Sun orientation calculation method based on energy constraint

Abstract

The invention discloses a sun orientation calculation method based on energy constraint, and the method comprises the steps: enabling a flight control computer to output charging power P to the flight control computer through a power controller as an input, calculating an included angle theta between a solar incident vector direction and a solar cell array, and outputting the included angle theta; the method comprises the following steps: decomposing an included angle theta between a solar incident vector direction and a solar cell array normal into information of three azimuth angles, determining measured solar vector orientations OA, OB and OC, constructing an index function J for eliminating system errors in combination with a theoretical solar vector orientation S, and performing quadratic form orthogonal transformation operation on the index function J to obtain an index function Jnew of a solar vector estimated value; solving an error estimation minimum value of the Jnew to obtain a calculated measured sun vector azimuth S, and calculating a maximum probability estimation value Ps of the sun vector direction by using the measured sun vector azimuth S; and calculating attitude adjusting azimuth angles alpha and beta of the solar panel to the sun according to Ps, inputting the attitude adjusting azimuth angles alpha and beta into a flight control computer, driving the solar panel to complete solar panel to the sun, and realizing energy supply of the spacecraft. According to the invention, the problem of contradiction between spacecraft flight reliability improvement and low cost is solved.

Description

Technical Field

[0001] The invention relates to a solar orientation calculation method based on energy constraints, belonging to the technical field of electrical overall design of aerospace transportation systems. Background Art

[0002] The solar panels are oriented to the sun to ensure that the spacecraft can obtain energy. To obtain solar vector information, it is usually necessary to equip the spacecraft with sensitive devices such as sun sensors, gyroscopes, and star sensors. Once these sensitive devices fail, the spacecraft will fail in its on-orbit mission. In recent years, the traditional solution is to increase the number of sensors to improve reliability, but there are constraints such as cost and limited structural layout.

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

[0004] The technical problem solved by the present invention is: to overcome the deficiencies of the prior art, to provide a solar orientation calculation method based on energy constraints, to improve the reliability of the spacecraft in the event of a primary failure of the attitude control sensor, to achieve solar orientation in a non-sensitive state, and to solve the energy supply problem of the spacecraft.

[0005] The technical solution of the present invention is: a solar orientation calculation method based on energy constraints, comprising:

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

[0007] The flight control computer uses the charging power P as input, calculates the angle θ between the direction of the solar incident vector and the solar array, and outputs it;

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

[0009] The measured solar vector azimuths OA, OB, and OC are determined by three azimuth information θ1, θ2, and θ3;

[0010] Combining the measured solar vector orientations OA, OB, OC and the theoretical solar vector orientation S, an index function J for eliminating system errors is constructed;

[0011] The index function J that eliminates the system error is transformed by quadratic orthogonal transformation to obtain the index function J of the solar vector estimate calculated by the least squares method. new ;

[0012] By using the indicator function J of the estimated value of the sun vector new Find the minimum error estimate and get the calculated measured sun vector position Use it to calculate the maximum probability estimate of the sun vector direction P s ;

[0013] According to the maximum probability estimate P of the sun vector direction s , calculate the solar sail panel's attitude adjustment angles α and β;

[0014] Input α and β into the flight control computer to drive the sailboard to face the sun and realize the energy supply of the spacecraft.

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

[0016] P(0)=∫V·p·n·dA

[0017] θ=cos -1 (P(θ) / P(0))

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

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

[0020] Preferably, the measured solar vector azimuths OA, OB, OC are determined by three azimuth information θ1, θ2, 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 system errors is:

[0027] J=|OA×S| 2 +|OB×S| 2 +|OC×S| 2

[0028] Among them, S represents the theoretical solar vector orientation, OA, OB, and OC are the measured solar vector orientations, that is, the measured values ​​of S.

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

[0030] J new =(DS)'Λ(DS)+λ(S'S-1)

[0031] Among them, D represents the orthogonal transformation matrix for solving the quadratic form, S represents the sun vector direction, λ represents the Lagrange multiplier term, Λ represents the matrix with the square of the eigenvalues ​​of the quadratic form matrix as diagonal elements, and 1 represents the unit matrix.

[0032] Preferably, the index function J of the estimated value of the sun vector is new Find the minimum error estimate and get the calculated measured sun vector position The method is:

[0033]

[0034] S represents the theoretical sun vector position, and λ represents the Lagrange multiplier term.

[0035] Preferably, the maximum probability estimate P of the sun vector direction is calculated s The method is:

[0036]

[0037] represents the calculated measured solar vector position, S represents the theoretical solar vector position, σ v express , D represents the orthogonal transformation matrix for solving the quadratic form, and λ represents the Lagrange multiplier term.

[0038] Preferably, according to the maximum probability estimate value P of the sun vector direction s , the method for calculating the solar sail panel's attitude adjustment angles α and β is:

[0039]

[0040]

[0041] Among them, y srepresents the component of the sun vector position on the y-axis of the spacecraft system, z s It represents the component of the sun vector position on the z-axis of the spacecraft system.

[0042] Compared with the prior art, the present invention has the following advantages:

[0043] (1) The present invention improves the energy design during the process of the spacecraft's orientation to the sun based on the output power of the solar array collected by the power controller, thereby solving the contradiction between the improvement of the flight reliability of the spacecraft and the low cost;

[0044] (2) The present invention introduces the least square method to reduce the error rate of solar vector measurement, thereby solving the problem of large solar vector calculation error caused by large solar array output power measurement error;

[0045] (3) The present invention provides a semi-physical testing method based on an air flotation simulator when the above algorithm is applied in the engineering stage, which has the ability to perform semi-physical testing on the prototype and final product of the spacecraft using this method. BRIEF DESCRIPTION OF THE DRAWINGS

[0046] Figure 1 The flowchart of the solar orientation calculation method based on the energy constraint of the solar array of the present invention is as follows;

[0047] Figure 2 It is a flow chart of the air flotation table test of the present invention. DETAILED DESCRIPTION

[0048] Inspired by the solar orientation technology of Taimin, this paper realizes a solar orientation calculation method based on the energy constraint of solar array, which can realize the solar orientation of solar array based on the output power of solar array in the non-Tamin state. A method of error optimization is proposed to reduce the problem of large solar vector positioning measurement error caused by power output error. Finally, in order to verify the correctness of this method, an experimental test method based on air flotation simulator is proposed.

[0049] The solar orientation calculation method based on solar array energy constraints consists of two parts: a power-based orientation determination algorithm and a probability estimation algorithm to reduce measurement errors. Figure 1 , the specific technical solution process is as follows:

[0050] S01: The spacecraft's power controller collects the solar array output current and bus voltage, and inputs the charging power P to the flight control computer.

[0051] S02: The flight control computer uses the power P as input, calculates the angle θ between the direction of the solar incident vector and the normal line of the solar cell array, and outputs it.

[0052] P(0)=∫V·p·n·dA

[0053] θ=cos -1 (P(θ) / P(0))

[0054] Wherein, V represents bus voltage, p represents standard solar intensity, n represents the direction of the sun's incident normal, A represents the cell microelement surface, P represents the input power of the flight control computer, P(θ) represents the cell output power when the angle between the sun's incident vector direction and the solar cell array normal is θ, and P(0) represents the cell output power when the angle between the sun's incident vector direction and the solar cell array normal is 0;

[0055] S03: By collecting the output power of the solar cell array in different directions, the angle θ between the direction of the solar incident vector and the normal of the solar cell array is decomposed into θ1, θ2, and θ3, where θ1 represents the angle between the solar incident vector and the x-axis of the spacecraft system, θ2 represents the angle between the solar incident vector and the y-axis of the spacecraft system, and θ3 represents the angle between the solar incident vector and the z-axis of the spacecraft system.

[0056] S04: Determine the measured solar vector azimuths OA, OB, and OC through the three azimuth information θ1, θ2, and θ3, using the following method:

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

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

[0059] OC=(OC x ,OC y ,OC z )

[0060]

[0061] S05: Combine the theoretical sun vector position S to construct the index function J for eliminating the system error. The index function for eliminating the system error is:

[0062] J=|OA×S| 2 +|OB×S| 2 +|OC×S| 2

[0063] Among them, OA, OB, OC represent the measured values ​​of the theoretical solar vector azimuth S in the plane perpendicular to the three body axes;

[0064] S06: Perform a quadratic orthogonal transformation operation on the index function J that eliminates the system error, and obtain the index function J that can be used to calculate the estimated value of the solar vector using the restricted least squares method new .

[0065] J new =(DS)'Λ(DS)+λ(S'S-1)

[0066] Where D represents the orthogonal transformation matrix for solving the quadratic form, S represents the theoretical sun vector position, λ represents the Lagrange multiplier term, Λ represents a matrix with the squares of the eigenvalues ​​of the quadratic form matrix as diagonal elements, and 1 represents the identity matrix.

[0067] S07: By calculating the indicator function J of the estimated value of the sun vector new The error estimate minimum is obtained to obtain the calculated measured sun vector azimuth Use it to calculate the maximum probability estimate of the sun vector direction P s .

[0068] in:

[0069] By calculating the indicator function J new The error estimate minimum is obtained The method is:

[0070]

[0071] Calculate the maximum probability estimate P of the sun vector direction s The method is:

[0072]

[0073] Among them, P s represents the maximum probability estimate of the sun vector direction, represents the calculated measured solar vector position, S represents the theoretical solar vector position, σ v express , D represents the orthogonal transformation matrix for solving the quadratic form, and λ represents the Lagrange multiplier term;

[0074] S08: Based on the maximum probability estimate of the sun vector direction P s , solve the solar sail panel's attitude adjustment angles α and β.

[0075]

[0076]

[0077] Among them, y s represents the component of the sun vector position on the y-axis of the spacecraft system, z sRepresents the component of the sun vector position on the z-axis of the spacecraft system;

[0078] S09: Input the solar panel's attitude adjustment angles α and β to the flight control computer, drive the panel to face the sun, and realize the energy supply of the spacecraft.

[0079] In order to verify the correctness of this method, an experimental test method based on an air flotation simulator is proposed. The experimental process is shown in Figure 2 , the specific test process is as follows:

[0080] S01: First, each system is powered on, and the current information of the air floating platform is read at the ground control terminal to complete the initial information collection.

[0081] S02: Adjust the air flotation platform, which includes a marble platform and a satellite, wherein the satellite includes a solar array, an integrated electronic module and an air flotation simulator; in addition, a simulated light source is arranged in the test environment so that the solar array can obtain incident power information of the simulated light source in three directions.

[0082] S03: Read the attitude adjustment information of the air-floating platform at the ground control end, complete ground information collection, calculate the vector direction of the light source relative to the satellite, and perform satellite attitude conversion matrix compensation calculation.

[0083] S04: Get the final P output from the satellite s , obtain the simulated light source vector position identification result, and verify the correctness of the simulated light source vector position calculation result.

[0084] The contents not described in detail in the specification of the present invention belong to the prior art known to the professional and technical personnel in this field.

Claims

1. A solar orientation calculation method based on energy constraints, characterized in that include: The spacecraft's power controller collects the solar array output current and bus voltage, and outputs the charging power P to the flight control computer; The flight control computer uses the charging power P as input, calculates the angle θ between the direction of the solar incident vector and the solar array, and outputs it; Decompose the angle θ between the direction of the solar incident vector and the normal line of the solar cell array into three azimuth angle information θ1, θ2, θ3; The measured solar vector azimuths OA, OB, and OC are determined by three azimuth information θ1, θ2, and θ3; Combining the measured solar vector orientations OA, OB, OC and the theoretical solar vector orientation S, an index function J for eliminating system errors is constructed; The index function J that eliminates the system error is transformed by quadratic orthogonal transformation to obtain the index function J of the solar vector estimate calculated by the least squares method. new ; By using the indicator function J of the estimated value of the sun vector new Find the minimum error estimate and get the calculated measured sun vector position Use it to calculate the maximum probability estimate of the sun vector direction P s ; According to the maximum probability estimate P of the sun vector direction s , calculate the solar sail panel's attitude adjustment angles α and β; Input α and β into the flight control computer to drive the sailboard to face the sun and realize the energy supply of the spacecraft.

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

3. The solar orientation calculation method based on energy constraints according to claim 1 is characterized in that: θ1 represents the angle between the solar incident vector and the x-axis of the spacecraft system, θ2 represents the angle between the solar incident vector and the y-axis of the spacecraft system, and θ3 represents the angle between the solar incident vector and the z-axis of the spacecraft system.

4. The solar orientation calculation method based on energy constraints according to claim 3 is characterized by: The measured solar vector azimuths OA, OB, and OC are determined by the three azimuth information 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 ) 5. The solar orientation calculation method based on energy constraints according to claim 1 is characterized in that: The index function J for eliminating system errors is: J=|OA×S| 2 +|OB×S| 2 +|OC×S| 2 Among them, S represents the theoretical solar vector orientation, OA, OB, and OC are the measured solar vector orientations, that is, the measured values ​​of S.

6. The solar orientation calculation method based on energy constraints according to claim 1 is characterized by: Index function J of the estimated value of the sun vector new for: J new =(DS)'Λ(DS)+λ(S'S-1) Among them, D represents the orthogonal transformation matrix for solving the quadratic form, S represents the sun vector direction, λ represents the Lagrange multiplier term, Λ represents the matrix with the square of the eigenvalues ​​of the quadratic form matrix as diagonal elements, and 1 represents the unit matrix.

7. The solar orientation calculation method based on energy constraints according to claim 1 is characterized by: By using the indicator function J of the estimated value of the sun vector new Find the minimum error estimate and get the calculated measured sun vector position The method is: S represents the theoretical sun vector position, and λ represents the Lagrange multiplier term.

8. The solar orientation calculation method based on energy constraints according to claim 1 is characterized by: Calculate the maximum probability estimate P of the sun vector direction s The method is: represents the calculated measured solar vector position, S represents the theoretical solar vector position, σ v express , D represents the orthogonal transformation matrix for solving the quadratic form, and λ represents the Lagrange multiplier term.

9. The solar orientation calculation method based on energy constraints according to claim 1 is characterized by: According to the maximum probability estimate P of the sun vector direction s , the method for calculating the solar sail panel's attitude adjustment angles α and β is: Among them, y s represents the component of the sun vector position on the y-axis of the spacecraft system, z s It represents the component of the sun vector position on the z-axis of the spacecraft system.

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