Gas turbine engine with compressor disk deflectors

Patent metadata

Field	Value
Publication number	US-10260524-B2
Application number	US-201414917981-A
Country	US
Kind code	B2
Filing date	Sep 10, 2014
Priority date	Oct 2, 2013
Publication date	Apr 16, 2019
Grant date	Apr 16, 2019

Abstract

A compressor section for use in a gas turbine engine is comprised of a plurality of compressor stages, with each stage including a disk having an inner periphery defining a bore that faces a shaft. A flow path flows in a generally axial direction between the shaft and the bores of each disk. At least one deflector is positioned between a pair of adjacent bores to direct the air flow radially outwardly into a cavity formed between an associated pair of adjacent disks.

First claim

The invention claimed is:

1. A compressor section for use in a gas turbine engine comprising:
a plurality of compressor stages, each stage including a disk having an inner hub with an inner periphery defining a bore that faces a shaft;
a flow path flowing in an axial direction between the shaft and the bores of the disks;
at least one deflector positioned between a pair of adjacent bores of the bores to direct the air flow radially outwardly into a cavity formed between an associated pair of adjacent disks of the disks, and wherein the at least one deflector comprises a main body having a forward end that is attached to or integrally formed with the hub of an upstream disk of the associated pair of adjacent disks and at least one portion that extends away from the main body such that axial air flow is redirected in a radial direction along a web of the upstream disk, and wherein the at least one portion is axially closer to the upstream disk than a downstream disk of the associated pair of adjacent disks;
wherein the at least one portion includes a first deflector portion extending away from the main body to block a portion of the flow path;
wherein the at least one portion further includes a second deflector portion extending away from the main body to direct blocked flow into the cavity; and
at least one inlet port formed within the main body upstream of the first and second deflector portions to direct the blocked flow into the cavity, and including at least one outlet port formed within the main body downstream of the first and second deflector portions to return flow to the flow path.

2. The compressor section for use in a gas turbine engine as set forth in claim 1 , wherein the first deflector portion extends radially inwardly toward the shaft and wherein the second deflector portion extends radially outwardly away from the shaft.

3. The compressor section for use in a gas turbine engine as set forth in claim 2 , wherein the first deflector portion extends into the flow path to redirect flow into the cavity and the second deflector portion extends into the cavity to direct flow along the web of the upstream disk, and wherein the first and second deflector portions are in radial alignment with each other.

4. The compressor section for use in a gas turbine engine as set forth in claim 1 , wherein the at least one inlet port is adjacent the upstream disk and the at least one outlet port is adjacent the downstream disk.

5. The compressor section for use in a gas turbine engine as set forth in claim 1 , wherein a rearward end of the main body is attached to the downstream disk.

6. The compressor section for use in a gas turbine engine as set forth in claim 1 wherein the at least one deflector is attached to the upstream disk and the downstream disk.

7. The compressor section for use in a gas turbine engine as set forth in claim 1 , wherein the inner hub of the upstream disk includes a rear face that faces a forward face of the inner hub of the downstream disk, and wherein the forward end of the main body is attached to at least a portion of the rear face.

8. The compressor section for use in a gas turbine engine as set forth in claim 7 , wherein the main body includes a rearward end that is attached to at least a portion of the forward face of the inner hub of the downstream disk.

9. A compressor section for use in a gas turbine engine comprising:
a plurality of compressor stages, each stage including a disk having an inner hub with an inner periphery defining a bore that faces a shaft;
a flow path flowing in an axial direction between the shaft and the bores of the disks; and
at least one deflector positioned between a pair of adjacent bores of the bores to direct the air flow radially outwardly into a cavity formed between an associated pair of adjacent disks of the disks, and wherein the at least one deflector comprises a main body having a forward end that is attached to or integrally formed with the hub of an upstream disk of the associated pair of adjacent disks and at least one portion that extends away from the main body such that axial air flow is redirected in a radial direction along a web of the upstream disk, and wherein the at least one portion is axially closer to the upstream disk than a downstream disk of the associated pair of adjacent disks, and wherein the at least one deflector is integrally formed with the hub of the upstream disk.

10. A compressor section for use in a gas turbine engine comprising:
a plurality of compressor stages, each stage including a disk having an inner periphery defining a bore that faces a shaft;
a flow path flowing in an axial direction between the shaft and the bores of the disks; and
at least one deflector positioned between a pair of adjacent bores of the bores to direct the air flow radially outwardly into a cavity formed between an associated pair of adjacent disks, and wherein the at least one deflector is positioned between bores of an upstream disk of the disks and an adjacent downstream disk of the disks, the at least one deflector being axially closer to the upstream disk than the downstream disk, and wherein the at least one deflector includes a portion that extends in a radially outward direction such that axial air flow is redirected in a radial direction along a web of the upstream disk, and wherein the at least one deflector is integrally formed with the upstream disk.

11. A gas turbine engine comprising:
at least one shaft defining an engine axis of rotation;
a compressor section and a turbine section connected to each other by the at least one shaft, wherein the compressor section comprises a plurality of compressor stages, each stage including a disk having an inner hub with an inner periphery defining a bore;
a flow path flowing in an axial direction along the bores of the disks;
at least one deflector positioned between a pair of adjacent bores of the bores to direct the air flow radially outwardly into a cavity formed between an associated pair of adjacent disks of the disks, and wherein the at least one deflector comprises a main body having a forward end that is attached to or integrally formed with the hub of an upstream disk of the associated pair of adjacent disks and at least one portion that extends away from the main body such that axial air flow is redirected in a radial direction along a web of the upstream disk, and wherein the at least one portion is axially closer to the upstream disk than a downstream disk of the associated pair of adjacent disks;
wherein the inner hub of the upstream disk includes a rear face that faces a forward face of the inner hub of the downstream disk, and wherein the forward end of the main body is attached to at least a portion of the rear face; and
wherein the main body extends to a rearward end that is free from attachment to the downstream disk.

12. The gas turbine engine as set forth in claim 11 , wherein the compressor section comprises a high pressure compressor section and the turbine section comprises a high pressure turbine section, and wherein the at least one shaft comprises at least a first shaft connecting the high pressure turbine and high pressure compressor sections and a second shaft connecting a low pressure turbine section to a low pressure compressor section, wherein the first shaft rotates at a faster speed than the second shaft.

13. The gas turbine engine as set forth in claim 12 , including a fan section upstream of the compressor section and driven by one of the first and second shafts.

14. The gas turbine engine as set forth in claim 13 , including a gear drive connecting the fan section to one of

Assignees

• United Technologies Corp

Inventors

• Hiester Paul J

Classifications

• Y02T50/676

  Cross-Sectional Technologies · mapped topic

• F01D5/082

  on the side of the rotor disc · CPC title

• F01D5/085

  cooling fluid circulating inside the rotor · CPC title

• F04D29/582 Primary

  specially adapted for elastic fluid pumps · CPC title

• F02C7/18

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title