Unducted, counterrotating gearless front fan engine

• US 4,976,102 A
• Filed: 05/09/1988
• Issued: 12/11/1990
• Est. Priority Date: 05/09/1988
• Status: Expired due to Fees

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Taylor, John B.
• Primary Examiner(s): Casaregola, Louis J.
• Application Number:
  US07/191,529
• Time in Patent Office: 946 Days
• Field of Search: 60/116.1, 60/116.3, 60/39.23, 60/226.1, 60/262, 60/263, 416/170 R, 416/129
• US Class Current: 60/226.1
• CPC Class Codes: B64C 11/308 automatic
• CPC Class Codes: B64C 11/48 Units of two or more coaxial propellers
• CPC Class Codes: B64D 2027/005 Aircraft with an unducted turbofan comprising contra-rotating rotors, e.g. contra-rotating open rotors [CROR]
• CPC Class Codes: F02C 3/067 having counter-rotating rotors F02C3/073 takes precedence
• CPC Class Codes: F02C 7/20 Mounting or supporting of plant; Accommodating heat expansion or creep
• CPC Class Codes: F02C 9/18 by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages F02C3/113 takes precedence
• CPC Class Codes: F02C 9/20 by throttling; by adjusting vanes
• CPC Class Codes: F02C 9/22 by adjusting turbine vanes
• CPC Class Codes: F02K 3/06 with front fan
• CPC Class Codes: F02K 3/072 with counter-rotating , e.g. fan rotors
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine having a core for generating combustion gases, a power turbine, an unducted fan section and a booster compressor. The power turbine includes two counterrotating turbine blade rows which are interdigitized and serve to rotate counterrotating first and second drive shafts, respectively. The unducted fan section also includes counterrotating spaced apart variable pitch fan blade rows which are respectively connected to the first and second drive shafts. A booster compressor is axially positioned between the spaced apart fan blade rows. The booster compressor likewise includes first and second blade rows which are counterrotating and interdigitated and are likewise driven by the first and second drive shafts. The engine is supported by two stationary support frames to permit the nacelle to be nonstructural. The core engine is a modular unit so that it can be separable from the rest of the engine.

Claims

1. A high bypass ratio gas turbine engine comprising:

a core engine effective for generating combustion gases passing through a main flow path;

a power turbine aft of said core engine and including first and second counter rotatable interdigitated turbine blade rows effective for counterrotating first and second drive shafts, respectively;

an unducted fan section forward of said core engine including a first fan blade row connected to said first drive shaft and a second fan blade row axially spaced aftward from said first fan blade row and connected to said second drive shaft, and

a booster compressor axially positioned between said first and second fan blade rows and including a plurality of first compressor blade rows connected to said first drive shaft and a plurality of second compressor blade rows connected to said second drive shaft.

2. A high bypass ratio gas turbine engine as in claim 1, wherein said first and second compressor blade rows are interdigitated.

3. A high bypass ratio gas turbine engine as in claim 1, wherein said first and second unducted fan blade rows are axially spaced apart approximately one and a half times the aero chord length of a blade of the first blade row.

4. A high bypass ratio gas turbine engine as in claim 1 and comprising a pylon for supporting the engine, and wherein the pylon is spaced from the second fan blade row by approximately one aero chord length of a blade of the second blade row.

5. A high bypass ratio gas turbine engine as in claim 1, wherein the booster compressor intake is separate from the fan airflow path and forward of the fan section.

6. A high bypass ratio gas turbine engine as in claim 1, and comprising means for varying the aerodynamic incidence angle of fan blades in said first and second fan blade rows.

7. A high bypass ratio gas turbine engine as in claim 1, wherein, the blades in both said compressor blade rows are fixed pitch blades.

8. A high bypass ratio gas turbine engine as in claim 2, wherein one of said plurality of compressor blades of said compressor blade rows are ganged together and coupled to a variable blade control mechanism to vary the pitch of the blades in at least one compressor blade row.

9. A high bypass ratio gas turbine engine as in claim 1, wherein fan blades in said first and second fane blade rows are all forward swept.

10. A high bypass ratio gas turbine engine as in claim 1, wherein said fan blades are all aft swept.

11. A high bypass ratio gas turbine engine as in claim 1, wherein one fan blade row comprises forward swept blades and the other fan blade row comprises aft swept blades.

12. A high bypass ratio gas turbine engine as in claim 11, wherein the blades of the first blade row are forward swept and the blades of the second blade row are aft swept.

13. A high bypass ratio gas turbine engine as in claim 1, and further comprising a bleed door positioned between said booster compressor and said core engine for relieving the back pressure on the booster compressor to control the stall margin.

14. A high bypass ratio gas turbine engine as in claim 1, and further comprising two counterrotating frames positioned along the main flow path, a respective one on either side of the booster compressor.

15. A high bypass ratio gas turbine engine as in claim 1, further comprising a pair of spaced apart annular stationary support frames on either side of the core engine, one of which is aft of the booster compressor and the other of which is forward of the power turbine.

16. A high bypass ratio gas turbine engine as in claim 15, and comprising an outer non-structurally supporting nacelle.

17. A high bypass ratio gas turbine engine as in claim 16, further comprising an exhaust system which rotates with one of said shafts.

18. A high bypass ratio gas turbine engine as in claim 1, wherein said core engine is removable.

19. A high bypass ratio gas turbine engine as in claim 1, wherein the bypass ratio is between 20 and 35.

20. A high bypass ratio gas turbine engine as in claim 1, further comprising a sharp conical inlet centerbody forward of said fan section.

21. A high bypass ratio gas turbine engine as in claim 1, further comprising a blunt spherical stub nose inlet centerbody forward of said fan section

22. A high bypass ratio gas turbine engine comprising:

a core engine;

a power turbine aft of said core engine;

an unducted fan section forward of said core engine coupled to said power turbine and comprising fan blades;

a booster compressor axially positioned between at least a part of said fan section and said core and including booster compressor blades;

means for varying the pitch on said fan blades, and

means for varying the pitch on said booster compressor blades, whereby the speed of the engine can be controlled through varying the pitch on both the fan and the booster compressor blades while maintaining the power level of the engine.

23. A high bypass ratio gas turbine engine as in claim 22, and comprising means for closing down the fan blades by decreasing the respective pitch angles of the blades and the booster compressor blades whereby the speed is increased for the same power level.

24. A high bypass ratio gas turbine engine as in claim 22, and comprising means for increasing the respective pitch angles of said fan blades and said booster compressor blades for slowing down the engine with a reduced noise condition.