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The paper discusses the issues of designing a control moment gyroscope electric drive with strict requirements in terms of the accuracy of ensuring a given rotation rate of the gyro motor suspension. A brief description of the control moment gyroscope electric drive applied currently is presented and the issues of improving the electric drive characteristics are discussed. As a solution, an electric drive is proposed which operates in the mode of feedback loop using angle sensors located on the axes of the gyroscope suspension and the engine rotor. The paper describes the arrangement of the control moment gyroscopes on advanced spacecraft for Earth remote sensing and presents the analytic expressions needed to calculate the control moments that affect the spacecraft. The moments are in the projection to the coordinate system brought into coincidence with the spacecraft. The paper compares spacecraft angular velocity stabilization errors for the cases of using the conventional scheme of control moment gyroscope electric drive and the newly developed one. The presented results can be used for developing control moment gyroscope electric drives to be mounted on spacecraft of different purpose with strict requirements on ensuring operation at specified rotational velocities.

The algorithm of the turn of the spacecraft from an initial arbitrary angular position at an arbitrary angular rate to a solar-oriented attitude is investigated. Minimum essential equipment of the motion control system required for the purpose of ensuring maintenance of solar orientation is defined: a solar sensor, a single-axis angular-rate sensor, low-thrust liquid rocket engines. A solution of the problem of defining the spacecraft angular rate vector by the measurements of the deviation of the optical axis of the solar sensor from the sun vector and the single-axis angular-rate sensor is presented. The conditions under which control action on the rocket engines for the purpose of changing the value of the angular-rate vector for the Sun to get into the field of viewing of the solar sensor or for the spacecraft stabilization are defined. Mathematical modeling of the spacecraft attitude control system with the unknown initial state vector of motion is carried out. The results of mathematical modeling confirmed the efficiency of the proposed algorithm in terms of reducing propellant fuel consumption and high-speed performance. In comparison with the known methods of solving the problem of reducing angular speed (a lengthy process with the use of a magnetic system or a fast process with the use of a three-axis angular-rate sensor and rocket engines) the duration of the process of reducing angular speeds is the same as in normal operation, however, at the same time the problem of bringing spacecraft into the solar oriented attitude is solved.

The efficiency of a magnetic system of kinetic moment dumping in solving the problem of stabilizing Aist-2D spacecraft without using information on angular velocity is investigated. We defined the conditions under which the magnetic moment produced by electromagnets constitutes a useful control mechanical moment. Boundary conditions for preventing emergence of the parasitic moment are also defined. The paper presents the solution of the problem of forming the vector of the kinetic moment of the system “spacecraft + control moment gyro” according to the information on the vector of geomagnetic induction of the Earth. We defined the dependence of the time taken by the algorithm of angular speed reduction on the value and direction of the projections of the vector of spacecraft angular speeds at the time of algorithm actuation. The results obtained in the course of mathematical modeling and natural experiment on the spacecraft, the process of dumping angular speeds in providing the solution of the problem of attitude determination with the use of star sensors are compared.