Reverse flow single spool core gas turbine engine

Patent metadata

Field	Value
Publication number	US-9989011-B2
Application number	US-201514639164-A
Country	US
Kind code	B2
Filing date	Mar 5, 2015
Priority date	Apr 15, 2014
Publication date	Jun 5, 2018
Grant date	Jun 5, 2018

Abstract

A bypass housing receives a fan and defines a front end. An airflow path delivers air into an inlet duct over a limited circumferential extent of the bypass housing. An airflow path passes across a low pressure compressor rotor. An airflow path passes through a core engine, which includes a high pressure compressor rotor, a combustor, and a high pressure turbine rotor. Products of combustion downstream of the high pressure turbine rotor pass into an intermediate duct and then across a low pressure turbine rotor. The low pressure turbine rotor is positioned closer to the front end of the engine than is the high pressure turbine rotor. The low pressure turbine rotor is positioned axially intermediate the low pressure compressor rotor and the fan. The low pressure turbine rotor drives both the fan and the low pressure turbine rotor. An aircraft is also disclosed.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a bypass housing receiving and circumscribing blades of a fan, and defining an axially front end, an airflow path downstream of said fan delivering air into an inlet duct of a low pressure compressor rotor over a limited circumferential extent of the bypass housing, an airflow path downstream of the inlet duct passing across the low pressure compressor rotor;
an airflow path downstream of said low pressure compressor rotor passing through a core engine, said core engine including a high pressure compressor rotor, a combustor, and a high pressure turbine rotor;
products of combustion downstream of the high pressure turbine rotor passing into an intermediate duct and then across a low pressure turbine rotor, with said low pressure turbine rotor being positioned closer to said front end of said engine than is said high pressure turbine rotor; and
said low pressure turbine rotor positioned axially intermediate said low pressure compressor rotor and said fan, and said low pressure turbine rotor driving both said fan and said low pressure compressor rotor, and wherein the inlet duct of the low pressure compressor rotor extends in a radial direction to at least a portion of an inner surface of the bypass housing at a location radially outward of the fan blades.

2. The gas turbine engine as set forth in claim 1 , wherein said low pressure turbine rotor driving a shaft which, in turn, drives said fan through a gear reduction.

3. The gas turbine engine as set forth in claim 1 , wherein said core engine rotates about a first axis, and said low pressure turbine rotor rotates about a second axis, with first and second axes being non-parallel.

4. The gas turbine engine as set forth in claim 1 , wherein said core engine rotates about a first axis and said low pressure turbine rotor rotates about a second axis with said first and second axes being parallel, but spaced from each other.

5. The gas turbine engine as set forth in claim 1 , wherein said core engine rotating about a first axis and said low pressure turbine rotor rotating about a second axis with said first and second axes being co-axial.

6. The gas turbine engine as set forth in claim 5 , wherein an airflow path downstream of said low pressure compressor rotor passing axially further into said engine and into said high pressure compressor rotor, and products of combustion downstream of the high pressure turbine rotor being turned back in a direction toward said axial front end of said engine before reaching said low pressure turbine rotor.

7. The gas turbine engine as set forth in claim 1 , wherein said low pressure turbine driving a shaft, with said shaft extending through said intermediate duct, and through said inlet duct to drive said low pressure compressor from said low pressure turbine rotor.

8. The gas turbine engine as set forth in claim 1 , wherein there being an exhaust duct downstream of said low pressure turbine rotor, with said exhaust duct delivering products of combustion to mix with bypass air from the bypass duct.

9. The gas turbine engine as set forth in claim 1 , wherein an airflow path downstream of said low pressure compressor turns into a direction heading axially back toward said front end of said engine before reaching said high pressure compressor rotor.

10. The gas turbine engine as set forth in claim 1 , wherein an airflow path downstream of said low pressure compressor rotor passing axially further into said engine and into said high pressure compressor rotor, and products of combustion downstream of the high pressure turbine rotor being turned back in a direction toward said axial front end of said engine before reaching said low pressure turbine rotor.

11. An aircraft comprising:
a pair of gas turbine engines in a tail area; and
the gas turbine engines having a bypass housing receiving and circumscribing blades of a fan, and defining an axially front end, an airflow path downstream of said fan delivering air into an inlet duct of a low pressure compressor rotor over a limited circumferential extent of the bypass housing, an airflow path downstream of the inlet duct passing across the low pressure compressor rotor, an airflow path downstream of said low pressure compressor rotor passing through a core engine, said core engine including a high pressure compressor rotor, a combustor, and a high pressure turbine rotor, products of combustion downstream of the high pressure turbine rotor passing into an intermediate duct and then across a low pressure turbine rotor, with said low pressure turbine rotor being positioned closer to said front end of said engine than is said high pressure turbine rotor, said low pressure turbine rotor positioned axially intermediate said low pressure compressor rotor and said fan, and said low pressure turbine rotor driving both said fan and said low pressure compressor rotor, and wherein the inlet duct of the low pressure compressor rotor extends in a radial direction to at least a portion of an inner surface of the bypass housing at a location radially outward of the fan blades.

12. The aircraft as set forth in claim 11 , wherein said low pressure turbine rotor driving a shaft which, in turn, drives said fan through a gear reduction.

13. The aircraft as set forth in claim 11 , wherein said core engine rotates about a first axis, and said low pressure turbine rotor rotates about a second axis, with said first and second axes being non-parallel.

14. The aircraft as set forth in claim 11 , wherein said core engine rotates about a first axis and said low pressure turbine rotor rotates about a second axis with said first and second axes being parallel, but spaced from each other.

15. The aircraft as set forth in claim 11 , wherein said core engine rotating about a first axis and said low pressure turbine rotor rotating about a second axis with said first and second axes being co-axial.

16. The aircraft as set forth in claim 15 , wherein an airflow path downstream of said low pressure compressor rotor passing axially further into said engine and into said high pressure compressor rotor, and products of combustion downstream of the high pressure turbine rotor being turned back in a direction toward said axial front end of said engine before reaching said low pressure turbine rotor.

17. The aircraft as set forth in claim 11 , wherein said low pressure turbine driving a shaft, with said shaft extending through said intermediate duct, and through said inlet duct to drive said low pressure compressor from said low pressure turbine rotor.

18. The aircraft as set forth in claim 11 , wherein there being an exhaust duct downstream of said low pressure turbine rotor, with said exhaust duct delivering products of combustion to mix with bypass air from the bypass duct.

19. The aircraft as set forth in claim 11 , wherein an airflow path downstream of said low pressure compressor turns into a direction heading axially back toward said front end of said engine before reaching said high pressure compressor rotor.

20. The aircraft as set forth in claim 11 , wherein an airflow path downstream of said low pressure compressor rotor passing axially further into said engine and into said high pressure compressor rotor, and products of combustion downstream of the high pressure turbine rotor being turned back in a direction toward said axial front end of said engine before reaching said low pressure turbine rotor.

21. The aircraft as set forth in claim 11 , wherein there is a

Assignees

• United Technologies Corp

Inventors

• Suciu Gabriel L
• Chandler Jesse M
• Kupratis Daniel Bernard

Classifications

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F02K3/04 Primary

  the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type · CPC title

• F02C3/06

  the compressor comprising only axial stages (F02C3/10 takes precedence) · CPC title

• F02K3/025

  the by-pass flow being at least partly used to create an independent thrust component · CPC title

• F05D2250/314

  the axes being inclined in relation to each other · CPC title