Key Techniques For Inertial Stellar Integrated Navigation System

This dissertation investigates in inertial stellar navigation system for aircraft under high altitude,long duration and for roundtrip spacecraft in near-earth orbit,taking research content mainly including air-born inertial stellar navigation technique based on small FOV(field of view)star tracker and low orbit inertial stellar navigation technique based on large FOV starlight sensor.Meanwhile,for high altitude,high speed and large maneuver vehicles,deep research is done for SINS updating algorithm and anti-interference initial alignment algorithm.The major work done is summed up as follows:1.For high altitude,high speed,large maneuver vehicles,SINS updating algorithm is improved based on traditional multi-sample algorithm.Within the basic framework of “high-frequency multi-sample optimization” plus “intermediate frequency updating calculation",a parameter interpolation operator is introduced to improve navigation performance under high speed and maneuvering conditions.The influence of earth rotation model error and gravity model error is studied in the view of navigation accuracy.The difference between WGS84 geodetic gravity model and spherical harmonic gravity model is analyzed according to latitude and altitude change,and concluded by simulation that the latter model is more suitable when the inertial navigation is under high altitude.2.Anti-interference alignment algorithm is investigated for aircraft take-off and rocket launching procedure,which composes of two steps as indirect resolving coarse alignment and state estimating fine alignment.The former algorithm is based on gravity vector observation in inertial frame,which is stable under angular maneuver.Construction method for gravity vectors is improved to reduce the error divergence,and align performance under linear interference is analyzed and enhanced in the view of angle between vectors.Furthermore,multi-vector determination methods as QUEST and FOAM are introduced into the algorithm,which expands the algorithm into moving vehicles.State estimating fine alignment is based on SINS error model and Kalman filter,which uses the coupling relationship between attitude error and velocity error to estimate attitude error.Precision of the fine alignment has great relationship with initial attitude error.To solve this problem,the influence of initial covariance distribution in Kalman filter is quantitatively analyzed,and an optimal distribution strategy is proposed,which achieves the same precision at any initial attitude error.3.Integrated navigation algorithm composed of small FOV star tracker and SINS is investigated,for autonomous navigation under high altitude,long duration aircraft.The working and navigation principle for star tracker is deduced,and the output models for elevation/azimuth error angle from one star observation and psi-angel from multi-star observation is given.The relationship function between star tracker outputs,attitude error and position error is derived.On this basis,two integration schemes respectively based on "position + orientation" correcting and SINS error model filtering are designed,and navigation accuracy is simulated and compared.Furthermore,an after-flight correcting algorithm is proposed in the case that the sky is covered and the star tracker can not work at low altitude.To further improve the navigation precision,a composed alignment algorithm is designed by adding star tracker measurement into fine alignment procedure,which can estimate both attitude error and accelerometer error.Finally,the influence of installation error is analyzed and calibration scheme is given both in laboratory and outside.4.The inertial/ stellar/ GNSS integration algorithm is investigated for round-trip aircraft.The matching measurement for large FOV starlight sensor and SINS under different navigation frame is derived,and the consistence is proved.The equivalence between star tracker and large FOV starlight sensor is proved when they are integrated with SINS,and the relationship function is deduced.Time delay effect of starlight sensor on navigation precision is analyzed,and improving solution is designed.Meanwhile,High precision translation algorithm between protocol inertial frame and protocol earth frame is designed and verified by STK trajectory data.For round-trip aircraft,a integrated navigation scheme is designed for different flight stages.Finally,SINS characteristic under vibration environment is deduced considering the vibration effect caused by friction between aircraft and atmosphere.5.Simulation and testing system for inertial stellar integration is investigated.Digital simulation system is designed based on flight trajectory simulation,which includes the flight simulation under different maneuvers,inertia device error,installation error,interference,flexure,arm error,and attack/slip angles.Simplified orbit trajectory is simulated under earth sphere model to validate the integration algorithm.Furthermore,a complicated orbit simulation algorithm is designed based on STK software.A stellar inertial testing system is established including inertial system,stellar equipment,dual-axis turntable and starlight simulator.And a landing vehicle testing system is established to verify the algorithm on ground.