This paper presents results of a thermal analysis performed on a gaseous oxygen/Polymethylmethacrilate hybrid rocket motor for a CubeSat application. This work was focused on deriving temperatures inside the combustion chamber and verifying the selected materials can withstand the high temperatures and stresses. Results show that the fuel starts to break-up before the end of the burn. The addition of a graphite post-combustion chamber was also evaluated but it is not recommended for long burn times due to high temperature reached in the steel. This work is an important intermediate step in the design process to determine system performance.