node.

[0058] Track maneuvers are performed based on the relationship between the actual value of the intersection period and the target value of the intersection period in order to maintain the track altitude.

[0059] In one embodiment, the present invention uses the period of the satellite passing through the ascending node as input and makes appropriate adjustments to the corresponding method to obtain an orbital altitude maintenance strategy using the period of the satellite passing through the descending node as input.

[0060] The method proposed in this invention uses only the satellite's position and velocity as input, and does not require high precision; the positioning accuracy of a typical GNSS receiver is sufficient. Moreover, it does not require high-frequency sampling of the satellite position data; only position data near the ascending node (or descending node) is needed as input for maintaining orbital altitude.

[0061] In one embodiment, performing orbital maneuvers based on the relationship between the actual value of the intersection period and the target value of the intersection period includes...

[0062] When the actual value of the intersection period is higher than the target value of the intersection period, no orbital maneuvering is applied in the next orbital period;

[0063] When the actual value of the intersection period is lower than the target value of the intersection period, and the difference between the target value of the intersection period and the actual value of the intersection period is less than the chemical propulsion start threshold, electric propulsion is used to maintain the orbital altitude and the start-up time is fixed in the next orbital period.

[0064] When the actual value of the intersection period is lower than the target value of the intersection period, and the difference between the target value of the intersection period and the actual value of the intersection period is greater than the chemical propulsion start threshold, chemical propulsion will be used to maintain the orbital altitude in the next orbital period and the start-up time will be calculated in orbit.

[0065] In one embodiment, using the position and velocity output from the satellite's onboard GNSS receiver as input, the actual value of the satellite's node period is calculated when the satellite is at its ascending or descending node.

[0066] When the satellite passes through the ascending node for the first time, the first moment when the satellite passes through the ascending node is calculated based on the time before and after the satellite passes through the ascending node and the Z-axis component of the position vector in the J2000 coordinate system.

[0067] When the satellite passes through the ascending node for the second time, the second moment when the satellite passes through the ascending node is calculated based on the time before and after the sampled satellite passes through the ascending node and the Z-axis component of the position vector in the J2000 coordinate system.

[0068] Calculate the time difference between two consecutive passages of the satellite through the ascending node to obtain the actual value of the node period.

[0069] In one embodiment,

[0070] First moment

[0071] Second moment

[0072] The actual value of the intersection period is T = t2 - t1.

[0073] Among them, t1 - The closest moment before the satellite's first passage through the ascending node, z1 - (z1- <0) represents the z-axis component of the satellite's position in the J2000 coordinate system at the closest moment before the satellite's first passage through the ascending node, t1 + To determine the closest point after the satellite's first passage through the ascending node, z1 + (z1 + >0) represents the z-axis component of the satellite's position in the J2000 coordinate system at the closest moment after the satellite's first passage through the ascending node; t2 - This is the closest point before the satellite's second passage through the ascending node, z2 - (z2 - <0) represents the z-axis component of the satellite's position in the J2000 coordinate system at the closest moment before the satellite's second passage through the ascending node, t2 + To determine the closest point after the satellite's second passage through the rising node, z2 + (z2 + >0) represents the z-axis component of the satellite position in the J2000 coordinate system at the closest moment after the satellite passes through the ascending node for the second time.

[0074] In one embodiment, maintaining orbital altitude using electric propulsion for a fixed duration during the next orbital cycle includes...

[0075] Calculate the satellite's orbital features based on its position and velocity at the most recent time it passed through the ascending node, and obtain the perigee argument.

[0076] Based on the obtained perigee angle, calculate the start and end times of the electric propulsion operation;

[0077] The calculation of the start and end times of electric propulsion based on the obtained perigee angle includes...

[0078] (1) When the perigee is in the Southern Hemisphere, the start time t of electric propulsion. i and the end time t f They are respectively

[0079]

[0080]

[0081] (2) When the perigee is in the Northern Hemisphere, the start time t of electric propulsion. i and the end time t f They are respectively

[0082]

[0083]

[0084] Where, Δt ep The preset fixed power-on duration, nJ It takes into account the average satellite orbital angular velocity perturbed by the J2 term;

[0085]

[0086] In the formula, μ is the Earth's gravitational constant, μ = 3.986 × 10⁻⁶. 14 m 3 / s 2 ; a is the semi-major axis of the satellite orbit; the semi-major diameter of the satellite orbit p = a(1-e 2 e is the satellite orbital eccentricity; i is the satellite orbital inclination.

[0087] In one embodiment, using chemical propulsion to maintain orbital altitude during the next orbital period includes performing two chemical propulsion maneuvers to maintain orbital altitude, with a half-orbital period between the two chemical propulsion maneuvers.

[0088] In one embodiment, the on-orbit computing power-on duration

[0089]

[0090] In the formula, m is the satellite mass, and μ is the Earth's gravitational constant μ = 3.986 × 10⁻⁶. 14 m 3 / s 2 F is the magnitude of chemical propulsion thrust, and T is the actual value of the intersection period. t The target value for the intersection period.

[0091] Optionally, the start time t of the first chemical propulsion is... 1i and the end time t 1f They are respectively

[0092]

[0093]

[0094] The start time t of the second chemical propulsion 2i and the end time t 2f They are respectively

[0095]

[0096]

[0097] Where n J It takes into account the average satellite orbital angular velocity perturbed by the J2 term.

[0098] This invention uses the satellite's position and velocity as input to determine the actual value of the intersection period, and determines the semi-major axis change for maintaining orbital altitude based on the relationship between the actual value of the intersection period and the target value of the intersection period. Furthermore, it designs an orbital altitude maintenance strategy that combines electric propulsion and chemical propulsion, which can take advantage of the advantages of low propellant consumption in electric propulsion and high thrust in chemical propulsion.

[0099] One specific embodiment,

[0100] The target orbital altitude for orbital altitude maintenance is 220km, with an orbital maintenance accuracy of ±1km.

[0101] The target value T of the intersection period t The orbital period, T, is set to the altitude of 220.9 km. t = 5342.55 sec; the time difference Δt between the chemical propulsion start-up thresholds c_th The value corresponding to a change in altitude of 1.5 km is taken as Δt. c_th = 1.82 sec.

[0102] The initial orbital altitude of the ultra-low Earth orbit satellite is 220.8 km, the initial mass of the satellite is taken as 200 kg, and the windward area is taken as 0.4 m². 2 The thrust amplitude of electric propulsion is set at 12 mN, and the specific impulse at 1200 sec; the thrust amplitude of chemical propulsion is set at 2 N, and the specific impulse at 220 sec. The operating time of electric propulsion within one orbital period is set at 720 sec. The three-axis position determination error of the spaceborne GNSS receiver is set at 10 m (1σ), and the three-axis velocity determination error is set at 0.1 m / s (1σ).

[0103] To verify the orbital altitude maintenance effect under the real Earth model, the simulated gravitational model includes the 21st order Earth gravitational field, atmospheric drag, gravitational force of the Sun, Moon and Third Body, and solar radiation pressure. The atmospheric drag model is the NRLMSISE model, and the atmospheric parameters are Cd = 3 and F107 = 150.

[0104] The method described in this invention is used to simulate the orbital altitude maintenance of a low Earth orbit satellite. Table 1 shows the simulation results of 50 orbital cycles obtained by the method described in this invention. Table 1 includes the intersection time, the actual value of the intersection cycle, the deviation between the target value and the actual value of the intersection cycle, and whether an orbital maneuver is applied in the next orbital cycle.

[0105] Based on the simulation results in Table 1, it can be concluded that within 50 orbital cycles, chemical propulsion orbital maneuvers are required for 3 orbital cycles, and electric propulsion orbital maneuvers are required for the remaining orbital cycles.

[0106] Figure 2This is the time history of the satellite's orbital altitude during the orbital altitude maintenance process. The results show that maintaining the orbital altitude within the range of 219 km to 221 km meets the orbital altitude maintenance target (220 ± 1 km).

[0107] Figure 3 This is the time history of satellite fuel consumption during orbital altitude maintenance. The results show that the mass changes slowly during electric propulsion and rapidly during chemical propulsion, thus combining the advantages of low propellant consumption in electric propulsion and high thrust in chemical propulsion, thereby reducing thruster propellant consumption.

[0108] Table 1 Summary of Simulation Results

[0109]

[0110] Continued table

[0111]

[0112] In the description of this invention, it should be noted that the terms "upper," "lower," etc., indicating the orientation or positional relationship are based on the orientation or positional relationship shown in the accompanying drawings, and are only for the convenience of describing the invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of the invention. Unless otherwise expressly specified and limited, the terms "installed," "connected," and "linked" 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 mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication between two elements. For those skilled in the art, the specific meaning of the above terms in this invention can be understood according to the specific circumstances.

[0113] It should also be noted that in the description of this invention, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitation, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0114] Obviously, the above embodiments of the present invention are merely examples for clearly illustrating the present invention, and are not intended to limit the implementation of the present invention. For those skilled in the art, other variations or modifications can be made based on the above description. It is impossible to exhaustively list all the implementation methods here. All obvious variations or modifications derived from the technical solutions of the present invention are still within the protection scope of the present invention.

Claims

1. A method for autonomously maintaining the orbital altitude of an ultra-low Earth orbit satellite, characterized in that, include: Using the position and velocity output from the satellite's onboard GNSS receiver as input, the actual value of the satellite's node period is calculated when the satellite is at its ascending or descending node. Track maneuvers are performed based on the relationship between the actual value of the intersection period and the target value of the intersection period in order to maintain the track altitude; The process of performing orbital maneuvers based on the relationship between the actual value and the target value of the intersection period includes... When the actual value of the intersection period is higher than the target value of the intersection period, no orbital maneuvering is applied in the next orbital period; When the actual value of the intersection period is lower than the target value of the intersection period, and the difference between the target value of the intersection period and the actual value of the intersection period is less than the chemical propulsion start threshold, electric propulsion is used to maintain the orbital altitude and the start-up time is fixed in the next orbital period. When the actual value of the intersection period is lower than the target value of the intersection period, and the difference between the target value of the intersection period and the actual value of the intersection period is greater than the chemical propulsion start threshold, chemical propulsion will be used to maintain the orbital altitude in the next orbital period and the start-up time will be calculated in orbit. Using the position and velocity output from the satellite's onboard GNSS receiver as input, the actual value of the satellite's node period is calculated when the satellite is at its ascending or descending node. When the satellite passes through the ascending node for the first time, the first moment when the satellite passes through the ascending node is calculated based on the time before and after the satellite passes through the ascending node and the Z-axis component of the position vector in the J2000 coordinate system. When the satellite passes through the ascending node for the second time, the second moment when the satellite passes through the ascending node is calculated based on the time before and after the sampled satellite passes through the ascending node and the Z-axis component of the position vector in the J2000 coordinate system. Calculate the time difference between two consecutive passages of the satellite through the ascending node to obtain the actual value of the node period; First moment Second moment Actual value of intersection period , in, This is the closest moment before the satellite first passes through the ascending node. The z-axis component of the satellite's position in the J2000 coordinate system corresponds to the closest moment before the satellite's first passage through the ascending node. , To determine the closest point after the satellite's first passage through the ascending node, The z-axis component of the satellite position in the J2000 coordinate system corresponds to the closest moment after the satellite's first passage through the ascending node. ; This is the closest point before the satellite passes through the ascending node for the second time. The z-axis component of the satellite position in the J2000 coordinate system corresponds to the closest moment before the satellite's second passage through the ascending node. , To determine the closest point after the satellite's second passage through the ascending node, The z-axis component of the satellite's position in the J2000 coordinate system corresponds to the closest moment after the satellite's second passage through the ascending node. .

2. The method according to claim 1, characterized in that, During the next orbital cycle, electric propulsion is used to maintain orbital altitude for a fixed duration, including... Calculate the satellite's orbital features based on its position and velocity at the most recent time it passed through the ascending node, and obtain the perigee argument. Based on the obtained perigee angle, calculate the start and end times of the electric propulsion operation; The calculation of the start and end times of electric propulsion based on the obtained perigee angle includes... (1) When the perigee is in the Southern Hemisphere, the start time of electric propulsion. and the end time They are respectively (2) When the perigee is in the Northern Hemisphere, the start time of electric propulsion. and the end time They are respectively in, The preset fixed power-on duration, It takes into account the average satellite orbital angular velocity perturbed by the J2 term; In the formula, , It is the Earth's gravitational constant. ; The semi-major axis of the satellite orbit; the semi-circular diameter of the satellite orbit. ; For satellite orbital eccentricity; This refers to the satellite's orbital inclination.

3. The method according to claim 2, characterized in that, Maintaining orbital altitude using chemical propulsion during the next orbital period includes Two chemical thrusts are performed to maintain orbital altitude, with a half-orbital period between the two chemical thrusts.

4. The method according to claim 3, characterized in that, The on-orbit computing power-on duration In the formula, It's about satellite quality. It is the Earth's gravitational constant. , This represents the chemical propulsion thrust amplitude, where T is the actual value of the intersection period. The target value for the intersection period.

5. The method according to claim 4, characterized in that, The start of the first chemical propulsion mission and the end time They are respectively The start of the second chemical propulsion mission and the end time They are respectively in It takes into account the average satellite orbital angular velocity perturbed by the J2 term.

6. A computer device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the program, it implements the method as described in any one of claims 1-5.

7. A computer-readable storage medium having a computer program stored thereon, characterized in that, When the program is executed by the processor, it implements the method as described in any one of claims 1-5.

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