## 1 Introduction  
In recent years, with the development of information technology, microelectromechanical systems (MEMSs), new material technology, and high-end manufacturing, the global micro/nano satellite industry is growing rapidly and showing a \"blowout\" development momentum [1]. Micro/nano satellites generally refer to satellites weighing between 1 and 100 kg [2; 3; 4]. Due to their advantages, such as small size, light weight, low cost, flexible mobility, short development cycle, and ease to launch in batches, etc. [5; 6; 7; 8], micro/nano satellites have become one of the most important directions of current space technology development [9]. More recently, micro/nano satellites have begun to transform from technical experiments to business operations and have gradually replaced part of traditional large satellites [10]. Micro/nano satellites have a very broad development prospect and have been widely used in many fields, such as remote sensing imaging, electronic reconnaissance, satellite communication, satellite navigation, environmental detection, and technical experiment [2; 11; 12; 13]. For most micro/nano satellites, due to the limitations of size, cost, etc., the solar panel is fixed on the main body of the satellite and cannot be rotated. Therefore, it can only rely on the attitude determination and control system (ADCS) of the satellite to continuously adjust the satellite attitude and keep the solar panel pointing to the sun. Generally, in order to ensure a sufficient and reliable energy supply of the onboard system, the satellite needs to maintain a sun-oriented attitude for most of the time in orbit. As one of the most important satellite attitude control modes, the sun-oriented attitude is a specific satellite attitude where a certain vector fixed to the satellite is aligned with the solar vector, i.e.,the unit vector pointing from the center of mass of the satellite to the sun. In many cases, whether a stable and reliable sun-oriented attitude can be achieved directly determines the success or failure of the flight mission. Therefore, it is of great importance to study how to use the simplest sensors and actuators and the most reliable algorithm to achieve the sun-oriented attitude, especially for the case that star sensor, reaction wheel, and other devices are malfunctioned. Many studies related to sun-oriented attitude control have been carried out, and a lot of valuable research results have been obtained [14; 15; 16; 17; 18; 19; 20; 21]. As early as 1970, [PERSON] et al. [14] studied active magnetic attitude control of the spinning sun-oriented satellite. A magnetometer and a sun sensor were selected as attitude measurement devices, and a magnetic coil installed around the spin axis was used to generate the control torques. In addition, nutation dampers were used to dampen the wobble of the satellite's spin axis. The method is simple and has low power consumption and can be used to control the satellite to achieve a sun-oriented spin-stabilized attitude. Since then, many sun-oriented attitude control methods have been proposed. For example, in 2002, [PERSON] et al. [15] proposed a spin-stabilized attitude control algorithm for CUBESAT, which is a picosatellite developed by Stanford University and the California Polytechnic State University for space experiments. CUBESAT is equipped with two orthogonal magnetic torquing coils and a CO\({}_{2}\) high-pressure cartridge as attitude actuators and a 2-axis sun sensor as the attitude sensor. The proposed algorithm is used to control CUBESAT to maintain a spin-stabilized attitude, even during the shadow zone of the orbit. In 2007, [PERSON] [16] designed a spin-stabilized attitude control method for solar sail spacecraft and studied the method of generating spinning angular rate. In addition, the attitude control characteristics of the proposed method were simulated and analyzed. In 2012, aiming at the spin-stabilized satellite equipped with three magnetic coils or torque rods as attitude actuators, [PERSON] et al.