As the errors of the Inertial Navigation System (INS) in a launch vehicle upper stage increase significantly with time and cannot meet the requirements of long working hours,an INS/CNS (Celestial Navigation System) integrated navigation system is studied. Firstly, the coordinate systems and the attitude transformation matrix are defined. Then, the state equation and measurement equation of the integrated navigation system are raised from the INS error propagation equations and measurements of the star tracker and earth sensor. Finally, a semi-physical simulation experiment system with the star tracker in the loop is designed based on Matlab/dSpace simulation environment. Experimental results indicate that the integrated navigation system decreases the inertial navigation position errors from 1.171 9×104 m to 1.036 7×103 m, the velocity errors from 11.282 7 m/s to 3.662 6 m/s, and the attitude errors from 0.1° to 5′. Furthermore,the experimental results show that the integrated navigation system can correct the inertial navigation errors effectively, and also confirm that the INS/CNS integrated navigation system is feasible and appropriate.