Patent ID: 7475545

Claim:
A supersonic turbojet engine comprising: a core engine including a multistage axial compressor having sequential stages of stator vanes and rotor blades, with the last stage of rear vanes being variable and discharging pressurized airflow to a combustor followed by a high pressure turbine, and said turbine is joined to said rotor blades by a rotor; an afterburner disposed coaxially with an aft end of said core engine for receiving combustion gases therefrom; a converging-diverging exhaust nozzle disposed coaxially with an aft end of said afterburner for discharging said combustion gases; a bypass duct surrounding said core engine and afterburner and terminating in flow communication with said exhaust nozzle; said compressor including a row of first stage fan blades extending from a supporting rotor disk, and each of said blades having an integral f lade extending from the tip thereof into said bypass duct; and said bypass duct includes a row of variable inlet guide vanes disposed forward of said blades for controlling airflow thereto.