Fladed supersonic missile turbojet

1. 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.

2. An engine according to claim 1 wherein:

said core engine includes an inner inlet in front of a first stage of said stator vanes;

said bypass duct includes an outer inlet at the forward end thereof in front of said guide vanes and surrounding said inner inlet; and

said guide vanes are operatively joined to a controller configured for controlling airflow into said bypass duct for matching airflow supply to said inner inlet to demand therefor in said core engine over an operating cycle including subsonic flight of a missile, transonic flight, and supersonic flight.

3. An engine according to claim 2 wherein said bypass duct further includes a row of outlet guide vanes disposed downstream behind said flades, and said flades are sized and configured for pressurizing airflow in said bypass duct for flow through said exhaust nozzle into said combustion gases discharged therethrough.

4. An engine according to claim 3 wherein:

said bypass duct includes in part the casing of said afterburner;

said exhaust nozzle includes an inlet duct converging to a throat of minimum flow area, and an outlet duct diverging aft therefrom and disposed in flow communication with said bypass duct; and

said flades are configured to pressurize said airflow for flow through said bypass duct to cool said afterburner casing and said outlet duct.

5. An engine according to claim 4 wherein said bypass duct terminates in a flow distribution manifold disposed at said exhaust nozzle, and said manifold is joined in flow communication with said exhaust nozzle for discharging airflow into said nozzle outlet duct.

6. An engine according to claim 5 wherein said controller is operatively joined to said rear vanes, and is further configured for scheduling rotary position of said rear vanes to limit speed of said rotor during supersonic flight requiring maximum airflow through said compressor to about the speed of said rotor during subsonic flight requiring less airflow through said compressor.

7. An engine according to claim 5 wherein:

a plurality of articulated primary flaps defining said inlet duct, and a plurality of articulated secondary flaps defining said outlet duct; and

said manifold is joined in flow communication with said secondary flaps.

8. An engine according to claim 7 wherein said manifold includes a plurality of distribution valves cooperating therewith and operatively joined to said controller for varying bypass flow into said exhaust nozzle to effect thrust vectoring.

9. An engine according to claim 5 wherein:

said exhaust nozzle comprises a nonvariable inlet duct and a nonvariable outlet duct, and said throat has a fixed flow area; and

said manifold is joined in flow communication with said throat for fluidically varying flow area thereof by injecting pressurized airflow from said bypass duct therein.

10. An engine according to claim 9 further comprising a variable area bypass injector disposed between said core engine and said afterburner in flow communication with said bypass duct for selectively injecting a portion of said airflow from said bypass duct inside said afterburner to join exhaust flow from said core engine in collective flow through said fixed throat.

11. A single-rotor turbojet engine for powering a supersonic missile comprising:

a core engine including a multistage axial compressor joined by said single 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;

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 flade 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.

12. A turbojet engine comprising:

an annular intake at a forward end of said turbojet engine;

a core engine disposed in direct flow communication with said intake, and 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;

a bypass duct surrounding said core engine and afterburner and terminating in flow communication with said exhaust nozzle;

a row of first stage fan blades extending from a supporting rotor disk in said compressor, and each of said blades having an integral flade extending from the tip thereof into said bypass duct;

said core engine further includes an inner inlet at the forward end thereof in front of said fan blades, and disposed in flow communication with said intake;

said bypass duct includes a row of variable inlet guide vanes disposed forward of said blades for controlling airflow thereto;

said bypass duct further includes an outer inlet at the forward end thereof in front of said guide vanes and surrounding said inner inlet in common flow communication with said intake; and

said guide vanes are operatively joined to a controller configured for controlling airflow into said bypass duct for matching airflow supply to said inner inlet to demand therefor in said core engine over an operating cycle including subsonic flight of a missile, transonic flight, and supersonic flight.

13. An engine according to claim 12 wherein said bypass duct further includes a row of outlet guide vanes disposed downstream behind said flades, and said flades are sized and configured for pressurizing airflow in said bypass duct for flow through said exhaust nozzle into said combustion gases discharged therethrough.

14. An engine according to claim 13 wherein:

said bypass duct includes in part the casing of said afterburner;

said exhaust nozzle includes an inlet duct converging to a throat of minimum flow area, and an outlet duct diverging aft therefrom and disposed in flow communication with said bypass duct; and

said flades are configured to pressurize said airflow for flow through said bypass duct to cool said afterburner casing and said outlet duct.

15. An engine according to claim 14 wherein:

said compressor includes a row of variable stator rear vanes in the last stage thereof; and

said controller is operatively joined to said rear vanes, and is further configured for scheduling rotary position of said rear vanes to limit speed of said rotor during supersonic flight requiring maximum airflow through said compressor to about the speed of said rotor during subsonic flight requiring less airflow through said compressor.

16. An engine according to claim 15 wherein said bypass duct terminates in a flow distribution manifold disposed at said exhaust nozzle, and said manifold is joined in flow communication with said exhaust nozzle for discharging airflow into said nozzle outlet duct.

17. An engine according to claim 16 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

said manifold is joined in flow communication with said secondary flaps.

18. An engine according to claim 17 wherein said manifold includes a plurality of distribution valves cooperating therewith and operatively joined to said controller for varying bypass flow into said exhaust nozzle to effect thrust vectoring.

19. An engine according to claim 16 wherein:

said exhaust nozzle comprises a nonvariable inlet duct and a nonvariable outlet duct, and said throat has a fixed flow area; and

said manifold is joined in flow communication with said throat for fluidically varying flow area thereof by injecting pressurized airflow from said bypass duct therein.

20. An engine according to claim 19 further comprising a variable area bypass injector disposed between said core engine and said afterburner in flow communication with said bypass duct for selectively injecting a portion of said airflow from said bypass duct inside said afterburner to join exhaust flow from said core engine in collective flow through said fixed throat.

21. An engine according to claim 19 wherein said manifold includes a plurality of distribution valves cooperating therewith and operatively joined to said controller for varying bypass flow into said exhaust nozzle to effect thrust vectoring.