A turbojet engine includes a core engine, an afterburner, and a converging-diverging exhaust nozzle in serial flow communication. A controller is operatively joined to the core engine and afterburner and configured for scheduling fuel thereto for operating the afterburner dry during subsonic flight
A turbojet engine includes a core engine, an afterburner, and a converging-diverging exhaust nozzle in serial flow communication. A controller is operatively joined to the core engine and afterburner and configured for scheduling fuel thereto for operating the afterburner dry during subsonic flight operation of the engine, wet during transonic flight, and dry during supersonic flight.
대표청구항▼
The invention claimed is: 1. A supersonic turbojet engine comprising: a core engine including a multistage axial compressor joined by a rotor to a single-stage high pressure turbine, with an annular combustor disposed therebetween; an afterburner disposed coaxially with an aft end of said core engi
The invention claimed is: 1. A supersonic turbojet engine comprising: a core engine including a multistage axial compressor joined by a rotor to a single-stage high pressure turbine, with an annular combustor disposed therebetween; 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; and a controller operatively joined to said core engine and afterburner and configured for scheduling fuel thereto in a predetermined schedule for operating said afterburner dry during subsonic flight of said turbojet engine, wet during transonic flight, and dry during supersonic flight. 2. An engine according to claim 1 wherein: said compressor includes a row of variable stator rear vanes in the last stage thereof, and a drive train for simultaneously rotating each of said rear vanes; and said controller is operatively joined to said drive train, and is further configured for scheduling rotary position of said vanes to limit speed of said rotor during dry supersonic flight requiring maximum airflow through said compressor to about the speed of said rotor during dry subsonic flight requiring correspondingly less airflow through said compressor. 3. An engine according to claim 2 wherein: said exhaust nozzle includes an inlet duct converging aft to a throat of minimum flow area, and an outlet duct diverging aft therefrom for diffusing said combustion gases discharged therefrom; and said controller is operatively joined to said exhaust nozzle, and is further configured for varying flow area of said throat to minimum area during said dry subsonic flight, maximum area during said wet transonic flight, and intermediate area during said dry supersonic flight. 4. An engine according to claim 3 wherein said compressor further comprises sequential stages of cooperating stator vanes and rotor blades terminating in said variable rear vanes and a cooperating row of last stage rotor blades joined to said rotor, and the first two stages of said vanes are variable and operatively joined by corresponding drive trains to said controller for scheduling rotary position thereof. 5. An engine according to claim 4 wherein said compressor comprises only five stages of said vanes and corresponding blades joined by a common rotor to said turbine having a single row of turbine blades. 6. An engine according to claim 4 wherein said afterburner comprises: a tubular combustion liner mounted concentrically inside an annular casing to define a bypass duct therebetween for receiving said combustion gases from said core engine; and a plurality of fuel spray bars extending radially inwardly at a forward end of said liner and operatively joined to said controller for scheduling fuel flow to said afterburner during operation. 7. An engine according to claim 4 wherein said exhaust nozzle further comprises: a plurality of articulated primary flaps defining said inlet duct, and a plurality of articulated secondary flaps defining said outlet duct; and a drive train operatively joined to said primary and secondary flaps for varying slope thereof and flow area at said throat, and operatively joined to said controller for scheduling said flow area. 8. An engine according to claim 4 further comprising a heat exchanger adjoining said controller and including a flow circuit to initially channel said fuel therethrough for cooling said controller. 9. An engine according to claim 4 wherein: said controller is distributed in discrete and separated modules around the surrounding envelope of said core engine for minimizing size thereof; said afterburner comprises: a tubular combustion liner mounted concentrically inside an annular casing to define a bypass duct therebetween for receiving said combustion gases from said core engine; and a plurality of fuel spray bars extending radially inwardly at a forward end of said liner and operatively joined to said controller for scheduling fuel flow to said afterburner during operation; said exhaust nozzle further comprises: a plurality of articulated primary flaps defining said inlet duct, and a plurality of articulated secondary flaps defining said outlet duct; and a drive train operatively joined to said primary and secondary flaps for varying slope thereof and flow area at said throat, and operatively joined to said controller for scheduling said flow area; and further comprising: a heat exchanger adjoining said controller and including a flow circuit to initially channel said fuel therethrough for cooling said controller. 10. An engine according to claim 9 mounted inside a surrounding tubular engine bay and further including: a power takeoff module and oil sump distributed axially along said core engine with said distributed controller modules for minimizing diameter of said bay; and said exhaust nozzle includes an ejector inlet at a forward end for ejector cooling said secondary flaps. 11. A turbojet engine for powering a supersonic missile comprising: a core engine including a multistage axial compressor joined by a rotor to a high pressure turbine, with an annular combustor disposed therebetween; 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; and a controller operatively joined to said core engine and afterburner and configured for scheduling fuel thereto in a predetermined schedule for operating said afterburner dry during subsonic flight of said missile, wet during transonic flight, and dry during supersonic flight. 12. A turbojet engine for powering a supersonic missile comprising: a core engine including a multistage axial compressor joined by a rotor to a high pressure turbine, with an annular combustor disposed therebetween, and said compressor includes a row of variable stator rear vanes in the last stage thereof; 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 controller operatively joined to said core engine and afterburner and configured for scheduling fuel thereto for operating said afterburner dry during subsonic flight of said missile, wet during transonic flight, and dry during supersonic flight; and said controller is operatively joined to said rear vanes, and is further configured for scheduling rotary position of said vanes to limit speed of said rotor during dry supersonic flight requiring maximum airflow through said compressor to about the speed of said rotor during dry subsonic flight requiring correspondingly less airflow through said compressor. 13. An engine according to claim 12 wherein: said exhaust nozzle includes an inlet duct converging aft to a throat of minimum flow area, and an outlet duct diverging aft therefrom for diffusing said combustion gases discharged therefrom; and said controller is operatively joined to said exhaust nozzle, and is further configured for varying flow area of said throat to minimum area during said dry subsonic flight, maximum area during said wet transonic flight, and intermediate area during said dry supersonic flight. 14. An engine according to claim 13 wherein said compressor further comprises sequential stages of cooperating stator vanes and rotor blades terminating in said variable rear vanes and a cooperating row of last stage rotor blades joined to said rotor, and the first two stages of said vanes are variable and operatively joined to said controller for scheduling rotary position thereof. 15. An engine according to claim 14 wherein said compressor comprises only five stages of said vanes and corresponding blades joined by said rotor to said turbine having a single stage. 16. An engine according to claim 14 wherein said afterburner comprises: a tubular combustion liner mounted concentrically inside an annular casing to define a bypass duct therebetween for receiving said combustion gases from said core engine; and a plurality of fuel spray bars extending radially inwardly at a forward end of said liner and operatively joined to said controller for scheduling fuel flow to said afterburner during operation. 17. An engine according to claim 14 wherein said exhaust nozzle further comprises: a plurality of articulated primary flaps defining said inlet duct, and a plurality of articulated secondary flaps defining said outlet duct; and a drive train operatively joined to said primary and secondary flaps for varying slope thereof and flow area at said throat, and operatively joined to said controller for scheduling said flow area. 18. An engine according to claim 14 further comprising a heat exchanger adjoining said controller and including a flow circuit to initially channel said fuel therethrough for cooling said controller. 19. An engine according to claim 14 wherein said controller is distributed in discrete and separated modules around the surrounding envelope of said core engine for minimizing size thereof. 20. An engine according to claim 19 in combination with said supersonic missile and mounted inside an engine bay at an aft end thereof, and said turbojet engine further includes an inlet duct extending forward from said core engine and forwardly through the side of said missile for receiving ambient air during subsonic to supersonic operation.
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