Closed-loop verification method for spacecraft energy system

Abstract

The invention provides a closed-loop verification method for an energy system of a spacecraft. The closed-loop verification method comprises the following steps: acquiring a spacecraft position coordinate, a spacecraft attitude coordinate, a solar wing rotation angle and star time; establishing a spacecraft position correction matrix according to the spacecraft position coordinates and the satellite time, rotating the spacecraft position coordinate system through matrix transformation, enabling the-Y axis to point to the sun direction, and obtaining a sun vector direction vector, the position of the spacecraft in a new coordinate system and a unit direction vector; according to the position coordinates of the spacecraft after position correction, whether the spacecraft is in a shadow area or a sunshine area is identified; when the spacecraft is in a sunny area, correcting the attitude of the spacecraft, and determining an included angle between a solar wing normal phase and a sun vector direction according to spacecraft attitude information and a solar wing rotation angle; according to the included angle between the solar wing normal phase and the solar vector direction, curve parameters of the solar array simulator are obtained, and according to the obtained curve parameters, the output state and the output power of the solar array simulator are set for spacecraft energy system verification.

Description

Technical Field

[0001] This invention belongs to the field of spacecraft testing technology, and specifically relates to a closed-loop verification method for spacecraft energy systems. Background Technology

[0002] In the integrated testing of spacecraft, the solar array simulator is a key piece of equipment. The solar array simulator powers the spacecraft through the output curves of the solar array. Traditional testing methods include two approaches: command control and cyclic list method. Command control involves writing pre-stored parameter curves into the solar array simulator via software to adjust its output. Cyclic list method involves repeatedly writing a set of pre-stored parameter curves into the solar array simulator according to a pre-designed timing sequence to achieve cyclic adjustment. As spacecraft power supply and distribution systems become increasingly complex, and their on-orbit maneuverability increases, the requirements for test realism become more stringent. The solar array simulator needs to adjust its power in real time according to the spacecraft's on-orbit illumination conditions to simulate the actual power generation state of the solar array, verifying the rationality of the spacecraft's power supply and distribution system design and the correctness of the energy balance analysis. Therefore, effective testing methods are required to ensure that the spacecraft's energy system testing closely approximates its actual on-orbit condition.

[0003] The power output of a spacecraft's solar array while in orbit is related to the intensity of sunlight. Influenced by the spacecraft's orbit, position, and attitude, the power output of the solar array varies continuously. Command control and cyclic list methods cannot simulate the continuous variation in solar array power output caused by changes in on-orbit lighting conditions; they can only utilize pre-stored discrete curves.

[0004] To address the aforementioned issues, this invention proposes a closed-loop verification method for spacecraft energy systems based on lightweight digital twin technology. A lightweight digital twin model is established that can rapidly predict the solar array's incident angle. Real-time telemetry data of the spacecraft's orbit and attitude are used as input data for the model. The solar array illumination is calculated in real time, and the output curve of the solar array simulator is dynamically adjusted to simulate the time-varying nature of the continuous changes in the solar array's power generation curve. This verifies whether the spacecraft's energy balance analysis meets design requirements and enables highly dynamic simulated flight of the spacecraft energy system. Summary of the Invention

[0005] To overcome the shortcomings of existing technologies, the inventors have conducted intensive research and provided a closed-loop verification method for spacecraft energy systems. This method uses real-time telemetry data of the spacecraft's orbit and attitude to calculate the solar array illumination in real time, dynamically adjusts the output curve of the solar array simulator, simulates the time-varying nature of the continuous changes in the solar array power generation curve, verifies whether the spacecraft's energy balance analysis meets the design requirements, and realizes high-dynamic simulated flight of the spacecraft energy system.

[0006] The technical solution provided by this invention is as follows: Firstly, a closed-loop verification method for a spacecraft energy system includes: Obtain the spacecraft's position coordinates, attitude coordinates, solar array rotation angle, and satellite time; Based on the spacecraft's position coordinates and satellite time, a spacecraft position correction matrix is ​​established. The spacecraft's position coordinate system is rotated through matrix transformation so that the -Y axis points towards the sun, and the sun vector direction vector, the spacecraft's position and unit direction vector in the new coordinate system are obtained. Based on the spacecraft's position coordinates after position correction, identify whether the spacecraft is in the shaded area or the sunlit area; When the spacecraft is in the sunlit area, the spacecraft attitude is corrected, and the angle between the normal phase of the solar array and the direction of the solar vector is determined based on the spacecraft attitude information and the rotation angle of the solar array. Based on the angle between the solar array normal and the solar vector direction, the curve parameters of the solar array simulator are obtained. Based on the obtained curve parameters, the output state and output power of the solar array simulator are set for spacecraft energy system verification.

[0007] Secondly, a closed-loop verification device for a spacecraft energy system includes: One or more processors; Storage device for storing one or more programs. When the one or more programs are executed by the one or more processors, the one or more processors implement the spacecraft energy system closed-loop verification method described in the first aspect.

[0008] Thirdly, a readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the spacecraft energy system closed-loop verification method described in the first aspect.

[0009] Fourthly, a computer program product comprising: a computer program (also referred to as code or instructions) that, when run, executes the spacecraft energy system closed-loop verification method described in the first aspect.

[0010] The closed-loop verification method for a spacecraft energy system provided by the present invention has the following beneficial effects: This invention provides a closed-loop verification method for a spacecraft energy system. It calculates the solar array illumination in real time using real-time telemetry data of the spacecraft's orbit and attitude, dynamically adjusts the output curve of the solar array simulator, simulates the time-varying nature of the continuous change of the solar array power generation curve, verifies whether the spacecraft's energy balance analysis meets the design requirements, and realizes high-dynamic simulated flight of the spacecraft energy system. Attached Figure Description

[0011] Figure 1This is a flowchart of the closed-loop verification method for the spacecraft energy system of the present invention; Figure 2 A schematic diagram showing the locations of the shaded and sunlit areas; Figure 3 The cosine of the solar incidence angle for two consecutive periods Examples of changes over time. Detailed Implementation

[0012] The features and advantages of the present invention will become clearer and more apparent from the following detailed description.

[0013] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments.

[0014] This invention provides a closed-loop verification method for a spacecraft energy system, such as... Figure 1 As shown, it includes the following steps: (1) Obtain the spacecraft's position coordinates, attitude coordinates, solar array rotation angle, and star time, and convert the star time into quantified digital information.

[0015] Energy generated by solar panels P cell Subject to the surface area of ​​solar cells A Solar panel power generation efficiency ŋ and the intensity of sunlight I The influence of these factors is as follows;

[0016] Sunlight intensity I and incident light angle α The intensity of sunlight is proportional to the cosine of 0. When the incident light is perpendicular to the solar array, the intensity of sunlight is at its maximum, expressed as: I max Sunlight intensity I for: I = I max * cos α 0 When the spacecraft is under normal illumination, I max Approximately equal to the solar constant. The solar constant is the flux density of incident light at a distance of one astronomical unit from the Sun. Rewrite the incident power of the solar array.

[0017]

[0018] Based on the above formula, we can conclude that, ideally, when the intensity of sunlight is constant, the spacecraft can be considered as a point mass, and the energy generated by the solar array... P The angle between the solar array's normal phase and the direction of the solar vector (i.e., the angle of incident light). α It is proportional to the cosine value of 0.

[0019] The angle between the solar array normal and the solar vector direction is related to the spacecraft's position, time, orbital altitude, spacecraft attitude, solar array mounting position, and solar array rotation angle. A mathematical model of the angle between the solar array normal and the solar vector direction is established using the aforementioned telemetry parameters, and the angle is calculated synchronously and in real-time based on spacecraft telemetry information.

[0020] This step involves acquiring the spacecraft's position coordinates, attitude coordinates, solar array rotation angle, and satellite time, including obtaining the semi-major axis. a eccentricity e Track inclination i Right ascension of ascending node Argument of the pericentric point oh True near point angle T Star Time T me Yaw angle ψ 0. Roll angle f 0. Pitch angle i 0. Solar panel rotation angle r 0.

[0021] (I) Calculate the distance between the spacecraft and the Earth's center r e + r a :

[0022] In the formula, r e The average length of the Earth's radius. r a This represents the orbital altitude of the spacecraft.

[0023] (II) Convert star time into quantified digital information to obtain time parameters. l , l It is represented as the angle between the line connecting the Earth and the Sun at this moment and the line connecting the Earth and the Sun at the vernal equinox; for example, June 12, 2024. l The calculation formula is as follows:

[0024] (III) Calculate the spacecraft's position in the orbital plane u : u = oh+ T (2) Establish the spacecraft position correction matrix, rotate the spacecraft position coordinate system through matrix transformation so that the -Y axis points to the sun, and obtain the sun vector direction vector, the spacecraft position and unit direction vector in the new coordinate system.

[0025] Rotating the inertial coordinate system via matrix transformation, so that the -Y axis of the new coordinate system points towards the sun, reduces the calculation of the X and Z components when performing vector multiplication operations (incident light angle calculation). Define the correction matrix. As shown in the following formula:

[0026] Correction Matrix This involves multiplying four rotation matrices to correct for orbital inclination, ascending node offset, Earth's polar tilt, and Earth's orbital offset. The rotations, interpreted from right to left according to the equations, represent clockwise rotations around the +X axis. i Rotate 90° clockwise around the +Z axis Rotate 23.5° counterclockwise around the +Y axis and rotate counterclockwise around the Z+ axis. l .

[0027] By left-multiplying the correction matrix The spacecraft's position in the new coordinate system is obtained, with its unit direction vector being... .

[0028]

[0029] In the formula, x n For the correction matrix The first element; y n For the correction matrix The second element; z n For the correction matrix The third element.

[0030] The direction vector of the solar vector is:

[0031] (3) Based on the spacecraft's position coordinates after position correction, identify whether the spacecraft is in the shadow area or the sun area.

[0032] Orient the spacecraft's orbit in the new coordinate system to a plane perpendicular to the Y-axis. XOZ (Plane) projection, constructing the projection function ;

[0033] Projection function With Earth XOZ The projection of the plane has four intersection points (intersection points 1, 2, 3, and 4), see... Figure 2 The intersection point satisfies the following formula:

[0034]

[0035] Pick y n The two intersection points (intersection points 1 and 2) when <0 are the boundary points between the shadow area and the sunlight area where the spacecraft is located. If the spacecraft is in the shadow area, the solar array simulator output is disabled; otherwise, if it is in the sunlit area, continue with the following steps.

[0036] (4) Correct the spacecraft attitude. Based on the spacecraft attitude information and the rotation angle of the solar array, determine the angle between the solar array normal phase and the solar vector direction.

[0037] Defined in yaw angle ψ =0 、 Roll angle f =0 、 Pitch angle i When = 0, the direction vectors of the three coordinate axes of the spacecraft's body coordinate system are respectively , , Yaw angle ψ = ψ 0 、 Roll angle f = f 0 、 Pitch angle i = i At time 0, the direction vectors of the three coordinate axes of the spacecraft's body coordinate system are respectively , , .according to , , The spacecraft's attitude is adjusted in sequence.

[0038] First, the spacecraft's body coordinate system revolves around Rotate counterclockwise ψ 0 got , :

[0039]

[0040] After that, around Rotate counterclockwise f 0 got , :

[0041]

[0042] Finally, around Rotate counterclockwise i 0 got , :

[0043]

[0044] Based on the property that the axis of rotation remains unchanged, the following result is obtained:

[0045]

[0046]

[0047] The spacecraft's yaw angle can be obtained through the above transformation. ψ = ψ 0 、 Roll angle f = f 0 、 Pitch angle i = i The direction vectors of the three coordinate axes of the spacecraft's body coordinate system at time 0 , , .

[0048] Solar wings around Rotation, when the solar array is horizontally zero, the normal vector of the solar array is... When the solar array rotates at an angle of... p=p At 0, the normal vector of the solar array can be considered as Around Rotation r 0. Define the solar array at the yaw angle. ψ = ψ 0 、 Roll angle f = f 0 、 Pitch angle i = i 0. Solar wing rotation angle p=p At 0, the normal phase vector of the solar array is .

[0049]

[0050] Therefore, the solar incidence angle of the solar array .

[0051] Figure 3 This diagram shows the cosine values ​​of the solar incidence angle for two consecutive cycles in a spacecraft's three-axis stabilized Earth-orbiting mode with the solar array horizontally zeroed out, in a specific example. Changes over time.

[0052] (5) Calculate the curve parameters of the solar array simulator based on the angle between the solar wing normal phase and the solar vector direction.

[0053] Searching for the solar incidence angle using piecewise linear interpolation α 0 corresponds to V oc 、V max 、I sc 、I max The four parameters found V oc 、V max 、I sc 、I max Write it into the solar array simulator. V oc Open circuit voltage, V max The voltage at the maximum power point. I sc This is the short-circuit current. I max This is the current at the maximum power point.

[0054] (6) Based on the calculated curve parameters, set the output state and output power of the solar array simulator.

[0055] Based on the above analysis, this invention establishes a lightweight digital model of the solar array's on-orbit illumination variation based on the spacecraft's orbit and attitude parameters. The output power of the solar array simulator is dynamically adjusted according to the model's output parameters to simulate the spacecraft's on-orbit energy variation process, making the spacecraft's energy system verification close to the actual on-orbit state. This effectively verifies the rationality of the spacecraft's power supply and distribution system design and the correctness of the energy balance analysis, realizing the testing of the spacecraft's energy system and high-dynamic flight simulation.

[0056] The present invention also provides a closed-loop verification device for a spacecraft energy system, comprising: One or more processors; Storage device for storing one or more programs. When the one or more programs are executed by the one or more processors, the one or more processors implement the spacecraft energy system closed-loop verification method described in the first aspect.

[0057] The present invention also provides a readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the spacecraft energy system closed-loop verification method described in the first aspect.

[0058] The readable storage media include, but are not limited to, USB flash drives, portable hard drives, read-only memory (ROM), random access memory (RAM), magnetic disks, optical disks, and other media capable of storing program code.

[0059] The present invention also provides a computer program product comprising: a computer program (also referred to as code or instructions), which, when run, executes the spacecraft energy system closed-loop verification method described in the first aspect.

[0060] In the above embodiments, implementation can be achieved, in whole or in part, through software, hardware, firmware, or any combination thereof. When implemented using software, it can be implemented, in whole or in part, as a computer program product. The computer program product includes one or more computer instructions. When the computer instructions are loaded and executed on a computer, all or part of the processes or functions described in the embodiments of this application are generated. The computer can be a general-purpose computer, a special-purpose computer, a computer network, or other programmable device. The computer instructions can be stored in a computer-readable storage medium or transmitted from one computer-readable storage medium to another. For example, the computer instructions can be transmitted from one website, computer, server, or data center to another via wired (e.g., coaxial cable, fiber optic, digital subscriber line (DSL)) or wireless (e.g., infrared, microwave, etc.) means.

[0061] Those skilled in the art will recognize that the units and algorithm steps of the various examples described in conjunction with the embodiments disclosed herein can be implemented in electronic hardware, or a combination of computer software and electronic hardware. Whether these functions are implemented in hardware or software depends on the specific application and design constraints of the technical solution. Those skilled in the art can use different methods to implement the described functions for each specific application, but such implementation should not be considered beyond the scope of this application.

[0062] Those skilled in the art will clearly understand that, for the sake of convenience and brevity, the specific working processes of the devices, apparatuses, and modules described above can be referred to the corresponding processes in the foregoing method embodiments, and will not be repeated here.

[0063] The present invention has been described in detail above with reference to specific embodiments and exemplary examples; however, these descriptions should not be construed as limiting the present invention. Those skilled in the art will understand that various equivalent substitutions, modifications, or improvements can be made to the technical solutions and embodiments of the present invention without departing from the spirit and scope of the invention, and all such modifications and improvements fall within the scope of the present invention. The scope of protection of the present invention is defined by the appended claims.

[0064] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A closed-loop verification method for a spacecraft energy system, characterized in that, include: Obtain the spacecraft's position coordinates, attitude coordinates, solar array rotation angle, and satellite time; Based on the spacecraft's position coordinates and satellite time, a spacecraft position correction matrix is ​​established. The spacecraft's position coordinate system is rotated through matrix transformation so that the -Y axis points towards the sun, and the sun vector direction vector, the spacecraft's position and unit direction vector in the new coordinate system are obtained. Based on the spacecraft's position coordinates after position correction, identify whether the spacecraft is in the shaded area or the sunlit area; When the spacecraft is in the sunlit area, the spacecraft attitude is corrected, and the angle between the normal phase of the solar array and the direction of the solar vector is determined based on the spacecraft attitude information and the rotation angle of the solar array. Based on the angle between the solar array normal and the solar vector direction, the curve parameters of the solar array simulator are obtained. Based on the obtained curve parameters, the output state and output power of the solar array simulator are set for spacecraft energy system verification.

2. The closed-loop verification method for a spacecraft energy system according to claim 1, characterized in that, The steps for obtaining the spacecraft's position coordinates, attitude coordinates, solar array rotation angle, and star time include: obtaining the semi-major axis. a eccentricity e Track inclination i Right ascension of ascending node Argument of the pericentric point ω True near point angle T Star Time T me Yaw angle ψ 0. Roll angle φ 0. Pitch angle θ 0. Solar panel rotation angle ρ 0; Convert star time into quantified digital information to obtain time parameters. λ , λ It is represented as the angle between the line connecting the Earth and the Sun at this moment and the line connecting the Earth and the Sun at the vernal equinox.

3. The closed-loop verification method for a spacecraft energy system according to claim 2, characterized in that, In the step of establishing the spacecraft position correction matrix, the correction matrix is ​​a product of four rotation matrices, which respectively corrects for orbital inclination, ascending node offset, Earth's polar axis tilt, and Earth's revolution offset. The correction matrix is ​​as follows: In the formula, This is the correction matrix.

4. The closed-loop verification method for a spacecraft energy system according to claim 2, characterized in that, In the steps of obtaining the solar vector direction vector, the spacecraft's position in the new coordinate system, and the unit direction vector, the solar vector direction vector is: .

5. The closed-loop verification method for a spacecraft energy system according to claim 3, characterized in that, In the step of obtaining the solar vector direction vector, the spacecraft's position in the new coordinate system, and the unit direction vector, the correction matrix is ​​multiplied by left. The spacecraft's position in the new coordinate system is obtained, with its unit direction vector being... : In the formula, x n For the correction matrix The first element; y n For the correction matrix The second element; z n For the correction matrix The third element; u The position of the spacecraft in the orbital plane. u = ω + T .

6. The closed-loop verification method for a spacecraft energy system according to claim 5, characterized in that, The step of identifying whether the spacecraft is in shadow or sunlight based on its position-corrected coordinates is implemented as follows: Orbit the spacecraft in the new coordinate system XOZ Plane projection, constructing the projection function ; In the formula, r e The average length of the Earth's radius. r a The orbital altitude of the spacecraft; Projection function With Earth XOZ The projection of the plane has 4 intersection points, which satisfy the following formula: Pick y n The two intersection points when <0, these two points are the boundary points between the shadow area and the sunlight area where the spacecraft is located. When the spacecraft is in the shadow zone, the solar array simulator output is disabled; otherwise, when it is in the sun, the solar array simulator is powered on.

7. The closed-loop verification method for a spacecraft energy system according to claim 2, characterized in that, The steps for correcting the spacecraft's attitude are implemented in the following manner: Defined in yaw angle ψ =0 、 roll angle φ =0 、 Pitch angle θ When = 0, the direction vectors of the three coordinate axes of the spacecraft's body coordinate system are respectively , , ; Yaw angle ψ = ψ 0 、 roll angle φ = φ 0 、 Pitch angle θ = θ At time 0, the direction vectors of the three coordinate axes of the spacecraft's body coordinate system are respectively , , ;according to , , The spacecraft's attitude is adjusted in sequence; First, the spacecraft's body coordinate system revolves around Rotate counterclockwise ψ 0 got , : After that, around Rotate counterclockwise φ 0 got , : Finally, around Rotate counterclockwise θ 0 got , : Based on the property that the axis of rotation remains unchanged, the following result is obtained: The spacecraft yaw angle is obtained through the above transformation. ψ = ψ 0 、 roll angle φ = φ 0 、 Pitch angle θ = θ The direction vectors of the three coordinate axes of the spacecraft's body coordinate system at time 0 , , .

8. The closed-loop verification method for a spacecraft energy system according to claim 7, characterized in that, The step of determining the angle between the solar array normal phase and the solar vector direction based on the spacecraft attitude information and the solar array rotation angle is implemented in the following manner: Define the solar array at the yaw angle ψ = ψ 0 、 roll angle φ = φ 0 、 Pitch angle θ = θ 0. Solar wing rotation angle ρ=ρ At 0, the normal phase vector of the solar array is ; Solar wing angle of incidence .

9. The closed-loop verification method for a spacecraft energy system according to claim 8, characterized in that, The step of obtaining the curve parameters of the solar array simulator based on the angle between the solar wing normal phase and the solar vector direction is implemented in the following manner: Searching for the solar incidence angle using piecewise linear interpolation α 0 corresponds to V oc 、V max 、I sc 、I max , V oc Open circuit voltage, V max The voltage at the maximum power point. I sc This is the short-circuit current. I max This is the current at the maximum power point.

10. A computer program product, characterized in that, The computer program product includes: a computer program that, when the computer program is run, executes the spacecraft energy system closed-loop verification method according to any one of claims 1 to 9.

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