Research on transcritical heat transfer characteristic of thrust chamber in high-thrust liquid oxygen/methane rocket engines

Thermal protection of the thrust chamber is essential for the stable operation of high-thrust liquid oxygen/methane rocket engines. Currently, regenerative cooling and film cooling are common and well-developed thermal protection methods. For the high-thrust liquid oxygen/methane rocket engine, the transcritical methane heat transfer characteristics of the thrust chamber have not been fully investigated. In this paper, a comprehensive model of heat transfer has been developed for the thrust chamber combining regenerative cooling and film cooling. The chamber wall temperature and heat flux are solved by using a set of semi-empirical heat transfer criteria formulas. In addition, the factors influencing the thrust chamber cooling are investigated. It is found that an increase in coolant mass flow rate results in a decrease in gas-side wall temperature and heat flux. When the film coolant mass flow rate increases from zero to 10% of the regenerative coolant mass flow rate, the maximum gas-side wall temperature reduces from 1119K to 807K, and the maximum heat flux reduces from 173MW/m² to 119MW/m². Also, the results indicate that the injection of film coolant near the head is an effective way for thermal protection of high-thrust liquid oxygen/methane engines. When the film coolant inlet position is at the head of the chamber, the near-wall gas temperature at the film coolant inlet is only 296K, increasing to 580K when the inlet is 0.1m downstream from the head. These results contribute to the design of the thrust chamber cooling system.