Aft mounted pusher fan for gas turbine engine

Patent metadata

Field	Value
Publication number	US-12510030-B2
Application number	US-202418646941-A
Country	US
Kind code	B2
Filing date	Apr 26, 2024
Priority date	Dec 23, 2022
Publication date	Dec 30, 2025
Grant date	Dec 30, 2025

Abstract

A fan drive turbine drives a drive shaft. A compressor section and a turbine section rotate about a rotational axis, with the turbine section positioned in a downstream direction. A fan has fan blades and a fan hub positioned radially within a fan case, with the fan case defining a bypass duct. A fan inlet guide vane case is upstream of the fan blades and a fan exit guide vane case is mounted downstream. The drive shaft is mounted on at least one input drive shaft bearing inwardly of the fan inlet guide vane case. The drive shaft drives a flexible coupling to in turn drive a gear reduction. The gear reduction drives a fan drive shaft to in turn drive the fan hub. At least two fan drive shaft bearings support the fan drive shaft within the fan exit guide vane case. An aircraft is also disclosed.

First claim

What is claimed is:

1 . A gas turbine engine comprising:
a compressor section, a combustor and a turbine section defining a core engine, a fan drive turbine driving a drive shaft, said compressor section and said turbine section rotating about a rotational axis, with said turbine section positioned in a downstream direction relative to said compressor section; a fan having fan blades and a fan hub positioned radially within a fan case, with the fan case defining a bypass duct; a fan inlet guide vane case upstream of the fan blades and a fan exit guide vane case mounted downstream of the fan blades; said drive shaft being mounted on at least one input drive shaft bearing inwardly of the fan inlet guide vane case, said drive shaft driving a flexible coupling to in turn drive a gear reduction, the gear reduction driving a fan drive shaft to in turn drive the fan hub; at least two fan drive shaft bearings supporting the fan drive shaft and mounted within the fan exit guide vane case; wherein there is an upstream fairing positioned forward of the fan inlet guide vane case; and wherein the upstream fairing is fixed to the fan inlet guide vane case and has a curved outer surface to guide air outwardly of the upstream fairing and approaching the fan inlet guide vane case.

2 . The gas turbine engine as set forth in claim 1 , wherein there is a static inner fairing intermediate the fan inlet guide vane case and the fan hub.

3 . The gas turbine engine as set forth in claim 1 , wherein said gear reduction is mounted aft of the fan drive shaft.

4 . The gas turbine engine as set forth in claim 3 , wherein drive shaft oil seals are positioned within the fan inlet guide vane case on each axial side of the at least one input drive shaft bearing.

5 . The gas turbine engine as set forth in claim 4 , wherein a fan shaft oil seal is positioned between the fan drive shaft and the flexible coupling.

6 . An aircraft comprising:
an aircraft body mounting at least two gas turbine engines; each of said gas turbine engines having a compressor section, a combustor and a turbine section defining a core engine, and a fan drive turbine driving a drive shaft, said compressor section and said turbine section rotating about a rotational axis, with said turbine section positioned in a downstream direction relative to said compressor section; a fan having fan blades and a fan hub positioned radially within a fan case, with the fan case defining a bypass duct; a fan inlet guide vane case upstream of the fan blades and a fan exit guide vane case mounted downstream of the fan blades; said drive shaft being mounted on at least one input drive shaft bearing inwardly of the fan inlet guide vane case, said drive shaft driving a flexible coupling to in turn drive a gear reduction, the gear reduction driving a fan drive shaft to in turn drive the fan hub; at least two fan drive shaft bearings supporting the fan drive shaft and mounted within the fan exit guide vane case; and wherein an upstream fairing is fixed to the fan inlet guide vane case and has a curved outer surface to guide air outwardly of the upstream fairing and approaching the fan inlet guide vane case.

7 . The aircraft as set forth in claim 6 , wherein said gear reduction is mounted aft of the fan drive shaft.

8 . The aircraft as set forth in claim 7 , wherein drive shaft oil seals are positioned within the fan inlet guide vane case on each axial side of the at least one input drive shaft bearing, and a fan shaft oil seal is positioned between the fan drive shaft and the flexible coupling.

9 . The aircraft as set forth in claim 6 , wherein the fan drive shaft is mounted radially outwardly of the flexible coupling.

10 . The aircraft as set forth in claim 6 , wherein the aircraft has a pair of wings and one of said gas turbine engines is mounted vertically above each of said wings.

11 . The aircraft as set forth in claim 6 , wherein the aircraft has a pair of wings with struts extending from a central fuselage to an underside of the wings, and one of said gas turbine engines is mounted beneath each of the wings.

12 . The aircraft as set forth in claim 6 , wherein the aircraft has a central fuselage with wings merging outwardly of the central fuselage and there being a plurality of said gas turbine engines mounted in each of two lateral sides of said central fuselage.

13 . The aircraft as set forth in claim 12 , wherein said fan drive turbines are positioned forward of the fans in the gas turbine engines.

14 . The aircraft as set forth in claim 6 , wherein said fan drive turbine is spaced axially from said core engine and separate from said core engine.

Assignees

• Rtx Corp

Inventors

• Sobanski Jon Erik
• Breault Andrew E

Classifications

• F05D2240/62

  Flexible · CPC title

• F05D2220/323

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

• F02K3/04

  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

• F02C3/10

  with another turbine driving an output shaft but not driving the compressor · CPC title