The combustion gas temperature measurement system of rocket motor was established after temperature measurement principle utilizing Alexandrite effect was introduced, and the influences of the engine diameter and the heat loss on the axis temperature of combustion chamber were studied by numerical simulation. The combustion gas temperatures of three composite propellants with 1%, 9%, 17% aluminum at atmosphere(0.1 MPa), combustion chamber and nozzle exit with the motor working pressure of 5 MPa were measured. Results show that the engine diameter and the heat loss have little effect on the axis temperature of combustion chamber. The experimental combustion gas temperatures at atmosphere are 2857, 3109 K and 3284 K respectively, and the theoretical combustion gas temperatures are 2712, 2891 K and 3049 K, which demonstrates that with the increasing of aluminum content, the experimental and theoretical values all increase. The measured combustion gas temperature at nozzle exit is 2200 K, which agree with 2278 K through theoretical calculation. The measured gas temperature at combustion chamber are 2300 K and 2450 K, considerably lower than the theoretical values of 3190 K and 3450 K, which is probably because of the transparent glass window contaminated by combustion gas during engine working process. In order to accurately measure the gas temperature of rocket engine, the temperature measurement system must be further improved.