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Research on Satellite Autonomous Navigation and Attitude Determination Based on Spaceborne Sensors

Author: WangPeng
Tutor: ZhangYingChun
School: Harbin Institute of Technology
Course: Control Science and Engineering
Keywords: Autonomous navigation Attitude Determination Continuous - discrete extended Kalman filter Information Fusion Observability and considerable measure
CLC: V448.22
Type: PhD thesis
Year: 2008
Downloads: 1115
Quote: 9
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Abstract


With the successful launch of China's Shenzhou series of spacecraft, lunar exploration program and the implementation of the plan of deep space exploration, spacecraft autonomy viability of concern, one of the key technologies of the new century. Spacecraft autonomous navigation and attitude determination technology is an important aspect of satellite autonomous spacecraft control today technology trends, and in reducing the burden of the ground-based monitoring system, reducing the running costs of the spacecraft, to improve the viability of the spacecraft, and the expansion of satellite The application potential significance. Therefore, use of the the spaceborne sensor using information fusion method improves independent satellite navigation and attitude determination accuracy and reliability are the key elements of the spacecraft design and hot issues. National Security Key Basic Research Program (973) This paper combines the new concept of \Specific work is as follows: only single determining line the spacecraft navigation parameters or attitude information compared magnetometer measurements of the geomagnetic field with traditional methods, proposed a gyroscope and magnetometer combined geomagnetic field measurement information to determine low Earth orbit Satellite navigation method parameters and attitude information. First detailed analysis of a mathematical model of the geomagnetic field and the magnetic observatories model; then derive the track six elements of orbital dynamics model and attitude quaternion kinematic model, the establishment of the equation of state of the combined system; Second magnetometer measured values ??with the international Geomagnetic Reference Field (IGRF) model estimates for poor, the establishment of a combination of systematic observation equation, the expression that contains both spacecraft navigation parameters and attitude information obtained by both differential analysis of the observed difference: only one observation, and finally design a state-of-the-art continuous - discrete extended Kalman filter, numerical simulation, and the simulation results and system performance analysis and discussion of the combined system. Directly sensitive to the horizon can be divided into three autonomous celestial navigation method based on the fusion of information: (1) According to the star sensor and earth sensor combination sensitive besylate way for the use of the single celestial sensitive autonomous navigation shortcomings and deficiencies, proposed with indirect sensitive besylate. When indirectly sensitive horizon observed refraction Star, the introduction of satellite earth sensor measurement distance and direction vector as a new observation information. (2) the sun sensor and magnetometer combined, the two modes of operation when the satellite has been operating in different regions, according to the working principle of the sun sensor, can be divided into the sun light district and the Sun Shadow. When the satellite is in a sun shaded area, the introduction of the International Geomagnetic Reference Field (IGRF) model to the the magnetometer measured value and the cosine of the angle of the IGRF model as a new observation information. (3) for the above shortcomings and deficiencies of the proposed two kinds of autonomous navigation method based on information fusion, proposed the above satellite celestial sensor combination for autonomous navigation. Were designed according to the characteristics of each of the above three autonomous navigation system, adaptive EKF algorithm based on information fusion of these systems, numerical simulation, and the simulation results and system performance analysis and discussion. Strapdown inertial navigation system (SINS) is an autonomous navigation system, have autonomy, concealment, wideband and comprehensive information, etc., but the inertial measurement component errors accumulate over time, difficult to work continuously for a long time the shortcomings and lack of the same is very obvious. The star sensor is currently the most widely spaceborne attitude sensors, high pointing accuracy, no posture accumulated error characteristics. Proposed strapdown inertial navigation system with the star sensor combination, the most commonly used SINS system error mathematical model. Sensor observing the stars in the the Earth solid joint coordinate system or geographic coordinate system, and the star is in the geocentric inertial coordinate system, so geocentric inertial coordinate system is derived the SINS error mathematical model and its transmission characteristics. Select the equation of state of the 40-dimensional state variables integrated navigation system; the attitude quaternion SINS and star sensor output attitude quaternion error, observation equation of integrated navigation system using advanced continuous - discrete extended Kalman filter numerical simulation, and analysis of simulation results. The observability of the system state and considerable measure is an important indicator to verify the accuracy and speed of the design of the filter convergence. Typical for the autonomous navigation system, three analytical system observability and considerable measure: (1) According to the definition of system observability and analysis methods directly sensitive to the horizon celestial navigation system observability and impact its a major factor in the numerical analysis; (2) of the piecewise linear time-invariant system (PWCS) observability analysis; (3) of PWCS observability matrix singular value decomposition, analysis of the state of the system can be observed degrees. SINS / CNS integrated navigation system based on the singular value decomposition of the observation and considerable measure analysis and numerical simulation of integrated navigation system before and after the dimensionality reduction, and simulation results, and system performance are analyzed and discussed.

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CLC: > Aviation, aerospace > Aerospace ( Astronauts ) > Space instrument,spacecraft devices,spacecraft guidance and control > Guidance and Control > Spacecraft guidance and control > Attitude Control System
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