Attitude and Orbit Control The attitude of a spacecraft is its orientation space with respect to some reference system, i.e. how the spacecraft body axes are oriented relative to an inertial or rotating coordinate system. The motion of a spacecraft is described by four aspects: position, velocity, attitude and attitude motion. Position and velocity are the subject of orbital mechanics. Attitude and attitude motion, however, are the subjects of attitude dynamics, i.e. the motion of the spacecraft about its center of mass. Since environmental torques (such as gravity gradient torques, solar wind torques and atmospheric torques) are a function spacecraft position and velocity, spacecraft orbit and attitude are coupled. That is, the spacecraft position and velocity determine the environmental torques which affect the attitude motion, where as in low attitude Earth orbits, the attitude of the spacecraft affects the atmospheric drag on the spacecraft and therefore its orbit. Attitude control is the process of orienting the spacecraft in a desired direction. This includes attitude stabilization (maintaining the attitude in a desired state) and attitude maneuver control (changing the attitude from one orientation, old state to another orientation, new state). This process involves the use of attitude control hardware (actuator such as reaction wheels or jets), on-board or remote computers to generate the commands and relevant software. The concept selected is given in Figure 10.3.8. Any spacecraft requires a type of attitude control and determination. For engineering purpose, attitude control is needed for functions such as pointing the antennas in a desired direction, pointing solar panels in a desired direction, pointing the control jets in the desired direction to accomplish efficient maneuvers. Assumptions The satellite is in a sun-synchronous orbit, the inclination selected is 100 degrees for an altitude of 1000 km. The antenna is a phased array type, the solar array does not need fine pointing so the body will be piloted with an accuracy of 0.5 degree which is classical for a LEO satellite. Looking at the assigned mission the satellite will be three axis stabilized. Sensors • Four gyroscope blocks of two axes (two redundant) • Two Earth sensors of Infrared scanning type (one) on +Z side • Two sun sensors on +Y side (one redundant) • Two sun acquisition sensors. One on +Y side, one on -Y side Actuators • Three reaction wheels, one in each satellite axis • Two magneto torquers • A propulsion system • For simplicity it is chosen to have a monopropelant system using Mono Methyl Hydrazine; the tanks will be pressurized on the upper part by nitrogen (ystem used on Spot). We estimate the consumption for 5 years to 110 kg ; the satellite will be equipped of two tanks of 75 kg MMH each. • Twelve 15N thrusters on the sides +,- Y and +,- X • Pressure transducers, electrical and pyromecanical valves, fill and drain valves, filters Platform Electrical Architecture The Electric Control System receives a small part of the electric power from the large solar array that generates the power to be transmitted to the Earth and distributes the appropriate power to each onboard equipment, thermal control system, attitude control system, orbit control system, TT&C (Telemetry, Command, and Control system) and so on, of the spacecraft. The total power for all the systems would be much smaller than the power transmitted to the Earth as the mission. If the orbit would be selected to have eclipses, batteries to provide power for all systems should be required. The electric power needed would then be collected by the electric control system directly from the solar array, or from the battery connected to the solar array. The electrical architecture is shown in Figure 10.3.9.
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