Spacecraft Attitude Determination and Control

Journal of Physics: Conference Series, 2013

This work discusses an attitude control study for the ASTER mission, the first Brazilian mission to the deep space. The study is part of a larger scenario that is the development of optimal trajectories to navigate in the 2001 SN263 asteroid system, together with the generation of orbit and attitude controllers for autonomous operation. The spacecraft attitude is defined from the orientation of the body reference system to the Local Vertical Local Horizontal (LVLH) of a circular orbit around the Alpha asteroid. The rotational equations of motion involve the dynamic equations, where the three angular speeds are generated from a set of three reaction wheels and the gravitational torque. The rotational kinematics is represented in the Euler angles format. The controller is developed via the linear quadratic regulator approach with output feedback. It involves the generation of a stability augmentation (SAS) loop and a tracking outer loop, with a compensator of desired structure. It was chosen the feedback of the p, q and r angular speeds in the SAS, one for each reaction wheel. In the outer loop, it was chosen a proportional integral compensator. The parameters are tuned using a numerical minimization that represents a linear quadratic cost, with weightings in the tracking error and controls. Simulations are performed with the nonlinear model. For small angle manoeuvres, the linear results with reaction wheels or thrusters are reasonable, but, for larger manoeuvres, nonlinear control techniques shall be applied, for example, the sliding mode control.

AIAA/AAS Astrodynamics Specialist Conference and Exhibit, 2004

This paper presents the attitude determination method for the Bifocal Relay Mirror Spacecraft Simulator. The simulator simulates three-axis motion of a spacecraft and has an optical system emulating a bifocal space telescope. The simulator consists of three control moment gyroscopes, rate gyros, two-axis analog sun sensor, and two inclinometers. The five-foot diameter platform is supported on a spherical air bearing to offer a low-torque environment. This paper demonstrates two attitude determination methods employing the measurements from a two-axis analog IR sensor, two inclinometers, and a triaxial gyroscope. The first method implements the conventional Kalman filter algorithm. The second method uses a nonlinear observer derived from the Lyapunov's direct method. Analytical and experimental results are presented to validate the proposed algorithm.

2011

Spin-stabilized spacecraft generally rely on sun and three-axis magnetic eld sensor measurements for attitude determination. This study experimentally determines the total accuracy of attitude determination solutions using modest quality sensors. This was accomplished by having a test spacecraft collect data during spinning motions. The data was then post-processed to nd the attitude estimates, which were then compared to the experimentally measured attitude. This same approach will be used to test the accuracy of the attitude determination system of the DICE spacecraft to be built by SDL/USU.

2019

Based on the three-dimensional dynamics of a rigid body and Newton's laws, the simplified dynamics of a spacecraft is studied and described through the systematical representation, mathematical modeling and also by a block diagram representation, to finally simulates the spacecraft dynamics in the Matlab programming environment called Simulink. It is paramount to be able to identify and recognize the attitude (often represented with the Euler angles) and position variables like the degrees of freedom (DOF) of the system and also the linear behavior. All this to conclude up about the non-linear behavior presented by the accelerations, velocities, positions and Euler angles (attitude) when those mentioned are plotted against time. In addition to this, the linearized system is found in order to facilitate the control analysis and stability analysis, at using linear analysis tools of Simulink and concepts like controllability and observability, reaching the point of determining under the previous concepts to proceed with the control design phase. Lastly, an uncertainty and sensitivity analysis is realized, by means the Monte-Carlo and the Linear regression method (in Simulink too), to find the torque like critical model input, since it has the greatest effect on the response variables in the system; and thus finally, to implement the Linear Quadratic Regulator (LQR) controller, at using the lqr Matlab function.

The spacial missions will have a high automati-zation level, making the pointing precision growing up. The control system trust will be very important. It is possible to see that the tests will have more hardware into de software mesh, so the process starts from a control system complete simulation and, slowly, the on board computer and sensors will be added, using the real system parts, and simulating at a computer only what is necessary.

2012

This paper describes a path toward the development of theory for using a low noise high frame rate camera as a star tracker for spacecraft attitude estimation. The benefit of using a low noise high frame rate camera is that s ar data can be sampled at a faster rate while allowing one to measure very di m stars, increasing the number of stars available for attitude estimation. The d evelopment of a noise model is discussed and an algorithm to process raw data is sho wn. An attitude estimation method is discussed and simulated data is shown. A simulated star tracker for attitude estimation is shown and attitude estim ation results are shown.

Advances in Estimation, Navigation, and Spacecraft Control, 2015

Attitude determination, along with attitude control, is critical to functioning of every space mission. In this paper, we investigate and compare, through simulation, the application of two autonomous sequential attitude estimation algorithms, adopted from the literature, for attitude determination using attitude sensors (sun sensor and horizon sensors) and rate-integrating gyros. The two algorithms include a direction cosine matrix (DCM) based steady-state Kalman Filter and the classic quaternion-based Extended Kalman Filter. To make the analysis realistic, as well as to improve the design of the attitude determination algorithms, detailed sensor measurement models are developed. Modifications in the attitude determination algorithms, through estimation of additional states, to account for sensor biases and misalignments have been presented. A modular six degree-of-freedom closed-loop simulation, developed in house, is used to observe and compare the performances of the attitude determination algorithms.

GOES-8 and Beyond, 1996

This paper presents a summary of the basic simulation parameters and results of a study for the Geostationary Operational Environmental Satellite (GOES). The study involves the simulation of minor modifications to the current spacecraft, so that the relative performance of these modifications can be analyzed. The first modification studied requires the placement of a baseline inertial reference unit, such as the Dry Rotor Inertia Reference Unit (DRIRU-II) or the Space Inertial Reference Unit (SIRU), onto the spacecraft. The second modification involves using the imager/sounder assembly for real-time on-board attitude determination information. The third modification studied is the addition of star trackers to provide precise attitude knowledge. Simulation results are presented for each modification.

2003

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Planetary penetrator missions offer excellent opportunities for examining the planetary structure in great detail. The success of such missions largely depends on navigating the penetrator to impact the target with the right orientation so that the scientific instruments onboard are safe. This in turn requires a sophisticated autonomous attitude determination and control subsystem (ADCS) onboard the penetrator. The aim of this paper is to propose a suitable mission sequence for such hard landers along with identification of appropriate sensors and actuators, and to examine the attitude determination and control strategies, desired for the right impact. A detailed investigation on using horizon detectors and Sun sensors for attitude determination along with the use of gas jets for attitude control will be discussed. The penetrator is a spin-stabilized platform, and hence a forced precession of the spin axis using the Sun sensor as the reference for firing the thrust pulses results in the penetrator following a rhumb-line attitude trajectory relative to the Sun direction. Results of simulations for the penetrator attitude control targeting the Lunar and Martian surfaces are presented.