A solar orientation calculation method based on energy constraints

Abstract

This invention discloses a solar orientation calculation method based on energy constraints. The flight control computer receives charging power P from the power controller as input, calculates and outputs the angle θ between the solar incident vector direction and the solar array. The angle θ between the solar incident vector direction and the solar array normal is decomposed into three azimuth angles to determine the measured solar vector azimuths OA, OB, and OC. Combined with the theoretical solar vector azimuth S, an index function J to eliminate system errors is constructed. A quadratic orthogonal transformation is performed on this function to obtain the index function J of the estimated solar vector value. new ; For J new Find the minimum error estimate to obtain the calculated solar vector azimuth. Use this to calculate the maximum probability estimate P of the solar vector direction. s According to P s The solar panel's attitude adjustment azimuth angles α and β relative to the sun are calculated and input into the flight control computer to drive the solar panel to align with the sun, thus providing energy to the spacecraft. This invention solves the contradiction between improving spacecraft flight reliability and reducing costs.

Description

Technical Field

[0001] This invention relates to a method for calculating solar orientation based on energy constraints, belonging to the field of electrical overall design technology for aerospace transportation systems. Background Technology

[0002] Solar panels' orientation to the sun is crucial for spacecraft to obtain energy. Acquiring solar vector information typically requires sensitive devices such as solar sensors, gyroscopes, and star sensors. Failure of these devices can lead to mission failure. In recent years, the traditional solution has been to increase the number of sensors to improve reliability, but this is constrained by cost and limitations in structural layout.

[0003] In summary, traditional solar orientation technology based on solar sensors presents a contradiction between low cost and high reliability. Summary of the Invention

[0004] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a solar orientation calculation method based on energy constraints, which improves the reliability of spacecraft under the condition of attitude control sensor failure, realizes solar orientation in the absence of attitude control sensor, and solves the energy supply problem of spacecraft.

[0005] The technical solution of this invention is: a method for calculating solar orientation based on energy constraints, comprising:

[0006] The spacecraft's power controller collects the output current and bus voltage of the solar array and outputs charging power P to the flight control computer.

[0007] The flight control computer takes the charging power P as input, calculates the angle θ between the solar incident vector direction and the solar cell array, and outputs the result.

[0008] The angle θ between the solar incident vector direction and the solar cell array normal is decomposed into three azimuth information θ1, θ2, and θ3;

[0009] The measured solar vector azimuths OA, OB, and OC are determined using three azimuth angles θ1, θ2, and θ3.

[0010] By combining the measured solar vector azimuths OA, OB, and OC with the theoretical solar vector azimuth S, an index function J for eliminating systematic errors is constructed.

[0011] A quadratic orthogonal transformation is performed on the index function J for eliminating systematic errors to obtain the index function J of the solar vector estimate calculated by the least squares method. new ;

[0012] Through the index function J of the solar vector estimate new By finding the minimum error estimate, the calculated measured solar vector azimuth can be obtained. Use it to calculate the maximum probability estimate P of the solar vector direction. s ;

[0013] Based on the maximum probability estimate P of the solar vector direction s Calculate the azimuth angles α and β of the solar panel relative to the sun;

[0014] Inputting α and β into the flight control computer drives the solar panel to align with the sun, thus supplying energy to the spacecraft.

[0015] Preferably, the method for calculating the angle θ between the solar incident vector direction and the solar cell array is as follows:

[0016] P(0)=∫V·p·n·dA

[0017] θ=cos -1 (P(θ) / P(0))

[0018] Where V represents the bus voltage, p represents the standard solar intensity, n represents the direction of the solar incident normal, A represents the micro-element of the solar cell, P represents the input power of the flight control computer, P(θ) represents the output power of the solar cell when the angle between the solar incident vector direction and the solar cell array normal is θ, and P(0) represents the output power of the solar cell when the angle between the solar incident vector direction and the solar cell array normal is 0.

[0019] Preferably, θ1 represents the angle between the solar incident vector and the x-axis of the spacecraft's main system, θ2 represents the angle between the solar incident vector and the y-axis of the spacecraft's main system, and θ3 represents the angle between the solar incident vector and the z-axis of the spacecraft's main system.

[0020] Preferably, the measured solar vector azimuths OA, OB, and OC are determined using three azimuth angles θ1, θ2, and θ3.

[0021] in:

[0022] OA = (OA x OA y OA z )

[0023] OB = (OB x ,OB y ,OB z )

[0024] OC = (OC x ,OC y ,OC z )

[0025]

[0026] Preferably, the index function J for eliminating systematic errors is:

[0027] J = |OA × S| 2 +|OB×S| 2 +|OC×S| 2

[0028] Where S represents the theoretical solar vector azimuth, and OA, OB, and OC are the measured solar vector azimuths, i.e., the measured values ​​of S.

[0029] Preferably, the index function J of the solar vector estimate new for:

[0030] J new =(DS)'Λ(DS)+λ(S'S-1)

[0031] Where D represents the orthogonal transformation matrix for solving the quadratic form, S represents the solar vector orientation, λ represents the Lagrange multiplier term, Λ represents the matrix with the squares of the eigenvalues ​​of the quadratic form matrix as diagonal elements, and 1 represents the identity matrix.

[0032] Preferably, the index function J of the solar vector estimate is used. new By finding the minimum error estimate, the calculated measured solar vector azimuth can be obtained. The method is as follows:

[0033]

[0034] S represents the theoretical solar vector orientation, and λ represents the Lagrange multiplier term.

[0035] Preferably, the maximum probability estimate P of the solar vector direction is calculated. s The method is as follows:

[0036]

[0037] The calculated measured solar vector azimuth is represented by S, and the theoretical solar vector azimuth is represented by σ. v express The statistical variance is given by , where D represents the orthogonal transformation matrix for solving the quadratic form, and λ represents the Lagrange multiplier term.

[0038] Preferably, based on the maximum probability estimate P of the solar vector direction s The method for calculating the azimuth angles α and β of the solar panel relative to the sun is as follows:

[0039]

[0040] Among them, y s The z-axis represents the component of the solar vector azimuth on the y-axis of the spacecraft's intrinsic system. sThis represents the component of the solar vector azimuth on the z-axis of the spacecraft's intrinsic system.

[0041] Compared with the prior art, the present invention has the following advantages:

[0042] (1) Based on the output power of the solar cell array collected by the power controller, the present invention improves the energy design during the process of spacecraft orientation to the sun, and solves the problem of the contradiction between improving the reliability of spacecraft flight and low cost.

[0043] (2) The present invention introduces the least squares method to reduce the error rate of solar vector measurement and solves the problem of large error in solar vector calculation caused by large error in solar cell array output power measurement;

[0044] (3) The present invention provides a semi-physical testing method based on an air-bearing simulator when the above algorithm is applied to the engineering stage, which has the ability to perform semi-physical testing on the prototype and formal products of spacecraft using the method. Attached Figure Description

[0045] Figure 1 This is a flowchart of the solar orientation calculation method based on solar cell array energy constraints of the present invention;

[0046] Figure 2 This is a flowchart of the air flotation platform test of the present invention. Detailed Implementation

[0047] Inspired by the solar array-based solar orientation technology, this invention realizes a solar orientation calculation method based on the energy constraint of a solar array. This method enables solar array orientation based on the output power of the solar array without a solar array. An error optimization method is proposed to reduce the large measurement error in solar vector positioning caused by power output errors. Finally, to verify the correctness of the method, an experimental testing method based on an air-bearing simulator is proposed.

[0048] The solar orientation calculation method based on solar array energy constraints consists of two parts: a power-based azimuth determination algorithm and a probabilistic estimation algorithm to reduce measurement errors. See flowchart below. Figure 1 The specific technical solution process is as follows:

[0049] S01: The spacecraft's power controller collects the output current and bus voltage of the solar array, and inputs the charging power P to the flight control computer.

[0050] S02: The flight control computer takes power P as input, calculates the angle θ between the solar incident vector direction and the normal line of the solar cell array, and outputs the result.

[0051] P(0)=∫V·p·n·dA

[0052] θ=cos-1 (P(θ) / P(0))

[0053] Where V represents the bus voltage, p represents the standard solar intensity, n represents the direction of the solar incident normal, A represents the micro-element of the solar cell, P represents the input power of the flight control computer, P(θ) represents the output power of the solar cell when the angle between the solar incident vector direction and the solar cell array normal is θ, and P(0) represents the output power of the solar cell when the angle between the solar incident vector direction and the solar cell array normal is 0.

[0054] S03: By collecting the output power of the solar cell array from different directions, the angle θ between the solar incident vector direction and the solar cell array normal is decomposed into θ1, θ2, and θ3. Among them, θ1 represents the angle between the solar incident vector and the x-axis of the spacecraft's own system, θ2 represents the angle between the solar incident vector and the y-axis of the spacecraft's own system, and θ3 represents the angle between the solar incident vector and the z-axis of the spacecraft's own system.

[0055] S04: The measured solar vector azimuths OA, OB, and OC are determined using three azimuth angles θ1, θ2, and θ3, using the following method:

[0056] OA = (OA x OA y OA z )

[0057] OB = (OB x ,OB y ,OB z )

[0058] OC = (OC x ,OC y ,OC z )

[0059]

[0060] S05: Combine the theoretical solar vector azimuth S to construct the index function J for eliminating systematic errors. The index function for eliminating systematic errors is:

[0061] J = |OA × S| 2 +|OB×S| 2 +|OC×S| 2

[0062] Among them, OA, OB, and OC represent the measured values ​​of the theoretical solar vector azimuth S in the plane perpendicular to the three body axes;

[0063] S06: Perform a quadratic orthogonal transformation on the index function J for eliminating systematic errors to obtain the index function J that can be used to calculate the solar vector estimate using the restricted least squares method. new .

[0064] J new =(DS)'Λ(DS)+λ(S'S-1)

[0065] Where D represents the orthogonal transformation matrix for solving the quadratic form, S represents the theoretical solar vector orientation, λ represents the Lagrange multiplier term, Λ represents a matrix with the squares of the eigenvalues ​​of the quadratic form matrix as its diagonal elements, and 1 represents the identity matrix.

[0066] S07: By calculating the index function J of the solar vector estimate new The minimum error estimate is used to obtain the calculated measured solar vector azimuth. Use it to calculate the maximum probability estimate P of the solar vector direction. s .

[0067] in:

[0068] By calculating the index function J new The minimum error estimate is obtained. The method is as follows:

[0069]

[0070] Calculate the maximum probability estimate P of the solar vector direction. s The method is as follows:

[0071]

[0072] Among them, P s This represents the maximum probability estimate of the solar vector direction. The calculated measured solar vector azimuth is represented by S, and the theoretical solar vector azimuth is represented by σ. v express The statistical variance is given by λ, where D represents the orthogonal transformation matrix for solving the quadratic form, and λ represents the Lagrange multiplier term.

[0073] S08: Based on the maximum probability estimate P of the solar vector direction s Calculate the azimuth angles α and β of the solar panel relative to the sun.

[0074]

[0075] Among them, y s The z-axis represents the component of the solar vector azimuth on the y-axis of the spacecraft's intrinsic system. s This represents the component of the solar vector azimuth along the z-axis of the spacecraft's intrinsic system;

[0076] S09: Input the azimuth angles α and β of the solar panel toward the sun into the flight control computer to drive the solar panel to complete the alignment with the sun and realize the energy supply of the spacecraft.

[0077] To verify the correctness of this method, an experimental test method based on an air flotation simulator is proposed. The experimental procedure is as follows: Figure 2 The specific testing process is as follows:

[0078] S01: First, all systems are powered on, and the current information of the air-floating platform is read at the ground control terminal to complete the initial information acquisition.

[0079] S02: Adjust the air-bearing platform, which includes a marble platform and satellites. The satellites include a solar cell array, an integrated electronic module, and an air-bearing simulator. Additionally, a simulated light source is arranged in the test environment to allow the solar cell array to acquire incident power information from the simulated light source in three directions.

[0080] S03: Read the attitude adjustment information of the air-bearing platform at the ground control terminal, complete the ground information acquisition, calculate the vector direction of the light source relative to the satellite, and perform satellite attitude transformation matrix compensation calculation.

[0081] S04: Obtain the final P output from the satellite. s The results of the simulated light source vector position identification are obtained, and the correctness of the simulated light source vector position calculation results is verified.

[0082] The contents not described in detail in this specification are existing technologies known to those skilled in the art.

Claims

1. A method for calculating solar orientation based on energy constraints, characterized in that... include: The spacecraft's power controller collects the output current and bus voltage of the solar array and outputs charging power P to the flight control computer. The flight control computer takes the charging power P as input, calculates the angle θ between the solar incident vector direction and the solar cell array, and outputs the result. The angle θ between the direction of the solar incident vector and the normal of the solar cell array is decomposed into three azimuth angles: θ1, θ2, and θ3. θ1 represents the angle between the solar incident vector and the x-axis of the spacecraft's own system, θ2 represents the angle between the solar incident vector and the y-axis of the spacecraft's own system, and θ3 represents the angle between the solar incident vector and the z-axis of the spacecraft's own system. The measured solar vector azimuths OA, OB, and OC are determined using three azimuth angles θ1, θ2, and θ3. By combining the measured solar vector azimuths OA, OB, and OC with the theoretical solar vector azimuth S, an index function J for eliminating systematic errors is constructed. A quadratic orthogonal transformation is performed on the index function J for eliminating systematic errors to obtain the index function J of the solar vector estimate calculated by the least squares method. new ; Through the index function J of the solar vector estimate new By finding the minimum error estimate, the calculated measured solar vector azimuth can be obtained. Use it to calculate the maximum probability estimate P of the solar vector direction. s ; Based on the maximum probability estimate P of the solar vector direction s Calculate the azimuth angles α and β of the solar panel relative to the sun; Inputting α and β into the flight control computer drives the solar panel to align with the sun, thus supplying energy to the spacecraft.

2. The solar orientation calculation method based on energy constraints according to claim 1, characterized in that: The method for calculating the angle θ between the solar incident vector direction and the solar cell array is as follows: P(0)=∫V·p·n·dA θ=cos -1 (P(θ) / P(0)) Where V represents the bus voltage, p represents the standard solar intensity, n represents the direction of the solar incident normal, A represents the micro-element of the solar cell, P represents the input power of the flight control computer, P(θ) represents the output power of the solar cell when the angle between the solar incident vector direction and the solar cell array normal is θ, and P(0) represents the output power of the solar cell when the angle between the solar incident vector direction and the solar cell array normal is 0.

3. The solar orientation calculation method based on energy constraints according to claim 1, characterized in that: The measured solar vector azimuths OA, OB, and OC are determined using three azimuth angles θ1, θ2, and θ3. in: OA=(OA x ,OA y ,OA z ) OB:(OB x ,WHETHER y ,WHETHER z ) OC=(OC x ,OC y ,OC z ) 4. The solar orientation calculation method based on energy constraints according to claim 1, characterized in that: The index function J for eliminating systematic errors is: J=|OA×S| 2 +|OB×S| 2 +|OC×S| 2 Where S represents the theoretical solar vector azimuth, and OA, OB, and OC are the measured solar vector azimuths, i.e., the measured values ​​of S.

5. The solar orientation calculation method based on energy constraints according to claim 1, characterized in that: The index function J of the solar vector estimate new for: J new =(DS)'Λ(DS)+λ(S'S-1) Where D represents the orthogonal transformation matrix for solving the quadratic form, S represents the solar vector orientation, λ represents the Lagrange multiplier term, Λ represents the matrix with the squares of the eigenvalues ​​of the quadratic form matrix as diagonal elements, and 1 represents the identity matrix.

6. The solar orientation calculation method based on energy constraints according to claim 1, characterized in that: Through the index function J of the solar vector estimate new By finding the minimum error estimate, the calculated measured solar vector azimuth can be obtained. The method is as follows: S represents the theoretical solar vector orientation, and λ represents the Lagrange multiplier term.

7. The solar orientation calculation method based on energy constraints according to claim 1, characterized in that: Calculate the maximum probability estimate P of the solar vector direction. s The method is as follows: σ represents the calculated measured solar vector azimuth. v express The statistical variance is given by , where D represents the orthogonal transformation matrix for solving the quadratic form, and λ represents the Lagrange multiplier term.

8. The solar orientation calculation method based on energy constraints according to claim 1, characterized in that: Based on the maximum probability estimate P of the solar vector direction s The method for calculating the azimuth angles α and β of the solar panel relative to the sun is as follows: Among them, y s The z-axis represents the component of the solar vector azimuth on the y-axis of the spacecraft's intrinsic system. s This represents the component of the solar vector orientation on the z-axis of the spacecraft's intrinsic system.

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