• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
• /
• 2006.05a
• /
• pp.376-379
• /
• 2006

Reaction Wheel Assembly (RWA) is one of the major disturbance sources that have influence upon the Line of Sight (LOS) of payload. A micro-vibration induced by RWA is propagated through the satellite structure and decrease the LOS stability performance of payload. This effect shall be analyzed through the jitter analysis. If a requirement or specification of payload jitter level is found to be not satisfied according to the jitter analysis campaign, some modification or redesign should be done on the satellite structure or a couple of isolator should be attached on the RWA interface in order to reduce the transmitted vibration level of RWA. The purpose of ???RWA isolator test? is to roughly evaluate the performance of vibration suppression level with a passive RWA isolator made of rubber. For this test, actual RWA is used as a vibration source and a couple of cube-shaped rubber mount designed for satellite is used as a passive isolator. There may be several considerations in order to accommodate RWA isolator to spacecraft such as not only vibration reduction performance but also thermal conduction problem, mechanical size, RWA alignment problem, etc. But in this report the feasibility of RWA isolator is analyzed only in a vibration suppression point of view. As a result, high frequency vibration of RWA above 50Hz is perfectly attenuated with isolators, however, first harmonic components below 50Hz became larger due to the additional low frequency resonance modes of roll, pitch, yaw rigid body motion of RWA+bracket.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
• /
• 2013.04a
• /
• pp.533-539
• /
• 2013

Number of components and payload systems installed in satellites were found to be exposed to various disturbance sources such as the reaction wheel assembly, the control moment gyro, coolers, and others. A micro-level of vibration can introduce jitter problems into an optical payload system and cause significant degradation of the image quality. Moreover, the prediction of on-orbit vibration effects on the performance of optical payloads during the development process is always important. However, analyzing interactions between subsystems and predicting the vibration level of the payloads is extremely difficult. Therefore, this paper describes the analytical and experimental approach to microvibration effects on satellite optical payload performance with integrated jitter analysis framework, micro vibration emulator and satellite structure testbed.

• Transactions of the Korean Society for Noise and Vibration Engineering
• /
• v.24 no.1
• /
• pp.5-13
• /
• 2014

Number of components and payload systems installed in satellites were found to be exposed to various disturbance sources such as the reaction wheel assembly, the control moment gyro, coolers, and others. A micro-level of vibration can introduce jitter problems into an optical payload system and cause significant degradation of the image quality. Moreover, the prediction of on-orbit vibration effects on the performance of optical payloads during the development process is always important. However, analyzing interactions between subsystems and predicting the vibration level of the payloads is extremely difficult. Therefore, this paper describes the analytical and experimental approach to microvibration effects on satellite optical payload performance with integrated jitter analysis frame-work, microvibration emulator and satellite structure testbed.

• Journal of Astronomy and Space Sciences
• /
• v.36 no.3
• /
• pp.181-186
• /
• 2019

The attitude information of spacecraft can be obtained by the sensors attached to it using a star tracker, three-axis magnetometer, three-axis gyroscope, and a global positioning signal receiver. By using these sensors, the spacecraft can be maneuvered by actuators that generate torques. In particular, electromagnetic-torque bars can be used for attitude control and as a momentum-canceling instrument. The spacecraft momentum can be created by the current through the electrical circuits and coils. Thus, the current around the electromagnetic-torque bars is a critical factor for precisely controlling the spacecraft. In connection with these concerns, a solar-cell array can be considered to prevent generation of a magnetic dipole moment because the solar-cell array can introduce a large amount of current through the electrical wires. The maximum value of a magnetic dipole moment that cannot affect precise control is $0.25A{\cdot}m^2$, which takes into account the current that flows through the reaction-wheel assembly and the magnetic-torque current. In this study, we designed a 300-W solar cell array and presented an optimal wire-routing method to minimize the magnetic dipole moment for space applications. We verified our proposed method by simulation.