Leading-to-trailing edge theoretical design of a generic scramjet

Abstract

The aerospace vehicle flying at hypersonic speed, using an airbreathing propulsion system based on supersonic combustion (scramjet) technology, requires a highly integrated system without moving parts, making the propulsion system and vehicle shape indistinguishable. A scramjet uses shock waves, generated by its inlet during the hypersonic flight to provide the temperature and velocity of atmospheric air to burn the hydrogen, at the combustor at supersonic speed. The divergent exhaust nozzle accelerates the combustion products providing the thrust. A two-dimensional hydrogen-powered generic scramjet has been designed to demonstrate supersonic combustion in atmospheric flight at a Mach number of 6.8 and an altitude of 30 km. Temperature and velocity at the combustion chamber are the most important key parameters for a preliminary design. The inlet configuration must provide a temperature higher than the ignition temperature of the fuel to guarantee spontaneous combustion and the velocity must remain supersonics. The nozzle exit velocity should be higher than the vehicle flight velocity to obtain sustained flight and produce enough thrust. For this theoretical analysis, the air is considered a calorically perfect gas without viscous effects. The compressible flow is analyzed based on oblique shock waves and one-dimensional compressible flow with friction and heat addition and Prandtl–Meyer coupled to the area ratio to describe each key component of the scramjet, such as compression, combustor, and nozzle sections, respectively. The inlet mass flow and hydrogen mass flow are also critical for the scramjet design to ensure stoichiometrically combustion at the combustion chamber, generating the high flow velocity at the nozzle to produce thrust.