Microsatellite Alignment to Position of Maximum Solar Panels Illumination Without a Sun Sensor

Abstract

To generate the maximum amount of electricity, solar panels of a spacecraft must be placed at the position of maximum illumination when the surface normal of the panels is close to the direction of the sun. In micro-and nanosatellites, solar panels are typically fixed to the body, without the possibility of autonomous rotation. Therefore, this issue is solved by an Attitude Determination and Control System that rotates the spacecraft by using its own actuators. This control mode is called “solar orientation mode” and to determine the required position of the satellite signals from solar sensors are usually used. This paper discusses a situation in which a solar sensor signal is missing owing to its failure or the shading of its field of view by satellite structural elements, or no solar sensor (using an Earth sensor or star sensor for attitude control). An algorithm was proposed for calculating control signals in the solar orientation mode using two direction cosine matrices: between the inertial and orbital reference frames, as well as between the orbital and body reference frames. The first matrix is calculated from the Keplerian elements, or from the spacecraft radius- and velocity- vectors obtained from the global navigation satellite system. The second matrix was determined using a strapdown inertial attitude system based on angular rate gyro sensors. The algorithm of such a system uses the equations of satellite kinematics in quaternions and considers the orbital angular rate. Numerical simulations confirmed the good characteristics.

Keywords

• Attitude determination and control system
• microsatellite
• solar panel orientation
• solar sensor failure
• strapdown inertial attitude system