CN-121997535-A - Closed-loop verification method for spacecraft energy system

CN121997535ACN 121997535 ACN121997535 ACN 121997535ACN-121997535-A

Abstract

The invention provides a closed loop verification method of a spacecraft energy system, which comprises the steps of obtaining spacecraft position coordinates, spacecraft attitude coordinates, solar wing rotation angles and star time, establishing a spacecraft position correction matrix according to the spacecraft position coordinates and the star time, rotating a spacecraft position coordinate system through matrix transformation to enable a Y axis to point to the sun direction, obtaining a solar vector direction vector, and a position and unit direction vector of the spacecraft under a new coordinate system, identifying whether the spacecraft is in a shadow area or a sun area according to the spacecraft position coordinates after position correction, correcting the spacecraft attitude when the spacecraft is in the sun area, determining a solar wing method phase and a solar vector direction included angle according to spacecraft attitude information and solar wing rotation angles, obtaining curve parameters of a solar matrix simulator according to the solar wing method phase and the solar vector direction included angle, and setting an output state and output power of the solar matrix simulator according to the obtained curve parameters for spacecraft energy system verification.

Inventors

GENG WENHAO
TIAN HANG
DUAN YUXIAO
CHANG JUNMING
HAN HAONAN
YU FENG
XIE ZHIYONG
LIU HE
ZHAO FENG
ZHAO ZHENG
NI LINNA
YING PENG
LIU YIFAN

Assignees

北京空间飞行器总体设计部

Dates

Publication Date: 20260508
Application Date: 20251211

Claims (10)

1. A method for closed loop verification of a spacecraft energy system, comprising: acquiring a spacecraft position coordinate, a spacecraft attitude coordinate, a solar wing rotation angle and a star time; According to the position coordinates of the spacecraft and the star time, a spacecraft position correction matrix is established, a spacecraft position coordinate system is rotated through matrix transformation, a-Y axis points to the sun direction, and a sun vector direction vector, the position of the spacecraft under a new coordinate system and a unit direction vector are obtained; identifying whether the spacecraft is in a shadow area or a sun area according to the spacecraft position coordinates after the position correction; correcting the attitude of the spacecraft when the spacecraft is in the sunshine area, and determining the normal phase of the solar wing and the included angle of the solar vector direction according to the attitude information of the spacecraft and the rotation angle of the solar wing; And acquiring curve parameters of the solar array simulator according to the included angle between the solar wing method and the solar vector direction, and setting the output state and the output power of the solar array simulator according to the acquired curve parameters for the verification of the spacecraft energy system.
2. The method for verifying closed loop of spacecraft energy system according to claim 1, wherein the step of obtaining spacecraft position coordinates, spacecraft attitude coordinates, solar wing rotation angle and star time comprises obtaining semi-major axis a, eccentricity e, orbit inclination i, ascending intersection point right ascent and descent The method comprises the steps of obtaining a near-heart point radial angle omega, a true near-heart point angle T, a star time T me , a yaw angle phi 0 , a roll angle phi 0 , a pitch angle theta 0 and a solar wing rotation angle rho 0 , converting the star time into quantized digital information, and obtaining a time parameter lambda, wherein lambda represents an included angle between a connecting line of the earth and the sun and a connecting line of the earth and the sun in spring.
3. The method for verifying closed loop of spacecraft energy system according to claim 2, wherein in the step of establishing a spacecraft position correction matrix, the correction matrix is a four-time rotation matrix multiplication, and the orbit inclination angle, the elevation intersection point offset, the earth polar axis inclination and the earth revolution offset are corrected respectively, and the correction matrix is: in the formula, Is a correction matrix.
4. The method for closed loop verification of a spacecraft energy system according to claim 2, wherein in the step of obtaining a solar vector direction vector, a position of the spacecraft in a new coordinate system, and a unit direction vector, the solar vector direction vector is: 。
5. a method of closed loop verification of a spacecraft energy system according to claim 3, wherein in said step of obtaining a solar vector direction vector, a position of the spacecraft in a new coordinate system and a unit direction vector, the correction matrix is corrected by a left-hand multiplication Obtaining the position of the spacecraft under the new coordinate system, wherein the unit direction vector is : Wherein x n is the correction matrix Y n is the correction matrix Z n is the correction matrix U is the position of the spacecraft in the orbit plane, u=ω+t.
6. The method for closed-loop verification of a spacecraft energy system according to claim 5, wherein the step of identifying whether the spacecraft is in a shadow area or a sun area according to the position coordinates of the spacecraft after the position correction is implemented by: projecting spacecraft orbit under new coordinate system to XOZ plane to construct projection function ; Wherein r e is the average length of the earth radius, and r a is the orbit height of the spacecraft; projection function There are 4 intersections with the projection of the earth on the XOZ plane, the intersections satisfy the following formula: Taking two intersection points when y n is less than 0, wherein the two intersection points are the boundary points of a shadow area and a sun area of the spacecraft, when When the spacecraft is in a shadow area, the output of the solar matrix simulator is forbidden, and when the spacecraft is in a sun area, the solar matrix simulator supplies power.
7. The method for closed loop verification of a spacecraft energy system of claim 2, wherein said step of correcting the attitude of the spacecraft is performed by: the direction vectors defining three coordinate axes of the spacecraft body coordinate system are respectively as follows when the yaw angle psi=0, the roll angle phi=0 and the pitch angle theta=0 , , The direction vectors of three coordinate axes of the spacecraft body coordinate system when the yaw angle psi=psi 0 , the roll angle phi=phi 0 and the pitch angle theta=theta 0 are respectively as follows , , According to 、、 Carrying out attitude adjustment on the spacecraft in the sequence of (1); First, the spacecraft body coordinate system is wound The rotation psi 0 is anticlockwise 、 : Then, wind around Anticlockwise rotation phi 0 to obtain 、 : Finally, wind around Anticlockwise rotation of theta 0 、 : Depending on the invariant nature of the rotation axis, the following results were obtained: The direction vectors of three coordinate axes of the spacecraft body coordinate system when the yaw angle psi=psi 0 , the roll angle phi=phi 0 and the pitch angle theta=theta 0 of the spacecraft are obtained through the transformation , , 。
8. The method for closed-loop verification of a spacecraft energy system according to claim 7, wherein the step of determining the angle between the normal phase of the solar wing and the direction of the solar vector according to the attitude information of the spacecraft and the rotation angle of the solar wing is implemented by: defining the solar wing, when the yaw angle psi=psi 0 , the roll angle phi=phi 0 , the pitch angle theta=theta 0 and the solar wing rotation angle rho=rho 0 , the solar wing normal direction vector is ; Solar incidence angle of solar wing 。
9. The method for closed-loop verification of a spacecraft energy system according to claim 8, wherein the step of obtaining the curve parameters of the solar array simulator according to the included angle between the solar wing method phase and the solar vector direction is implemented by the following steps: V oc 、V max 、I sc 、I max ,V oc corresponding to the solar incident angle α 0 is an open-circuit voltage, V max is a maximum power point voltage, I sc is a short-circuit current, and I max is a maximum power point current.
10. Computer program product, characterized in that it comprises a computer program which, when executed, performs the spacecraft energy system closed-loop verification method of one of claims 1 to 9.

Description

Closed-loop verification method for spacecraft energy system Technical Field The invention belongs to the technical field of spacecraft testing, and particularly relates to a closed-loop verification method for a spacecraft energy system. Background In the comprehensive testing process of the spacecraft, the solar array simulator is main equipment for testing the spacecraft. The solar array simulator supplies power to the spacecraft through the output curve of the solar wing, and the traditional test modes comprise an instruction control method and a circulation list method. The command control method is to write a pre-stored parameter curve into the solar array simulator through software to realize the output adjustment of the solar array simulator. The circulation list method is to write a group of prestored parameter curves into the solar array simulator in a circulation mode according to a pre-designed time sequence to realize circulation adjustment of the solar array simulator. As the power supply and distribution system of the spacecraft becomes more and more complex, the on-orbit maneuvering capability of the spacecraft becomes stronger and the requirement for testing the authenticity becomes higher and higher. The solar array simulator is required to adjust power in real time according to the on-orbit illumination condition of the spacecraft so as to simulate the real power generation state of the solar wing, and the rationality of the design of the power supply and distribution system of the spacecraft and the correctness of energy balance analysis are verified. Therefore, an effective test means is needed to make the spacecraft energy system test close to the on-orbit real state. The power generation power of the solar wing of the spacecraft during on-orbit is related to the solar illumination intensity, and the change of the power generation power of the solar wing is continuous under the influence of the orbit, the position and the gesture of the spacecraft. The command control method and the circulation list method can not simulate the characteristic of continuous change of solar wing power generation power caused by the change of on-orbit illumination conditions of the spacecraft, and only a prestored discrete curve can be called. Aiming at the problems, the invention provides a closed-loop verification method of a spacecraft energy system based on a lightweight digital twin technology, which establishes a lightweight digital twin model capable of rapidly predicting the illumination incidence angle of a solar wing, takes real-time telemetry data of the orbit and the attitude of the spacecraft as input data of the model, calculates the illumination condition of the solar wing in real time, dynamically adjusts the output curve of a solar array simulator, simulates the time-varying property of continuous change of a solar-wing power generation curve, verifies whether the energy balance analysis of the spacecraft meets design requirements, and realizes the high-dynamic simulated flight of the spacecraft energy system. Disclosure of Invention In order to overcome the defects in the prior art, the inventor performs intensive research and provides a closed-loop verification method for a spacecraft energy system, the illumination condition of a solar wing is calculated in real time by using real-time telemetry data of the orbit and the attitude of the spacecraft, the output curve of a solar array simulator is dynamically regulated, the time variability of continuous change of a solar wing power generation curve is simulated, whether the energy balance analysis of the spacecraft meets the design requirement is verified, and the high-dynamic simulated flight of the spacecraft energy system is realized. The technical scheme provided by the invention is as follows: in a first aspect, a closed loop verification method for a spacecraft energy system includes: acquiring a spacecraft position coordinate, a spacecraft attitude coordinate, a solar wing rotation angle and a star time; According to the position coordinates of the spacecraft and the star time, a spacecraft position correction matrix is established, a spacecraft position coordinate system is rotated through matrix transformation, a-Y axis points to the sun direction, and a sun vector direction vector, the position of the spacecraft under a new coordinate system and a unit direction vector are obtained; identifying whether the spacecraft is in a shadow area or a sun area according to the spacecraft position coordinates after the position correction; correcting the attitude of the spacecraft when the spacecraft is in the sunshine area, and determining the normal phase of the solar wing and the included angle of the solar vector direction according to the attitude information of the spacecraft and the rotation angle of the solar wing; And acquiring curve parameters of the solar array simulator according to the included angle between the solar wing met