Abstract: The compressibility correction of the calculation method is conducted to improve the simulation ability of high Mach number scramjet. The three-dimensional numerical simulations of the scramjet at Mach 12 flight condition are carried out. The shock system, parameters characteristics and combustion performance of the scramjet are analyzed. The results indicate that the position and intensity of shock wave calculated by the modified method are consistent with the experimental data. The modified method shows better ability in shock wave and high Ma scramjet simulation. The shock wave and reflected shock systems are formed in the scramjet. The basic structure of shock system through the flow path will not be changed by the combustion. The angle and number of shock waves will increase with the increase of equivalence ratio. The temperature rise and pressure rise induced by the intersection of shock waves are conducive to combustion heat release. The increase of wall heat flux caused by shock waves gradually reduces with the weakening of reflected shock waves along the flow direction. Most of the combustion is non-premixed in the flow field. The average temperature in the combustor exceeds 2500 K so that the efficiency of H₂-O₂ combustion deteriorates and the complete combustion product H₂O decreases. The available effective heat release increases in the forepart of the combustor and decreases in the rear section. O atom recombination mainly occurs in the nozzle. The chemical reactions are mainly conducted in the forepart of the combustor at the equivalence ratio of 0.5, while the reaction distance is longer at the equivalence ratio of 1.0. The difference between combustor drags of the two cases is small and the increase of total thrust coefficient is mainly contributed by the nozzle. Combustion will lead to the reduction of the combustor friction and whole model friction, while the changes of inlet friction and nozzle friction are not obvious.