Thermal analysis and control design are prerequisite essential to design the satellite. In the space environment, it makes satellite survive from extreme hot and cold conditions. In recent years CubeSat mission is developed for many kinds of purpose. Triplet Ionospheric Observatory (TRIO)-CubeSat for Ion, Neutral, Electron, MAgnetic fields (CINEMA) is required to weigh less than 3 kg and operate on minimal 3 W power. In this paper we describe the thermal analysis and control design for TRIO-CINEMA mission. For this thermal analysis, we made a thermal model of the CubeSat with finite element method and NX6.0 TMG software is used to simulate this analysis model. Based on this result, passive thermal control method has been applied to thermal design of CINEMA. In order to get the better conduction between solar panel and chassis, we choose aluminum 6061-T6 for the material property of standoff. We can increase the average temperature of top and bottom solar panels from -70°C to -40°C and decrease the average temperature of the magnetometer from +93°C to -4°C using black paint on the surface of the chassis, inside of top & bottom solar panels, and magnetometer.
This paper presents an estimation scheme of mass property and momentum actuator alignment of the rigid-body spacecraft with considering external torque. In many previous researches, it is assumed that external torque acts on a spacecraft is negligible. This assumption makes it feasible to build an estimator based on constant angular momentum vector of rigid-body spacecraft in inertial frame. However, the influence of external torque increases dramatically in low orbit Earth observation mission because of gravity-gradient torque. This paper develops a novel estimator that formulates gravity-gradient torque to get exact system dynamics equation. The performance of an estimator is verified by high-fidelity simulator.
Nomenclature= Direction cosine matrix from inertial frame (N) to body frame at time (k) = Transpose of matrix A 1 = m-by-m identity matrix = Inertial parameters = [ ] [ ] = Body-fixed inertial matrix ω = Body rate of spacecraft in body frame , = Inertial of th reaction-wheel n = Number of reaction-wheels = Rotational rate of th reaction-wheel = Alignment vector of th reaction-wheel in body frame , , , = Extended Kalman filter system dynamics relevant components (system dynamics, state, control input, and process noise respectively) , ℋ, = Extended Kalman filter measurement equation relevant components (system measurement: means measurement for variable x, measurement equation, and measurement noise respectively) h = Total angular momentum of spacecraft in inertial frame T = Rotational kinetic energy P = Actuator misalignment parameters = [ 11 12 21 22 ⋯ 2] ∆ = Error value of variable x ̂ = Best knowledge of variable x = Position vector from spacecraft to center of the Earth in body frame = Standard deviation of random variable , ℎ = Additional linearization errors of rotational kinetic energy and angular momentum respectively = Gravity constant (3.986012 × 10 6 3 2 ⁄ ) = Gravity-gradient torque at time time
In this paper, a real-time on-board orbit determination algorithm using the high precise orbit propagator is suggested and its performance is analyzed. Orbit determination algorithm is designed with the Extended Kalman Filter. And it utilizes the orbit calculated from the Pseudo-range as observed data. The performance of the on-board orbit determination method implemented in the GPS-12 receiver is demonstrated using the GNSS simulator. Orbit determination performance using high precise orbit propagator was analyzed in comparison to the orbit determination result using orbit propagator. The
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