| Along with the Shenzhou series spacecraft, the lunar exploration project and the deep space exploration Project, the self-surviving performance of the spacecraft is regarded as one of the key technologies in 21st century. The autonomous navigaition and attitude determination technology is not only important aspect of spacecraft autonomy but also development trend of spacecraft control technology. It is important meaning for the modern satellite to reduce burden of ground-assisted system, decrease the cost of performance missions, greatly strengthen the probe's survival ability and extend the potential space application of the spacecraft. One of the new and widely researched problems in the subject of spacecraft design is to autonomously determinate orbit parameters by using measurement information of multi-spaceborne sensors and to improve system precision and reliability based on using information fusion method. With the support of the Nation Security Magnitude Foundation Research Foundation–Research on the new concept and new mechanism of minisize spacecraft, this dissertation deals with automous navigation and attitude determination based on multi-spaceborne sensors. The main contributions are as follows:The tradition methods using measurement of geomagnetic field based on magnetometer only determinate navigation or attitude parameter in singles. The method of determinating navigation and attitude parameter of LEO satellite simultaneitily using geomagnetic field information is presented based on integrated system of magnetometer and gyro. Firstly, the geomagnetic mathematic model and magnetometer measurement model are analyzed in detail. Secondly, the state equation of integrated system is established on deducing the model of the oribt dynamic model based on the orbit six elements and attitude kinematic model based on attitude quaternion. Thirdly, the observed equation of integrated system is established by using the difference of the measurement value of magnetometer and the estimated value of IGRF model. The observed equation includes spacecraft navigation and attitude parameter by analyzing the difference value synchronously. Compared with traditional method, the new method can estimate the spacecraft navigation and attitude parameter synchronously. Finally, the continuous-discrete extend kalman filter(CDEKF) algorithm is designed based on integrated system. The simulation result and system performance are analyzed and discussed after numerical simulation.Contraposing disadvantage and deficiency only using single spaceborne celestial sensor, the three methods of autonomous celestial navigation are presented based on information fusion. Firstly, the sensing earth modes include directly sensing horizon and indirect sensing horizon. When the refraction stars can't be observed, the distance and direction vector from satellite to earth is imported as the new observation information. Secondly, autonomous navigation method is presented based on sun sensor and magnetometer. According to the operational principle of sun sensor, the operational mode can be divided into two parts: sun shining area form and sun shadow area form. The estimated value of IGRF model is imported as the new observation information. The new observation value is got by the angle cosine based on the measurement value of magnetometer and the estimated value of IGRF model. Finally, Contraposing disadvantage and deficiency of the above-mentioned methods, a new autonomous navigation method is presented based on multi-celestial sensors. The Federal Extended Kalman Filter algorithm based on information fusion method is designed by analyzing the characteristic of aforementioned antonomous navigation methohds. The simulation result and system performance is analyzed and discussed after numerical simulation.The strapdown inertial navigation system(SINS) is a autonomous navigation system including the advantage of independence, concealment,wide frequency band and all-around information. As the drift errors of the inertial measurement component increase with time, the disadvantage and deficiency of SINS is obvious, which is unsuitable for long-time operating continuously. As a kind of used widely spaceborne sensor, the star sensor gains popularity for its high accuracy, without accumulative and intelligence. A new integrated navigation mode of SINS and star sensor is presented to correct the navigation parameter. The SINS error mathematic model is in the earth fixed coordinate or the geography coordinate usually, then the star sensor observes navigation stars in earth-centered inertia coordinate. The error mathematic model and diffusion characteristic of SINS are deduced in earth-centered inertia coordinate. The state equtation of integrated navigation system is established by 40 dimension state variables. The observation equtation of integrated navigation system is established by attitude quaternion error defined by attitude quaternion of SINS and star sensor. The integrated navigation system of SINS/CNS simulates by aforementioned CDEKF and the simulation results are analyzed.The observability and observable degree of the system state variables are the key indicators to check the convergence accuracy and velocity of designed filter. The three analysis methods of the system observability and observable degrees are presented contraposing aforementioned typical autonomous navigation systems. Firstly, the observability and influence factor of the directly sensing horizon celestial navigation system is analyzed based on the observability and observability degree. Secondly, the observability of piece-wise constant system(PWCS) is analyzed. Finally, the observability and observable degree of integrated navigation system of SINS/CNS is analyzed by the singular value decomposition and descend the integrated navigation system dimensions. The simulation result and system performance is analyzed and compared with undescended system and descended system after numerical simulation.
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