In the aerospace field, aiming at the requirements of thermal response monitoring of composite structures in space service environment, a monitoring method for carbon fiber honeycomb sandwich structure based on fiber Bragg grating (FBG) reflection spectrum characteristics analysis under different thermal loads was studied. The fiber Bragg grating (FBG) sensors were implanted in different layers of carbon fiber honeycomb sandwich structure respectively. By monitoring the fiber Bragg grating (FBG) reflection spectrums of each layer under different thermal loads, the associated thermal strain characteristics of the honeycomb sandwich structure can be obtained. The results show that, certain differences exist in the thermal strain characteristics of different material layers in the carbon fiber honeycomb sandwich structure. As the temperature increases, the center wavelength of reflection spectrum of fiber Bragg grating (FBG) implanted between the outer skin surface and the glass cloth shifts to the long wave direction and the waveform of the obtained reflection spectrum changes insignificantly. As the temperature decreases, the reflection spectrum of fiber Bragg grating (FBG) implanted between the second and the third layers of the carbon fabric prepreg appears chirp effects gradually, such as sidelobe, multi-peak, etc. The main-peak center wavelength and the right sub-peak center wavelength of the reflection spectrum both shift to the short wave direction gradually. The peak amplitude of the main-peak does not show significant change and its temperature sensitivity is about 5.56×10-3 dBm·℃-1. The peak amplitude of the right sub-peak increases significantly and its temperature sensitivity is about 40.32×10-3 dBm·℃-1. The full width at half maximum（FWHM）of reflection spectrum of fiber Bragg grating (FBG) implanted between the inner skin and the honeycomb core increases gradually at a rate of 3.19 pm·℃-1 with the temperature decreasing. The reflection spectrum appears a multi-peak trend significantly due to the uneven thermal stress distributed between the layers. The interlaminar thermal strain of each implanted layer increases in similar trends within the increasing temperature range of -70 ℃ to +60 ℃. But at the temperature range of +60 ℃ to +120 ℃, the changing trend of the interlaminar thermal strain between each implanted layers shows significant differences. These characteristics can provide useful help to the following on-orbit status monitoring of composite spacecraft structure in space environment.