초록 ▼

A turbojet engine includes a core engine, an afterburner, and a converging-diverging exhaust nozzle in serial flow communication. A thrust vectoring system is joined between a compressor and the nozzle. A controller is operatively joined to the thrust vectoring system for selectively varying distrib

A turbojet engine includes a core engine, an afterburner, and a converging-diverging exhaust nozzle in serial flow communication. A thrust vectoring system is joined between a compressor and the nozzle. A controller is operatively joined to the thrust vectoring system for selectively varying distribution of air bled from the compressor into the exhaust nozzle for vectoring propulsion thrust.

대표청구항 ▼

Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims in which I claim: 1. A supersonic turbojet engine comprising: a core engine including a multistage axial compressor joined by a rotor to a high pr

Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims in which I claim: 1. A supersonic turbojet engine 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; an exhaust nozzle including 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; a thrust vectoring system operatively joined to said compressor for receiving pressurized air therefrom, and operatively joined to said nozzle inlet and outlet ducts for selectively discharging said pressurized air therein; and a controller operatively joined to said thrust vectoring system for varying distribution of said pressurized air around said nozzle and thereby vector thrust from said combustion gases discharged through said nozzle. 2. An engine according to claim 1 wherein said thrust vectoring system comprises: a plurality of distribution valves joined in flow communication with said compressor by a plurality of inlet conduits, and joined in flow communication with said nozzle inlet and outlet ducts by a plurality of outlet conduits spaced circumferentially therearound; and said controller is operatively joined to said valves and configured for adjusting flow of said pressurized air through said valves for varying distribution thereof to vector said thrust. 3. An engine according to claim 2 wherein said outlet conduits comprise: a plurality of forward outlet conduits joined to said converging inlet duct for discharging said pressurized air therein; and a plurality of aft outlet conduits joined to said diverging outlet duct for discharging said pressurized air therein. 4. An engine according to claim 3 wherein said thrust vectoring system further comprises: a distribution manifold joined in flow communication with said distribution valves by said inlet conduits; and a plurality of bleed conduits joining said manifold in flow communication with said compressor. 5. An engine according to claim 4 wherein said distribution valves are configured for splitting said pressurized air between said forward and aft outlet conduits for discharge into said inlet and outlet ducts, respectively. 6. An engine according to claim 5 wherein said distribution valves are configured for increasing flow to said aft outlet conduits while decreasing flow to said forward outlet conduits, and vice versa. 7. An engine according to claim 6 wherein said distribution valves are configured for distributing flow uniformly through said forward outlet conduits to effect pressure recovery in said nozzle, and distributing flow nonuniformly through said aft outlet conduits to effect said thrust vectoring. 8. An engine according to claim 7 wherein: said compressor includes sequential stages for progressively increasing pressure of said air; and said bleed conduits are joined in flow communication with an intermediate stage of said compressor for bleeding interstage air therefrom. 9. An engine according to claim 7 wherein: said compressor includes sequential stages for progressively increasing pressure of said air; and said bleed conduits are joined in flow communication with the discharge end of said compressor for bleeding compressor discharge air therefrom. 10. An engine according to claim 7 wherein: said bleed conduits, manifold, and inlet conduits provide an unobstructed flowpath from said compressor to said distribution valves; and said distribution valves are configured for continuous flow. 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 having a variable area throat; and a thrust vectoring system operatively joined to said compressor for receiving pressurized air therefrom, and operatively joined to said nozzle for selectively discharging said pressurized air therein axially spaced from said throat; and a controller operatively joined to said thrust vectoring system for varying distribution of said pressurized air around said nozzle and thereby vector thrust from said combustion gases discharged through said nozzle. 12. An engine according to claim 11 wherein said thrust vectoring system comprises: a plurality of distribution valves joined in flow communication with said compressor by a plurality of inlet conduits, and joined in flow communication with said nozzle by a plurality of outlet conduits spaced circumferentially therearound; and said controller is operatively joined to said valves and configured for adjusting flow of said pressurized air through said valves for varying distribution thereof to vector said thrust. 13. 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 including an inlet duct joined coaxially with an aft end of said afterburner and converging aft to a throat followed by a diverging outlet duct for discharging said combustion gases; a thrust vectoring system operatively joined to said compressor for receiving pressurized air therefrom, and including forward outlet conduits joined to said converging inlet duct forward from said throat, and aft outlet conduits joined to said diverging outlet duct aft of said throat for selectively discharging said pressurized air therein; and a controller operatively joined to said thrust vectoring system for varying distribution of said pressurized air through said forward and aft outlet conduits and thereby vector thrust from said combustion gases discharged through said nozzle. 14. An engine according to claim 13 wherein said thrust vectoring system further comprises: a plurality of distribution valves joined in flow communication with said compressor by a plurality of inlet conduits, and joined in flow communication with said nozzle by said forward and aft outlet conduits spaced circumferentially around said nozzle; said distribution valves are configured for splitting said pressurized air between said forward and aft outlet conduits for discharge into said inlet and outlet ducts, respectively; and said controller is operatively joined to said valves and configured for adjusting flow of said pressurized air through said valves for varying distribution thereof to vector said thrust. 15. An engine according to claim 14 wherein said distribution valves are configured for increasing flow to said aft outlet conduits while decreasing flow to said forward outlet conduits, and vice versa. 16. An engine according to claim 14 wherein said distribution valves are configured for distributing flow uniformly through said forward outlet conduits to effect pressure recovery in said nozzle, and distributing flow nonuniformly through said aft outlet conduits to effect said thrust vectoring. 17. An engine according to claim 14 wherein said thrust vectoring system further comprises: a distribution manifold joined in flow communication with said distribution valves by said inlet conduits; and a plurality of bleed conduits joining said manifold in flow communication with said compressor. 18. An engine according to claim 17 wherein: said compressor includes sequential stages for progressively increasing pressure of said air; and said bleed conduits are joined in flow communication with an intermediate stage of said compressor for bleeding interstage air therefrom. 19. An engine according to claim 17 wherein: said compressor includes sequential stages for progressively increasing pressure of said air; and said bleed conduits are joined in flow communication with the discharge end of said compressor for bleeding compressor discharge air therefrom. 20. An engine according to claim 17 wherein: said bleed conduits, manifold, and inlet conduits provide an unobstructed flowpath from said compressor to said distribution valves; and said distribution valves are configured for continuous flow. 21. An engine according to claim 17 in combination with said supersonic missile and mounted inside an engine bay at an aft end thereof, and said turbojet engine further includes an intake extending forward from said core engine and forwardly through the side of said missile for receiving ambient air during subsonic to supersonic operation.