Gas turbine engine diffuser system for a high pressure (HP) compressor

Patent metadata

Field	Value
Publication number	US-9726032-B2
Application number	US-201314137987-A
Country	US
Kind code	B2
Filing date	Dec 20, 2013
Priority date	Mar 8, 2013
Publication date	Aug 8, 2017
Grant date	Aug 8, 2017

Abstract

A gas turbine engine includes a compressor assembly that is rotationally coupled to a shaft. The engine includes a radial diffuser assembly coupled to a shroud of the compressor assembly and positioned to receive compressed air from a centrifugal impeller. The radial diffuser assembly includes a first arcuate wall and a second wall, and a near axial diffuser coupled to the radial diffuser assembly and positioned to receive the compressed air from the radial diffuser assembly. The gas turbine includes a gas seal coupled between the second wall and a wall of the near axial diffuser, the gas seal configured to prevent the compressed air from passing through the seal while allowing relative motion between the radial diffuser assembly and the near axial diffuser.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a compressor assembly rotationally coupled to a shaft, the compressor assembly having a centrifugal impeller;
a radial diffuser assembly coupled to a shroud of the compressor assembly and positioned to receive compressed air from the centrifugal impeller, the radial diffuser assembly having a first arcuate wall and a second wall;
another diffuser coupled to the radial diffuser assembly and positioned to receive the compressed air from the radial diffuser assembly;
a gas seal coupled between the second wall and a wall of the another diffuser, the gas seal configured to prevent the compressed air from passing through the seal while allowing relative motion between the radial diffuser assembly and the another diffuser;
a bleed slot positioned within the shroud proximate the centrifugal impeller; and
a bleed plenum formed in part by a wall of the shroud and by a non-load bearing material that is attached to the radial diffuser assembly and to the shroud via a pair of flanges that are in direct contact with one another, wherein bleed air that passes through the bleed slot passes directly from the centrifugal impeller, through the bleed slot, and into the bleed plenum.

2. The gas turbine engine as claimed in claim 1 , wherein the first arcuate wall includes a variable thickness about an arcuate portion thereof.

3. The gas turbine engine as claimed in claim 2 , the gas turbine engine further comprising an outer structure that is bolted at a bolted joint to the radial diffuser assembly wherein the variable thickness includes a first location with a first thickness and a second location with a second thickness, wherein the first thickness is greater than the second thickness.

4. The gas turbine engine as claimed in claim 1 , wherein the gas seal includes a w-shaped seal positioned between the radial diffuser assembly and the another diffuser.

5. The gas turbine engine as claimed in claim 1 , wherein the gas seal comprises:
a first seal flange attached to the wall of the another diffuser;
a second seal flange attached to the second wall of the radial diffuser assembly; and
a w-seal positioned between the first seal flange and the second seal flange.

6. The gas turbine engine of claim 1 , wherein the bleed slot positioned within the shroud and proximate the centrifugal impeller is positioned within the shroud along an arcuate inner surface along vanes of the centrifugal impeller.

7. The gas turbine engine of claim 1 , wherein the pair of flanges extend approximately orthogonal with respect to a centerline of the gas turbine engine.

8. A method of manufacturing a gas turbine engine comprising:
coupling a radial diffuser assembly to a shroud of a compressor assembly, wherein the radial diffuser assembly is positioned to receive compressed air from a centrifugal impeller, and wherein the radial diffuser assembly includes a first arcuate wall and a second wall;
coupling a second diffuser to the radial diffuser assembly to receive the compressed air from the radial diffuser assembly; and
coupling a gas seal between the second wall and a wall of the second diffuser, the gas seal configured to prevent the compressed air from passing through the seal while allowing relative motion between the radial diffuser assembly and the second diffuser;
forming a bleed slot within the shroud proximate the centrifugal impeller;
forming a bleed plenum with the steps of:
attaching a non-load bearing material to the radial diffuser assembly via a pair of flanges that are in direct contact with one another; and
attaching the non-load bearing material to the shroud;

wherein bleed air that passes through the bleed slot passes directly from the centrifugal impeller, through the bleed slot, and into the bleed plenum.

9. The method as claimed in claim 8 , further comprising forming the first arcuate wall of the gas turbine having a variable thickness about an arcuate portion of the first arcuate wall.

10. The method as claimed in claim 9 , further comprising:
forming an outer structure; and
attaching a radial diffuser assembly to the outer structure;
wherein the variable thickness includes a first location with a first thickness and a second location with a second thickness, wherein the first thickness is greater than the second thickness, and wherein the first location is closer to a bolted joint than the second location.

11. The method as claimed in claim 8 , the method further comprising:
attaching a first seal flange to the wall of the second diffuser;
attaching a second seal flange to the second wall of the radial diffuser assembly; and
positioning a w-seal between the first and second seal flanges to form the gas seal coupled between the second wall and the wall of the second diffuser.

12. The method of claim 8 , wherein the bleed slot positioned within the shroud and proximate the centrifugal impeller is positioned within the shroud along an arcuate inner surface along vanes of the centrifugal impeller.

13. The method of claim 8 , wherein attaching the non-load bearing material to the radial diffuser assembly via the pair of flanges, further comprises attaching such that the pair of flanges extend approximately orthogonal with respect to a centerline of the gas turbine engine.

14. A gas turbine engine comprising:
a radial diffuser coupled to a shroud of a compressor assembly, the compressor assembly rotationally coupled to a shaft, the compressor assembly having a centrifugal impeller, wherein the radial diffuser is positioned to receive compressed air from the centrifugal impeller;
a near-axial diffuser coupled to the radial diffuser and positioned to receive compressed air from the radial diffuser;
a gas seal coupled between a wall of the radial diffuser and a wall of the near-axial diffuser;
a bleed slot positioned within the shroud proximate the centrifugal impeller; and
a bleed plenum formed in part by a wall of the shroud and by a non-load bearing material that is attached to the radial diffuser assembly and to the shroud via a pair of flanges that are in direct contact with one another, wherein bleed air that passes through the bleed slot passes directly from the centrifugal impeller, through the bleed slot, and into the bleed plenum.

15. The gas turbine engine as claimed in claim 14 , wherein the gas seal is configured to prevent compressed air from passing through the seal while allowing relative motion between the radial diffuser and the near-axial diffuser.

16. The gas turbine engine as claimed in claim 14 , wherein the radial diffuser includes a first arcuate wall and a second wall, and wherein the first arcuate wall includes a variable thickness about an arcuate portion thereof.

17. The gas turbine engine as claimed in claim 16 , the gas turbine engine further comprising an outer structure that is bolted at a joint to the radial diffuser, wherein the variable thickness includes a first location with a first thickness and a second location with a second thickness, wherein the first thickness is greater than the second thickness;
wherein the near-axial diffuser further comprises a vaned assembly having one or more service struts configured to pass a fluid therethrough.

18. The gas turbine as claimed in claim 14 , wherein the gas seal comprises:
a first seal flange attached to the wall of the near-axial diffuser;
a second seal flange attached to the second wall of the radial diffuser; and
a w-seal positioned between the first flange and the second flange.

19. The

Assignees

• Rolls Royce Nam Tech Inc
• Rolls-Royce American Tech Inc
• Rolls Royce Plc

Inventors

• Ress Jr Robert A
• Swift Andrew

Classifications

• F02C7/28

  Arrangement of seals · CPC title

• F04D29/083

  especially adapted for elastic fluid pumps · CPC title

• F01D11/005 Primary

  Sealing means between non relatively rotating elements · CPC title

• F04D29/444

  Bladed diffusers · CPC title

• F01D9/065

  Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids (see also F01D25/16, F01D25/24 and F01D25/26) · CPC title