Reverse core gear turbofan

• US 9,239,004 B2
• Filed: 03/27/2012
• Issued: 01/19/2016
• Est. Priority Date: 03/27/2012
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Kupratis, Daniel Bernard, Moon, Francis R.
• Primary Examiner(s): Goyal, Arun
• Application Number:
  US13/431,048
• Publication Number:
  US 20130255224A1
• Time in Patent Office: 1,393 Days
• US Class Current: 1/1
• CPC Class Codes: F01D 17/141 by means of shiftable members or valves obturating part of the flow path
• CPC Class Codes: F02C 3/05 the compressor and the turbine being of the radial flow type
• CPC Class Codes: F02C 3/10 with another turbine driving an output shaft but not driving the compressor
• CPC Class Codes: F02C 3/107 with two or more rotors connected by power transmission
• CPC Class Codes: F02C 3/13 having variable working fluid interconnections between turbines or compressors or stages of different rotors controlling flow ratio between different flows of multi-flow jet-propulsion plant, e.g. ducted fan F02K3/075
• CPC Class Codes: F02C 3/145 the combustion chamber being in the reverse flow-type
• CPC Class Codes: F02C 6/02 Plural gas-turbine plants having a common power output
• CPC Class Codes: F02C 7/36 Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user F02C3/107 - F02C3/13 and F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H
• 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: F02K 3/06 with front fan
• CPC Class Codes: F02K 3/065 with front and aft fans
• CPC Class Codes: F02K 3/075 controlling flow ratio between flows
• CPC Class Codes: F05D 2220/32 in gas turbines
• CPC Class Codes: F05D 2220/323 for aircraft propulsion, e.g. jet engines
• CPC Class Codes: F05D 2250/36 in inner-outer relationship, e.g. shaft-bearing arrangements
• CPC Class Codes: F05D 2260/4023 through a friction clutch
• CPC Class Codes: F05D 2260/4031 as in toothed gearing
• CPC Class Codes: F05D 2270/053 Explicitly mentioned power

Abstract

A gas turbine engine has a fan at an axially outer location. The fan rotates about an axis of rotation. The fan delivers air into an outer bypass duct, and across a booster fan positioned radially inwardly of the outer bypass duct. The booster fan delivers air into a radially middle duct, and across a cold turbine into a radially inner core duct being directed into a compressor. From the compressor, air flows axially in a direction back toward the fan through a combustor section, and across an exhaust of the turbine section as directed into the middle duct. A gear reduction drives the fan from a fan drive turbine section. A method is also disclosed.

Claims

1. A gas turbine engine comprising:

a fan at an axially forward location, said fan rotating about an axis of rotation;

said fan delivering air into an outer bypass duct, and across a booster fan positioned radially inwardly of said outer bypass duct, said booster fan delivering air into a radially middle duct, and said booster fan delivering air across a cold turbine section into a radially inner core duct wherein said middle duct is located radially inside of said outer bypass duct and radially outside of said inner core duct;

air from said cold turbine section in said inner core duct being directed into a compressor section, and then flowing axially in a direction back toward said fan via a duct located radially inside of said inner core duct through a combustor section, and across a core turbine section, and then being directed into said middle duct; and

a gear reduction for driving said fan from a fan drive turbine section.

2. The gas turbine engine as set forth in claim 1, wherein a shaft downstream of said gear reduction relative to said fan drive turbine section also drives said booster fan.

3. The gas turbine engine as set forth in claim 2, wherein said shaft downstream of said gear reduction is also connected to rotate with said cold turbine section.

4. The gas turbine engine as set forth in claim 3, wherein said fan booster, and said cold turbine section rotate with a clutched shaft separate from a fan shaft driving said fan, and a clutch selectively connecting said clutched shaft to said fan shaft such that said fan shaft can selectively drive said clutched shaft.

5. The gas turbine engine as set forth in claim 1, wherein there are at least a plurality of core turbine sections, with one of said plurality of said core turbine sections driving said fan through said gear reduction and a second of said plurality of said core turbine sections driving said fan booster.

6. The gas turbine engine as set forth in claim 1, wherein said compressor section includes at least a first compressor section and a second compressor section downstream of said first compressor section, and said core turbine section includes at least a first core turbine section and a second core turbine section, with said first core turbine section driving said second compressor section and said second core turbine section driving said first compressor section, with said second turbine section and said first compressor section operating at a slower speed and at lower pressures than said first turbine section and said second compressor section.

7. The gas turbine engine as set forth in claim 1, wherein a first cold turbine section is positioned adjacent said booster fan, and a second cold turbine section is positioned downstream of the first cold turbine section in a path of air flowing through said inner core duct, and upstream of said compressor section.

8. The gas turbine engine as set forth in claim 7, wherein one of said first cold turbine section and said second cold turbine section is provided with a flow diverter that allows bypass of air around a rotor associated with said one of said first cold turbine section and said second cold turbine section.

9. The gas turbine engine as set forth in claim 8, wherein a radially outer extent of blades associated with said one of said first cold turbine section and said second cold turbine section is spaced inwardly of a radially outer position for said flow diverter to allow bypass of air radially outwardly of said radially outermost extent of the blades of said one of said first cold turbine section and said second cold turbine section.

10. The gas turbine engine as set forth in claim 9, wherein there are a pair of flow diverters, with said pair of flow diverters being movable between a position allowing bulk of the air delivered to the compressor section to bypass the rotor by passing radially outwardly of the radially outermost extent of the blades, and said pair of flow diverters being movable to an alternative position wherein a great bulk of the air delivered across said one of said first cold turbine section and said second cold turbine section passes radially inwardly of the radially outermost extent of the blades.

11. The gas turbine engine as set forth in claim 8, wherein said flow diverter is associated with said first cold turbine section.

12. The gas turbine engine as set forth in claim 8, wherein said flow diverter is associated with said second cold turbine section.

13. The gas turbine engine as set forth in claim 1, wherein said cold turbine section is provided with a flow diverter that allows bypass of air around a rotor associated with said cold turbine section.

14. The gas turbine engine as set forth in claim 13, wherein a radially outer extent of blades associated with said cold turbine section is spaced inwardly of a radially outer position for said flow diverter to allow bypass of air radially outwardly of said radially outermost extent of the blades in said cold turbine section.

15. The gas turbine engine as set forth in claim 14, wherein there are a pair of flow diverters, with said pair of flow diverters being movable to a position allowing bulk of the air delivered to the compressor section to bypass the rotor by passing radially outwardly of the radially outermost extent of the blades, and said pair of flow diverters being movable to an alternative position with a great bulk of the air delivered across said cold turbine section passes radially inwardly of the radially outermost extent of the blades.

16. The gas turbine engine as set forth in claim 15, wherein said cold turbine section associated with said pair of flow diverters is positioned adjacent to said booster fan.

17. The gas turbine engine as set forth in claim 15, wherein said cold turbine section associated with said pair of flow diverters is positioned at a location adjacent to said compressor section on an axial side of said compressor section spaced away from said fan.

18. The gas turbine engine as set forth in claim 13, wherein said cold turbine section associated with said flow diverter is positioned adjacent to said booster fan.

19. The gas turbine engine as set forth in claim 13, wherein said cold turbine section associated with said flow diverter is positioned at a location adjacent to said compressor section on an axial side of said compressor section spaced away from said fan.

20. The gas turbine engine as set forth in claim 13, wherein said cold turbine section associated with said flow diverter is positioned adjacent to said booster fan.