Turbojet engines of multi-shaft and multi-flow construction

• US 3,886,737 A
• Filed: 08/22/1973
• Issued: 06/03/1975
• Est. Priority Date: 08/22/1972
• Status: Expired due to Term

• Current Assignee: Hubert Grieb
• Original Assignee: Hubert Grieb
• Inventors: Grieb, Hubert
• Primary Examiner(s): Husar, C. J.
• Assistant Examiner(s): Garrett, Robert E.
• Application Number:
  US05/390,485
• Time in Patent Office: 650 Days
• Field of Search: 60/224,226R,226A,262,263,271 239/265.13 181/33HB,33HC,33HD
• US Class Current: 60/226.1
• CPC Class Codes: F02K 3/077 the plant being of the multiple flow type, i.e. having three or more flows
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A turbojet engine constructed for minimizing noise level generation which includes two separate annular concentric bypass air ducts extending around the hot exhaust gas stream from the combustion chamber. The fans energizing the respective annular ducts and the ducts themselves are configured such that the exit velocity for the annular bypass air ducts is decreased from the longitudinal centerline of the engine outwardly in a manner corresponding approximately to the velocity profile of the turbulent free jet formed downstream of the propulsion nozzle arrangement. Additional silencing means are provided within the duct and between the fans. Further silencing measures include specific arrangement of splitters for the flow through the various ducts from the more forward fans as well as specific configuration of the fan blades to minimize sound generation from the fan blades themselves.

Claims

1. A turbojet engine comprising:

first fluid energizing means for energizing a first fluid stream, second fluid energizing means for energizing a second fluid stream, third fluid energizing means for energizing a third fluid stream, first duct means for conducting said first fluid stream to a propulsion nozzle arrangement where said first fluid stream is discharged at a first jet velocity, second duct means for conducting said second fluid stream to said propulsion nozzle arrangement where said second fluid stream is discharged in annular surrounding relationship to said first fluid stream and at a second jet velocity, third duct means for coNducting said third fluid stream to said propulsion nozzle arrangement where said third fluid stream is discharged in annular surrounding relationship to said second fluid stream and at a third jet velocity, wherein said first jet velocity is greater than said second jet velocity and said second jet velocity is greater than said third jet velocity. wherein said third fluid energizing means includes a radially outer portion of a first rotatable fan means arranged immediately upstream of inlet means to said third duct means, wherein said second fluid energizing means includes a second rotatable fan means arranged downstream of said first fan means and immediately upstream of inlet means of said second duct means, wherein flow into all of said duct means passes through said first fan means, wherein flow into only said first and second duct means passes through said second fan means, and wherein said first and second fan means operate at different rotational speeds with respect to one another.

2. A turbojet engine according to claim 1, wherein the relationship between said first, second and third jet velocities corresponds to an optimum velocity profile of the turbulent free jet formed downstream of the propulsion nozzle arrangement.

3. A turbojet engine according to claim 1, further comprising acoustic silencer means in at least one of said duct means.

4. A turbojet engine according to claim 3, wherein said acoustic silencer means are provided in both said second and third duct means.

5. A turbojet engine comprising:

first fluid energizing means for energizing a first fluid stream, second fluid energizing means for energizing a second fluid stream, third fluid energizing means for energizing a third fluid stream, first duct means for conducting said first fluid stream to a propulsion nozzle arrangement where said first fluid stream is discharged at a first jet velocity, second duct means for conducting said second fluid stream to said propulsion nozzle arrangement where said second fluid stream is discharged in annular surrounding relationship to said first fluid stream and at a second jet velocity, third duct means for conducting said third fluid stream to said propulsion nozzle arrangement where said third fluid stream is discharged in annular surrounding relationship to said second fluid stream and at a third jet velocity, wherein said first jet velocity is greater than said second jet velocity and said second jet velocity is greater than said third jet velocity, wherein said first fluid energizing means includes combustion chamber means such that said first fluid stream includes hot combustion chamber exhaust products, wherein said third fluid energizing means includes a radially outer portion of a first rotatable fan means arranged immediately upstream of inlet means to said third duct means, wherein said second fluid energizing means includes a second rotatable fan means arranged downstream of said first fan means and immediately upstream of inlet means of said second duct means, wherein flow into all of said duct means passes through said first fan means, wherein flow into only said first and second duct means passes through said second fan means, and wherein said first and second fan means operate at different rotational speeds with respect to one another.

6. A turbojet engine according to claim 5, wherein each of said second and third duct means are in bypassing relationship to said combustion chamber means.

7. A turbojet engine according to claim 6, wherein the relationship between said first, second and third jet velocities corresponds to an optimum velocity profile of the turbulent free jet formed downstream of the propulsion nozzle arrangement.

8. A turbojet engine according to claim 5, wherein said first and second fan means are arranged on separate shafts driven by separate turbines, and wherein said turbines are arranged downstream of said combustion chamber means and are driven by exhauSt gas flow from said combustion chamber means.

9. A turbojet engine according to claim 7, wherein said first and second fan means are arranged on separate shafts driven by separate turbines, and wherein said turbines are arranged downstream of said combustion chamber means and are driven by exhaust gas flow from said combustion chamber means.

10. A turbojet engine according to claim 8, wherein said second fan means is operatively drivingly unified with a compressor arranged immediately downstream thereof, and wherein flow into said first duct means passes through said compressor.

11. A turbojet engine according to claim 10, wherein said compressor is arranged immediately upstream of said combustion chamber means and supplies combustion air to said combustion chamber means.

12. A turbojet engine according to claim 5, wherein said first and second fan means are drivingly coupled to one another by reducer gear means.

13. A turbojet engine according to claim 5, wherein all of said duct means extend coaxially with the longitudinal centerline of the engine, and wherein said duct means exit from said propulsion nozzle arrangement in respective exit planes staggerd relative to one another in the longitudinal direction of the engine.

14. A turbojet engine according to claim 13, wherein the exit plane for said duct means is behind the exit plane for said second duct means and the exit plane for said second duct means is behind the exit plane for said third duct means in the direction of fluid flow through the propulsion nozzle arrangement.

15. A turbojet engine according to claim 14, wherein each of said second and third duct means are in bypassing relationship to said combustion chamber means.

16. A turbojet engine according to claim 14, wherein said first and second fan means are arranged on separate shafts driven by separate turbines, and wherein said turbines are arranged downstream of said combustion chamber means and are driven by exhaust gas flow from said combustion chamber means.

17. A turbojet engine according to claim 14, wherein said second fan means is operatively drivingly unified with a compressor arranged immediately downstream thereof, and wherein flow into said first duct means passes through said compressor.

18. A turbojet engine according to claim 5, further comprising acoustic silencer means positioned within each said second and third duct means and between the first and second fan means.

19. A turbojet engine according to claim 13, further comprising acoustic silencer means positioned within each of said second and third duct means and between the first and second fan means.

20. A turbojet engine according to claim 17, further comprising acoustic silencer means positioned within each of said second and third duct means and between the first and second fan means.

21. A turbojet engine according to claim 5, wherein fan blades of at least one of said fan means are provided with splitters extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades and supplied to respective ones of said duct means.

22. A turbojet engine according to claim 18, wherein fan blades of at least one of said fan means are provided with splitters extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades and supplied to respective ones of said duct means.

23. A turbojet engine according to claim 13, wherein fan blades of at least one of said fan means are provided with splitters extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades and supplied to respective ones of said duct means.

24. A turbojet engine according to claim 5, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuratioN of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are aerodynamically inactive so as to cause no increase in pressure to the fluid flow therethrough.

25. A turbojet engine according to claim 24, wherein said splitter means is arranged at the junction between respective inlets to said third and second duct means.

26. A turbojet engine according to claim 7, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are aerodynamically inactive so as to cause no increase in pressure to the fluid flow therethrough.

27. A turbojet engine according to claim 13, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are aerodynamically inactive so as to cause no increase in pressure to the fluid flow therethrough.

28. A turbojet engine according to claim 5, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are designed for a very moderate pressure ratio.

29. A turbojet engine according to claim 28, wherein said splitter means is arranged at the junction between respective inlets to said third and second duct means.

30. A turbojet engine according to claim 13, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are designed for a very moderate pressure ratio.

31. A turbojet engine according to claim 5, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are designed to keep the relative velocity through the throat area of the blade ducts at a relatively high mach number in the approximate range of 0.7 to 0.8 so as to dampen the noise.

32. A turbojet engine according to claim 31, wherein said splitter means is arranged at the junction between respective inlets to said third and second duct means.

33. A turbojet engine according to claim 7, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are designed to keep the relative velocity through the throat area of the blade ducts at a relatively high mach number in The approximate range of 0.7 to 0.8 so as to dampen the noise.

34. A turbojet engine according to claim 13, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are designed to keep the relative velocity through the throat area of the blade ducts at a relatively high mach number in the approximate range of 0.7 to 0.8 so as to dampen the noise.

35. A turbojet engine according to claim 18, wherein said first fan means is provided with splitter means extending coaxially with the engine centerline for radially separating differential axial flows caused by the configuration of the blades of said first fan means and supplied to respective ones of said duct means, and wherein those portions of the blades of the first fan means which lie radially inwardly of said splitter means are designed to keep the relative velocity through the throat area of the blade ducts at a relatively high mach number in the approximate range of 0.7 to0.8 so as to dampen the noise.

36. A turbojet engine according to claim 5, wherein said second fan means is formed by outer blade tips of a compressor arranged upstream of said combustion chamber means, and wherein a further compressor is arranged upstream of said combustion chamber means and downstream of said first mentioned compressor, and wherein three separate turbines drive respective ones of said first and second fan means and said further compressor by way of three concentric separately rotatable shafts.

37. A turbojet engine according to claim 5, wherein a multistage compressor is arranged downstream of said second fan means for compressing combustion air supplied to said combustion chamber means, and wherein three separate turbines drive respective ones of said first and second fan means and said multi-stage compressor by way of three concentric separately rotatable drive shafts.