Low earth orbit satellite design method and system with normal attitude keeping function

Abstract

The invention provides a low-earth-orbit satellite design method and system with a normal attitude keeping function. The method comprises the following steps: setting basic parameters of a low-earth-orbit satellite; based on the basic parameters, a required low-orbit satellite model is obtained through satellite coordinate system design, satellite layout design and satellite system design; and carrying out rolling angle design on the attitude of the satellite orbit based on the low-orbit satellite model so as to maintain the normal attitude of the low-orbit satellite. The method reduces the balancing difficulty of the center of mass of the whole satellite, is easy to control the inertia of the whole satellite, and is friendly to the attitude control of the satellite; irreversible damage to a camera mounted on the ground can be avoided by means of heat generated by earth reflection and infrared radiation; the design difficulty of the sailboard can be reduced, meanwhile, the expenditure for installing SADA to obtain energy is reduced, and the design complexity of the whole satellite is reduced; a data transmission antenna axially points to the earth center, and communication with a data transmission station on the ground is achieved in a phased array electric scanning mode.

Description

Technical Field

[0001] The present invention relates to the field of satellite design technology, and in particular to a design method and system for a low-orbit satellite with a normal attitude maintenance function based on mission, structure, energy, and data transmission coupling, as well as a corresponding computer terminal and computer-readable storage medium. Background Art

[0002] With the rise of commercial spaceflight in recent years, the cost control requirements for low-orbit satellites have become extremely stringent, and therefore higher requirements are placed on the design of satellites. The coupling design of the various subsystems that constitute the satellite must be comprehensively considered to achieve the optimal overall design.

[0003] At a time when commercial spaceflight is booming, the need for satellite sail panels to be as small as possible is driven by low-cost design concepts and the long lifespan of low-orbit satellites. Small sail panels naturally save on the high cost of solar cells, while ensuring a small windward surface that reduces the impact of atmospheric drag on the orbital altitude attenuation of high-speed satellites. To fully capture energy, after the satellite enters orbit and derotates in its normal attitude (eliminating the satellite's angular velocity caused by the launch and release of the satellite), the attitude is adjusted to face the sun normal to fully capture energy. Multifunctional satellites integrating communications and remote sensing need to carry laser communication payloads and visible light camera payloads. The operating conditions and temperature control accuracy of these two payloads are closely related, and they require significant energy support during operation. Therefore, the coupling between mission, structure, thermal control, and energy is very strong. Attitude adjustments due to the need for energy capture will also affect ground communications for measurement, control, and data transmission, leading to a coupled design of the entire satellite design system.

[0004] However, existing satellite design solutions usually have the following technical problems:

[0005] Installing one-dimensional or even two-dimensional SADA without considering the cost will lead to complex control, increase weight and lose attitude control accuracy; if SADA is not installed, the traditional ground orientation method will result in insufficient energy required for sun-synchronous orbit satellites in near-meridian orbits, and the solar orientation method will cause the camera lens of the remote sensing camera to face the cold sky for 1 / 3 of the time, resulting in increased energy demand and also causing the problem of excessive energy consumption caused by reaction wheel maneuvers. Summary of the Invention

[0006] In response to the above-mentioned deficiencies in the prior art, the present invention provides a method and system for designing a low-orbit satellite with a normal attitude maintenance function based on mission, structure, energy, and data transmission coupling, and also involves a corresponding computer terminal and computer-readable storage medium.

[0007] According to one aspect of the present invention, a method for designing a low-orbit satellite with a normal attitude maintenance function is provided, comprising:

[0008] Set the basic parameters of low-orbit satellites;

[0009] Based on the basic parameters, the required low-orbit satellite model is obtained through satellite coordinate system design, satellite layout design and satellite system design;

[0010] Based on the low-orbit satellite model, the roll angle of the satellite orbit attitude is designed to achieve normal attitude maintenance of the low-orbit satellite.

[0011] According to another aspect of the present invention, a low-orbit satellite design system with a normal attitude maintenance function is provided, comprising:

[0012] Constraint design module, which is used to set the basic parameters of low-orbit satellites;

[0013] A satellite design module, which obtains the required low-orbit satellite model through satellite coordinate system design, satellite layout design, and satellite system design based on the basic parameters;

[0014] The attitude roll design module designs the roll angle of the satellite orbit attitude based on the low-orbit satellite model, so as to achieve normal attitude maintenance of the low-orbit satellite.

[0015] According to a third aspect of the present invention, a computer terminal is provided, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, the terminal can be used to execute the method described above in the present invention, or to execute the system described above in the present invention.

[0016] According to a fourth aspect of the present invention, a computer-readable storage medium is provided, on which a computer program is stored. When the computer program is executed by a processor, it can be used to execute the method described above in the present invention, or to run the system described above in the present invention.

[0017] Due to the adoption of the above technical solution, the present invention has at least one of the following beneficial effects compared with the prior art:

[0018] The present invention adopts symmetrical sailboard installation, which reduces the difficulty of balancing the center of mass of the entire satellite, makes the inertia of the entire satellite easy to control, is friendly to the attitude control of the satellite, and improves the attitude control accuracy of satellites of the same scale.

[0019] Through satellite system design, the present invention can avoid irreversible damage to ground-mounted cameras by utilizing the heat generated by Earth's reflection and infrared radiation when the on-board heater hardware is knocked over by a single particle, thus improving the reliability of the payload and, in turn, the reliability of the entire satellite system mission.

[0020] The present invention can reduce the difficulty of sailboard design, while reducing the cost of installing SADA to obtain energy and reducing the complexity of the entire satellite design, providing design ideas and engineering reference experience for commercial low-cost satellites.

[0021] The present invention can take into account both ground transmission of data transmission: the data transmission antenna is axially pointed to the center of the earth, and communicates with the ground data transmission station through phased array electronic scanning. Based on the small angle adjustment of -26° relative to the ground orientation based on the attitude roll, it does not affect the pointing direction of the data transmission phased array, and can well complete the attitude compatibility design. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Other features, objects and advantages of the present invention will become more apparent upon reading the detailed description of non-limiting embodiments with reference to the following drawings:

[0023] Figure 1 The figure is a workflow diagram of a method for designing a low-orbit satellite with a normal attitude maintenance function according to an embodiment of the present invention.

[0024] Figure 2 Schematic diagram of the components of a system designed for a low-orbit satellite with a normal attitude maintenance function in one embodiment of the present invention.

[0025] Figure 3 Schematic diagram of satellite launch state coordinates in a preferred embodiment of the present invention; wherein (a) is the direction of the remote sensing camera (+Z), and (b) is the direction of the sailboard installation surface (-Z).

[0026] Figure 4 This is a schematic diagram of the layout of a single satellite unit outside the satellite in a preferred embodiment of the present invention.

[0027] Figure 5 This is a second schematic diagram of the satellite off-site stand-alone layout in a preferred embodiment of the present invention.

[0028] Figure 6 Schematic diagram of the third satellite off-site stand-alone layout in a preferred embodiment of the present invention.

[0029] Figure 7 Schematic diagram of the fourth satellite off-site stand-alone layout in a preferred embodiment of the present invention.

[0030] Figure 8 Schematic diagram of satellite orientation to the earth in a preferred embodiment of the present invention.

[0031] Figure 9 Schematic diagram of two-dimensional satellite orientation to the sun (sunlit area) in a preferred embodiment of the present invention.

[0032] Figure 10 Schematic diagram of two-dimensional satellite orientation toward the sun in a preferred embodiment of the present invention (shaded area).

[0033] Figure 11 This is a diagram showing the angle between the camera axis in satellite 2 orbit and the vector from the satellite to the center of the earth, as well as the change in light intensity in a preferred embodiment of the present invention.

[0034] Figure 12 This is a diagram showing the change in the satellite's orbital solar angle Beta within one year in a preferred embodiment of the present invention.

[0035] Figure 13 This is a schematic diagram of the oblique installation of satellite sailboards in a preferred embodiment of the present invention.

[0036] Figure 14 Schematic diagram of satellite attitude roll orientation in a preferred embodiment of the present invention.

[0037] Figure 15 1 is a graph showing the insolation factor cosθ and the incident angle when the satellite is rolling at a -26° attitude in a preferred embodiment of the present invention.

[0038] Figure 16 This is a curve showing the load current of about 2.3A on average (2:25:37 to 12:35:54 on February 4) when the satellite is rolling at -26° in a preferred embodiment of the present invention.

[0039] Figure 17 This is a curve diagram of the bus voltage (2:25:37 to 12:35:54 on February 4) with an average of about 28V when the satellite rolls at a -26° attitude in a preferred embodiment of the present invention.

[0040] Figure 18 This is a curve diagram of the load current (0:25:37-10:35:54 on March 18) under two-dimensional satellite solar orientation in a preferred embodiment of the present invention, with an average of approximately 3.22A.

[0041] Figure 19 This is a curve diagram of the bus voltage (0:25:37 to 10:35:54 on February 4) under two-dimensional satellite solar orientation in a preferred embodiment of the present invention, with an average of about 28V.

[0042] Figure 20 This is a flow chart for calculating the orbital solar angle Beta in a preferred embodiment of the present invention.

[0043] In the figure, 1 is the space tracking and control transceiver antenna, 2 is the propulsion module, 3 is the navigation short message transceiver antenna, 4 is the ground tracking and control transceiver antenna, 5 is the phased array antenna, 6 is the camera, 7 is the laser optical machine, 8 is the star sensor, 9 is the solar cell array, 10 is the satellite and rocket release mechanism, and 11 is the laser communication payload optical machine equipment. DETAILED DESCRIPTION

[0044] The following is a detailed description of an embodiment of the present invention. This embodiment is implemented based on the technical solution of the present invention, and provides a detailed implementation method and specific operation process. It should be noted that those skilled in the art may make various modifications and improvements without departing from the scope of the present invention, and these modifications and improvements fall within the scope of protection of the present invention.

[0045] In response to the technical problems existing in the prior art, an embodiment of the present invention provides a low-orbit satellite design method with a normal attitude maintenance function. The method is based on mission, structure, energy, and data transmission coupling to achieve an overall optimal satellite design engineering implementation plan based on a low-cost design concept.

[0046] Specifically, if Figure 1 As shown, the method for designing a low-orbit satellite with a normal attitude maintenance function provided by this embodiment may include:

[0047] S1, set the basic parameters of the low-orbit satellite;

[0048] S2, based on basic parameters, obtains the required low-orbit satellite model through satellite coordinate system design, satellite layout design, and satellite system design;

[0049] S3, based on the low-orbit satellite model, designs the roll angle of the satellite orbit attitude to achieve normal attitude maintenance of the low-orbit satellite.

[0050] In some preferred embodiments, the above S1, setting basic parameters of the low-orbit satellite, may further include:

[0051] The total weight of the satellite is set to not exceed 100kg;

[0052] Set a sun-synchronous circular orbit with an altitude of 510 km and a descending node time of 10:00-10:30 AM local time.

[0053] Set the visible light remote sensing camera startup temperature: 5℃~35℃, operating temperature: 20℃~25℃;

[0054] Set the laser load starting temperature: 0℃~35℃, and the operating temperature: 20℃~25℃.

[0055] In some preferred embodiments, the above S2, satellite coordinate system design, may further include:

[0056] S211, define the satellite structure coordinate system as Os-XsYsZs, where:

[0057] Os is the center point of the satellite-rocket separation plane, that is, the origin;

[0058] The Xs axis is perpendicular to the separation plane of the satellite and rocket, and points to the propulsion module along the short side of the satellite, corresponding to the flight direction under the premise of stable attitude towards the earth on orbit;

[0059] The Ys axis is parallel to the separation plane of the rocket and satellite, and points to the direction of the laser payload along the long side of the release mechanism, corresponding to the negative normal of the orbital plane under the premise of stable attitude towards the earth on orbit;

[0060] The Zs axis is orthogonal to Xs and Zs in the right-hand coordinate system, and corresponds to the direction of the center of the earth under the premise of stable attitude on orbit;

[0061] S212, translate the satellite structure coordinate system to the satellite center of mass, and obtain the satellite body coordinate system Ob—XbYbZb, abbreviated as O—XYZ.

[0062] like Figure 3 As shown in (a) and (b).

[0063] In some preferred embodiments, the above S2, satellite layout design, may further include:

[0064] S221, integrated design of remote sensing camera payload;

[0065] S222, the entire star configuration is designed to be a long L-shaped configuration;

[0066] S223, the satellite is designed to be a flat and long strip configuration. The satellite solar array adopts a single-wing secondary expansion structure. After the satellite enters orbit and is unlocked by pyrotechnics, the middle plate is expanded outward 180° and then locked. At the same time, the two outer plates are also expanded outward 180° and then locked.

[0067] In some preferred embodiments, the above S221, integrating the remote sensing camera payload, may further include:

[0068] Given the satellite's largest size and heaviest weight, the remote sensing camera payload necessitates a central design for the overall satellite layout. The satellite itself measures 470mm (X) × 960mm (Y) × 660mm (Z), while the remote sensing camera payload measures Φ390 (X and Y directions) × 750 (Z direction). This indicates that the remote sensing camera payload extends beyond the satellite itself in the Z direction, occupying 83% in the X direction and 41% in the Y direction. The remote sensing camera payload weighs 13kg, representing 14% of the total satellite weight. The remote sensing camera payload consists of a camera lens and electronics. The camera lens is exposed, while the electronics are located inside the satellite.

[0069] In some preferred embodiments, the above S222, designing the entire star configuration into a long L-shaped configuration, may further include:

[0070] The entire satellite is designed as a three-cabin structure, in which the integrated electronic cabin and the remote sensing camera payload cabin have the same size in the Z direction, and the laser communication payload cabin has a size of 61% of the Z direction, and is L-shaped when viewed along the Y direction. Figure 4 The remote sensing camera payload compartment contains its electronic components and the non-exposed camera barrel.

[0071] In some preferred embodiments, the above S223, wherein the satellite is designed to be a flat long strip, and the satellite solar array adopts a single-wing double-expanding structure, may further include:

[0072] The active launch envelope of the satellite measures 700mm (X) × 1200mm (Y) × 930mm (Z), with the largest dimension in the Y direction. The release and separation mechanism is installed on the -X plane (i.e., the YOZ plane, with the positive direction being the -X direction). This ensures that the largest YOZ plane fits neatly within the launch fairing, fully utilizing the launcher's external space. To conserve launch space and prevent the sail panels from obstructing the laser communication payload's field of view after orbit, the satellite's solar array utilizes a single-wing, double-expandable structure, ensuring that the stowed state remains within the launcher's envelope.

[0073] In some preferred embodiments, the above S2, satellite system design, may further include any one of the following:

[0074] S231, unfolded sailboard tilted -26°;

[0075] S232, windsurfing suit, two-dimensional sun-fighting during no mission;

[0076] S233, windsurfing suit, attitude roll -26° during non-mission period.

[0077] In some preferred embodiments, the above-mentioned S231, unfolding the sailboard and tilting it at -26°, may further include:

[0078] Based on the installation of the sailboard at a -26° tilt, under the premise of 10:00-10:30 local time at the descending node, the normal direction of the satellite sailboard can be as close as possible to the sun on the left side of the flight direction (within the -Y half plane) to ensure energy acquisition. However, compared with the other two designs, this design has a higher structural design difficulty, and the deployed sailboard area is only about 1 / 2 of the body, which has a smaller contribution to energy. Figure 13 shown.

[0079] In some preferred embodiments, the above-mentioned S232, windsurfing upright, two-dimensionally facing the sun during no mission period, may further include:

[0080] The surface of the sailboard is parallel to the XOY plane, and the sailboard normal (-Z direction) points to the sun during the no-load mission to ensure maximum energy acquisition efficiency. However, in this solution, the relationship between the camera and the earth changes periodically within one orbit, which will cause the camera to be exposed to cold air in the shadow area, which will require a large amount of heating power for the camera. At the same time, frequent attitude maneuvers may damage the life of the reaction wheel and increase the burden of energy demand. Figure 11 As shown in the figure, the angle between the camera axis and the vector from the satellite to the center of the earth in the second orbit of the satellite and the curve of the change of the light intensity are given.

[0081] In some preferred embodiments, the above S233, wherein the windsurfing board is mounted upright and the attitude roll is -26° during no mission, may further include:

[0082] Under the premise of a traditional symmetrical design of the entire satellite's sail panels, a -26° attitude roll on orbit achieves energy acquisition between the two-dimensional solar and Earth-facing positions, and the entire satellite's energy consumption also falls between the two positions, maintaining a stable on-orbit attitude while reducing the difficulty of the entire satellite's structural design. Therefore, this design is the optimal one.

[0083] In some preferred embodiments, the above S3, based on the low-orbit satellite model, designs the roll angle of the satellite orbit attitude, and may further include:

[0084] S31, calculates the orbital solar angle Beta based on the satellite's on-orbit flight conditions;

[0085] S32, performing initial default parameter settings according to the recommended roll angle value.

[0086] In some preferred embodiments, the above S31, calculating the orbital solar angle Beta based on the satellite's on-orbit flight conditions, may further include:

[0087] S311: Determine the initial orbit insertion parameters based on the designed low-orbit satellite model;

[0088] S312, determine whether it is a sun-synchronous orbit; if so, calculate the Beta angle for a set period of time (e.g., 1 year) and jump to S314; if not, obtain the short period T of the Beta angle change through theoretical calculation and simulation, and continue to execute the following steps;

[0089] S313, calculating the Beta angle within a short period T and the Beta angle within a set period (e.g., within 1 year);

[0090] S314 , providing recommended values of the rolling angle at different times within a set period of time (eg, within 1 year) in intervals.

[0091] In some preferred embodiments, the above S32, performing initial default parameter setting according to the recommended roll angle value, may further include:

[0092] For the attitude pointing in the non-mission phase, based on the orbital solar angle Beta calculated by S31, the most appropriate value for the orbit entry time is selected from the system default guidance parameters as the initial default parameter, and is set to be stored in a form that can be configured through ground commands.

[0093] Based on the same inventive concept, an embodiment of the present invention further provides a low-orbit satellite design system with a normal attitude maintenance function.

[0094] like Figure 2 As shown, the low-orbit satellite design system with a normal attitude maintenance function provided by this embodiment may include:

[0095] Constraint design module, which is used to set the basic parameters of low-orbit satellites;

[0096] Satellite design module, which obtains the required low-orbit satellite model based on basic parameters through satellite coordinate system design, satellite layout design and satellite system design;

[0097] The attitude roll design module is based on the low-orbit satellite model and designs the roll angle of the satellite orbit attitude to achieve normal attitude maintenance of the low-orbit satellite.

[0098] It should be noted that the steps in the method provided by the present invention can be implemented using corresponding modules, devices, units, etc. in the system. Those skilled in the art can refer to the technical solution of the method to implement the composition of the system, that is, the embodiments in the method can be understood as preferred examples of constructing the system, which will not be elaborated here.

[0099] The working contents of the functional modules constituting the system provided by the above embodiment of the present invention are further described below.

[0100] 1. Constraint design module. The main design constraints of this module are as follows:

[0101] 1. Constraint 1: The total weight of the satellite shall not exceed 100 kg;

[0102] 2. Constraint 2: Sun-synchronous circular orbit, orbital altitude 510 km, descending node 10:00-10:30 AM local time;

[0103] 3. Constraint 3: Visible light remote sensing camera startup temperature: 5℃~35℃, operating temperature: 20℃~25℃;

[0104] 4. Constraint 4: Laser payload starting temperature: 0℃~35℃, operating temperature: 20℃~25℃.

[0105] 2. Satellite design module: This module mainly includes the following functional units:

[0106] 1. Satellite coordinate system design unit, used for coordinate system definition:

[0107] The satellite structure coordinate system (Os—XsYsZs) is a reference coordinate system fixed to the satellite body with the center point of the satellite-rocket separation plane as its origin. The three axes are defined as follows:

[0108] (1) Xs axis—perpendicular to the separation plane, along the short side of the satellite, pointing toward the propulsion module, corresponding to the flight direction under the premise of stable attitude toward the Earth on orbit;

[0109] (2) Ys axis - parallel to the separation plane of the rocket and satellite, pointing to the direction of the laser payload along the long side of the release mechanism, corresponding to the negative normal of the orbital plane under the premise of stable attitude towards the earth on orbit;

[0110] (3) Zs axis - orthogonal to Xs and Zs in the right-hand coordinate system, corresponding to the direction of the center of the earth under the premise of stable attitude relative to the earth on orbit.

[0111] The satellite structure coordinate system is translated to the satellite mass center, which is the satellite body coordinate system (Ob—XbYbZb), which is abbreviated as O—XYZ in this paper. Figure 3 As shown in (a) and (b).

[0112] 2. Satellite layout design unit, used for satellite layout design:

[0113] (1) The remote sensing camera payload is the largest and heaviest payload on the satellite, so the structural configuration is prioritized for integrated design around it to improve the integration of the entire satellite.

[0114] (2) The laser communication payload is equipped with a two-dimensional turntable with an azimuth rotation range of 0±120° and a pitch angle of -30° to +60°. The obstruction area is minimized within the rotation range. To avoid the payload field of view, a long L-shaped configuration is preferred in the entire satellite configuration to minimize field of view obstruction.

[0115] (3) In order to meet the requirements of launching multiple satellites with one rocket, the satellite adopts a flat and long strip configuration, making full use of the effective space between the fairing and the adapter. The satellite solar array adopts a single-wing secondary expansion design. After the satellite enters orbit, the middle plate of the solar array is expanded outward 180° and then locked after the pyrotechnic unlocking. At the same time, the two outer plates are also expanded outward 180° and then locked. Figures 4 to 7 shown.

[0116] It should be noted that: Figure 7 The medium phased array antenna is a data transmission antenna used to transmit mission data to the ground.

[0117] 3. System design unit, used for satellite system design:

[0118] A diagram of a satellite's attitude toward the Earth, such as Figure 8 shown.

[0119] Depend on Figure 8 It can be seen that in the geodirectional attitude, the external heat flow generated by the earth's reflection and infrared radiation has a heating effect on the visible light remote sensing camera installed on the +Z surface, which will save the power consumption required for camera heating.

[0120] A two-dimensional diagram of a satellite's attitude toward the Sun, such as Figure 9 and Figure 10 shown.

[0121] Depend on Figure 9 and Figure 10 It can be seen that in the two-dimensional solar attitude, the -Z axis (normal to the sailboard) points to the solar vector, and the relationship between the camera and the Earth changes periodically within one orbit, which will cause the camera to face the cold sky in the shadow area, and the heating power consumption demand of the camera is relatively large. Figure 11 The angle between the camera axis and the vector from the satellite to the center of the earth and the curve of light intensity variation in the satellite's two orbits are given.

[0122] According to the second constraint "descending node local time 10:00~10:30AM", the relationship between orbital solar angle Beta and time can be obtained as follows: Figure 12 shown.

[0123] Figure 12 The orbital solar angle, Beta, is defined as the angle between the sun vector and the orbital plane. It is positive if the sun and the orbital plane normal (the direction of the orbital angular velocity, i.e., the -Y direction of the body in the Earth-oriented attitude) are on the same side of the orbital plane; negative otherwise. As shown in the figure, Beta ranges from 17° to 34°.

[0124] The satellite's two-dimensional solar orientation can fully obtain energy in the sunlit area. However, considering that frequent attitude adjustments need to be executed through the reaction flywheel of the on-board actuator, it will increase energy consumption and reduce the life of the reaction flywheel of the rotating component. Therefore, the satellite is designed to always face the sun in two dimensions when there is no mission requirement. This brings about the problem that the camera in the shadow area needs more energy to heat the cold air.

[0125] Based on the -26° tilt of the sailboard, the normal direction of the satellite sailboard can be as close to the sun as possible on the left side of the flight direction (within the -Y half plane) to ensure energy acquisition. This solution will bring difficulties to the structural design, and the unfolded sailboard area is about 1 / 2 of the body, which does not contribute much to energy. Figure 13 shown.

[0126] Taking all the above into consideration, the satellite system designs adopted include the following three schemes. The advantages and disadvantages of these three schemes are compared, as shown in Table 1.

[0127] Table 1 Comparison of the advantages and disadvantages of three satellite system design schemes

[0128]

[0129]

[0130] Schematic diagram of attitude roll -26°, as shown Figure 14 shown.

[0131] Depend on Figure 14 It can be seen that the camera can still receive the heating effect of the earth's radiation and infrared radiation, and can effectively obtain energy. The angle curve between the sun vector and the sailboard normal in the sunlit area of 60 minutes in one track is given and the effective duration of direct sunlight is calculated to be 35 minutes. In 60 minutes, there are 55.6 minutes when the angle between the sun vector and the sailboard normal is less than 90°. The simulation results are as follows Figure 15 shown.

[0132] Compared with the situation that the shadow area and the sunlit area are all two-dimensionally facing the sun during the non-mission period, the above attitude is 2 The satellite of the sailboard will reduce the illumination and charging time of the sailboard by 25 minutes, and the corresponding energy is about 200×0.8×25 / 60=67Wh. The heating power consumption requirement of the visible light remote sensing camera is about 40W, and the heating power consumption requirement of the laser is about 35W. Calculated based on a 50% duty cycle, the heating demand for one orbit is about 71Wh.

[0133] Figure 16 and Figure 17 The load current and bus voltage in the no-mission rolling -26° attitude mode for approximately 10 hours on February 4th.

[0134] Figure 18 and Figure 19 The load current and bus voltage in the two-dimensional solar mode without mission for about 10 hours on March 18.

[0135] Based on the statistical data in the above figure, in the absence of a mission, the power consumption at a -26° 2D roll relative to the Sun increases by (3.22-2.3)*28=25.76W. For a single orbit, the increased energy demand is approximately 25.76*95 / 60=41Wh. The solar cells at a -26° 2D roll relative to the Sun generate an additional 67Wh of power, a difference of only 37%. Furthermore, due to the general trend of cost reduction in commercial satellites, the majority of micro-nanosatellite components are industrial-grade, significantly increasing the probability of being knocked over by a single particle in orbit. A -26° roll relative to the Sun attitude allows for full utilization of Earthshine and infrared radiation to generate heating, allowing for reduced camera usage (maintaining a temperature around 5°C) in the event of a heater driver chip failure that prevents the heater from turning on, thus preventing permanent device failure.

[0136] 3. Posture Scrolling Design Module

[0137] Based on the above analysis, this module is used to design an attitude rolling scheme applicable to all satellite orbits. The specific implementation steps are as follows:

[0138] Step 1: Calculate the orbital solar angle Beta based on the satellite's on-orbit flight conditions. Figure 20 shown.

[0139] Step 2: Set the initial default parameters of the software according to the recommended roll angle value.

[0140] For attitude pointing during the non-mission phase, the software has default guidance parameters. Based on the orbital solar angle Beta calculated in the previous step, the most appropriate value for the orbit entry time is selected as the software default value. This value is set in a form that can be configured through ground commands and stored in the non-volatile memory (NVM) of the onboard computer to prevent the loss of configured parameters due to computer software reset or hardware restart.

[0141] According to a third aspect of the present invention, a computer terminal is provided, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, the terminal can be used to execute the above-mentioned method of the present invention, or to execute the above-mentioned system of the present invention.

[0142] Optionally, the memory is used to store programs; the memory may include volatile memory (English: volatile memory), such as random-access memory (English: random-access memory, abbreviated: RAM), such as static random-access memory (English: static random-access memory, abbreviated: SRAM), double data rate synchronous dynamic random access memory (English: Double Data Rate Synchronous Dynamic Random Access Memory, abbreviated: DDR SDRAM), etc.; the memory may also include non-volatile memory (English: non-volatile memory), such as flash memory (English: flash memory). The memory is used to store computer programs (such as applications, functional modules, etc. that implement the above-mentioned methods), computer instructions, etc., and the above-mentioned computer programs, computer instructions, etc. can be partitioned and stored in one or more memories. In addition, the above-mentioned computer programs, computer instructions, data, etc. can be called by the processor.

[0143] The processor is configured to execute the computer program stored in the memory to implement the various steps of the method or various modules of the system involved in the above embodiments. For details, please refer to the relevant descriptions in the above method and system embodiments.

[0144] The processor and memory can be independent structures or integrated structures. When the processor and memory are independent structures, the memory and processor can be coupled via a bus.

[0145] According to a fourth aspect of the present invention, a computer-readable storage medium is provided, on which a computer program is stored. When the computer program is executed by a processor, it can be used to perform the above-mentioned method of the present invention, or to run the above-mentioned system of the present invention.

[0146] Computer-readable media include computer storage media and communication media, wherein communication media include any medium that facilitates the transfer of computer programs from one location to another. The storage medium can be any available medium that can be accessed by a general-purpose or special-purpose computer. An exemplary storage medium is coupled to a processor so that the processor can read information from the storage medium and write information to the storage medium. Of course, the storage medium can also be an integral part of the processor. The processor and storage medium can be located in an ASIC. Alternatively, the ASIC can be located in a user device. Of course, the processor and storage medium can also exist as discrete components in a communication device.

[0147] The low-orbit satellite design method and system provided by the above-mentioned embodiment of the present invention take into account the symmetrical installation of sail panels, which reduces the difficulty of balancing the center of mass of the entire satellite, makes the inertia of the entire satellite easy to control, and is friendly to the attitude control of the satellite. The fault situation in which the heater hardware on the satellite is unable to turn on due to being knocked over by a single particle is taken into account. With the help of the heat generated by the earth's reflection and infrared radiation, irreversible damage to the ground-mounted cameras can be avoided. The difficulty of sail panel design can be reduced, while the cost of installing SADA to obtain energy and the complexity of the entire satellite design can be reduced. The ground transmission of data transmission can be taken into account: the data transmission antenna is axially pointed to the center of the earth, and communicates with the ground data transmission station through phased array electronic scanning.

[0148] Matters not mentioned in the above embodiments of the present invention are well known in the art.

[0149] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art may make various variations or modifications within the scope of the claims, which do not affect the essence of the present invention.

Claims

1. A method for designing a low-orbit satellite with a normal attitude maintenance function, characterized in that: include: Set the basic parameters of low-orbit satellites; Based on the basic parameters, the required low-orbit satellite model is obtained through satellite coordinate system design, satellite layout design and satellite system design; Based on the low-orbit satellite model, the roll angle of the satellite orbit attitude is designed to achieve normal attitude maintenance of the low-orbit satellite.

2. The method for designing a low-orbit satellite with a normal attitude maintenance function according to claim 1, characterized in that: The basic parameters of the low-orbit satellite are set as follows: The total weight of the satellite is set to not exceed 100kg; Set a sun-synchronous circular orbit with an altitude of 510 km and a descending node time of 10:00-10:30 AM local time. Set the visible light remote sensing camera startup temperature: 5℃~35℃, operating temperature: 20℃~25℃; Set the laser load starting temperature: 0℃~35℃, and the operating temperature: 20℃~25℃.

3. The method for designing a low-orbit satellite with a normal attitude maintenance function according to claim 1, characterized in that: The satellite coordinate system design includes: The satellite structure coordinate system is defined as Os—XsYsZs, where: Os is the center point of the satellite-rocket separation plane, that is, the origin; The Xs axis is perpendicular to the separation plane of the satellite and rocket, and points to the propulsion module along the short side of the satellite, corresponding to the flight direction under the premise of stable attitude towards the earth on orbit; The Ys axis is parallel to the separation plane of the rocket and satellite, and points to the direction of the laser payload along the long side of the release mechanism, corresponding to the negative normal of the orbital plane under the premise of stable attitude towards the earth on orbit; The Zs axis is orthogonal to Xs and Zs in the right-hand coordinate system, and corresponds to the direction of the center of the earth under the premise of stable attitude on orbit; By translating the satellite structure coordinate system to the satellite center of mass, we obtain the satellite body coordinate system Ob—XbYbZb, abbreviated as O—XYZ.

4. The method for designing a low-orbit satellite with a normal attitude maintenance function according to claim 1, characterized in that: The satellite layout design includes: Integrated design of remote sensing camera payload; The entire satellite is designed as a long L-shaped configuration, and the satellite is designed as a flat long strip configuration. The satellite solar array adopts a single-wing secondary deployment structure. After the satellite enters orbit, the middle plate of the solar array is unfolded outward 180° and then locked after being unlocked by pyrotechnics. At the same time, the two outer plates are also unfolded outward 180° and then locked.

5. The method for designing a low-orbit satellite with a normal attitude maintenance function according to claim 1, characterized in that: The satellite system design includes any of the following: - Unfold the sailboard and tilt it at 26°; - Windsurfing formal wear, no two-dimensional sun during mission; -Sailboarding in full gear, attitude roll -26° during non-mission period.

6. The method for designing a low-orbit satellite with a normal attitude maintenance function according to claim 1, characterized in that: The step of designing a roll angle of the satellite orbit attitude based on the low-orbit satellite model includes: Calculate the orbital solar angle Beta based on the satellite's on-orbit flight conditions; Initial default parameter settings are made based on the roll angle recommendations.

7. The method for designing a low-orbit satellite with a normal attitude maintenance function according to claim 6, characterized in that: The calculation of the orbital solar angle Beta according to the satellite's on-orbit flight conditions includes: M1, determine the initial orbit insertion parameters based on the designed low-orbit satellite model; M2: Determine whether it is a sun-synchronous orbit. If so, calculate the Beta angle for a set number of years and jump to M4. If not, obtain the short period T of Beta angle change through theoretical calculation and simulation, and continue to execute the following steps. M3, calculates the Beta angle within a short period T and the Beta angle within a set number of years; M4, gives the recommended values of the rolling angle at different times within the set period of time; The initial default parameter setting according to the recommended roll angle value includes: For the attitude pointing in the non-mission phase, based on the orbital solar angle Beta, the most appropriate value for the orbit entry time is selected from the system default guidance parameters as the initial default parameter, and is set to be stored in a form that can be configured through ground commands.

8. A low-orbit satellite design system with a normal attitude maintenance function, characterized in that: include: Constraint design module, which is used to set the basic parameters of low-orbit satellites; A satellite design module, which obtains the required low-orbit satellite model through satellite coordinate system design, satellite layout design, and satellite system design based on the basic parameters; The attitude roll design module designs the roll angle of the satellite orbit attitude based on the low-orbit satellite model, so as to achieve normal attitude maintenance of the low-orbit satellite.

9. A computer terminal comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein: When executing the computer program, the processor can be used to perform the method according to any one of claims 1 to 7, or run the system according to claim 8.

10. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, it can be used to perform the method according to any one of claims 1 to 7, or to run the system according to claim 8.

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