The strap-down inertial navigation system/celestial navigation system(SINS/CNS) integrated system can be divided into the loosely coupled integrated mode and the tightly coupled integrated mode. Because the loosely coupled integrated mode requires the star sensor to observe at least three stars at the same time to obtain the attitude matrix, the tightly coupled integrated mode, which can still work with only one star is more practical in reality, especially for aircraft that need to work in the daytime. The star sensor is a vital part of the CNS, while its installation error impacts the navigation accuracy seriously in both modes. Therefore, the star sensor installation error of SINS/CNS integrated navigation system must be corrected. The installation error calibration has been settled for the loosely coupled SINS/CNS integrated navigation system when the attitude matrix is available. However, it is still a problem for the tightly coupled mode. This paper proposes a fast calibration method of the star sensor installation error for the tightly coupled SINS/CNS integrated navigation system based on maneuvers and observability analysis. Simulations indicate that the proposed method is feasible and effective when only one star is visible at a time. The mean position error of the tightly coupled SINS/CNS integrated navigation decreases by about 63.23% after compensation for the star sensor installation error.

The combination of celestial measurement and ground Doppler measurement has been successfully used for orbit determination in many deep-space exploration missions. However, one of the inherent drawbacks of ground tracking is the long communication delay. Using the Doppler measurement from the sun instead of the ground station is a feasible strategy for real-time navigation, but the accuracy of this Doppler measurement is affected by the frequency fluctuation of the solar spectrum. To solve this problem, a differential Doppler measurement-aided celestial navigation method is proposed in this paper, which uses the differential Doppler measurement to reduce the influence caused by instability of the solar spectrum frequency and improve the navigation accuracy. Simulations demonstrate the feasibility and effectiveness of this method.

Solar oscillations cause significant variations in sunlight spectral wavelength and intensity over short time periods. These oscillations have been studied in detail over the years, both theoretically and using actual observations. Through detecting the sunlight spectral central wavelengths and intensity and recording the moment when solar oscillation occurs, the time delay between the sunlight coming from the Sun directly and the sunlight reflected by a celestial body such as the satellite of a planet or asteroid can be obtained. Because the solar oscillation time delay is determined by the relative positions of the spacecraft, the reflective celestial body and the Sun, it can be adopted as the navigation measurement to provide the spacecraft's position information. In this paper, a novel celestial navigation method using the solar oscillation time delay measurement is proposed. The implicit measurement model of time delay is built, and the Implicit Unscented Kalman Filter (IUKF) is applied. Simulation results indicate that the position error and velocity error of the proposed method for the transfer orbit are about 3.55 km and 0.077 m/s, respectively, and those for the surrounding orbit are about 1.76 km and 1.57 m/s, respectively. The impact of four factors on the navigation performance is also investigated.