IMPELLER BACKFACE SHROUD FOR USE WITH A GAS TURBINE ENGINE

• US 20110299972A1
• Filed: 06/04/2010
• Published: 12/08/2011
• Est. Priority Date: 06/04/2010
• Status: Active Grant

• Current Assignee: Honeywell International Inc.
• Original Assignee: Honeywell International Inc.
• Inventors: Morris, Mark C., MirzaMoghadam, Alexander, Poon, Kin, Howe, Jeff, Tiltman, Alan G., Hosseini, Khosro Molla
• Granted Patent: US 8,801,364 B2
• Time in Patent Office: Days
• US Class Current: 415/1
• CPC Class Codes: F01D 3/025 with a centrally disposed radial stage

Abstract

An impeller or axial stage compressor disk backface shroud for use with a gas turbine engine is disclosed. The backface shroud includes, but is not limited to, a substantially funnel shaped body having a surface. The substantially funnel shaped body is configured to be statically mounted to the gas turbine engine substantially coaxially with the impeller or axial stage compressor disk. The surface and a backface of the impeller or axial stage compressor disk form a cavity that guides an airflow portion to a turbine when the substantially funnel shaped body is mounted coaxially with the impeller or axial stage compressor disk and axially spaced apart therefrom. The airflow portion has a tangential velocity and a recessed groove in the surface of the backface shroud is oriented generally transversely to the tangential velocity to at least partially interfere with the airflow portion, thus affecting static pressure in the cavity.

Claims

1. An impeller or axial stage compressor disk backface shroud for use with a gas turbine engine having an impeller or axial stage compressor disk, the impeller or axial stage compressor disk backface shroud comprising:

a substantially funnel shaped body having a surface, the substantially funnel shaped body configured to be statically mounted to the gas turbine engine in a position that is substantially coaxial with the impeller or axial stage compressor disk, the surface and a backface of the impeller or axial stage compressor disk forming a cavity configured to guide an airflow portion from the impeller to a turbine when the substantially funnel shaped body is mounted to the gas turbine engine coaxially with the impeller or axial stage compressor disk and axially spaced apart therefrom in an aft direction; and

a recessed groove defined in the surface,

wherein the airflow portion has a tangential velocity and wherein the recessed groove is oriented generally transversely to the tangential velocity of the airflow portion and configured to at least partially interfere with the airflow portion, whereby a static pressure in the cavity is affected.

2. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove has a cross section having a substantially square aspect ratio.

3. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove has a cross section having a rectangular low aspect ratio.

4. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove has a cross section having a rectangular high aspect ratio.

5. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove has a cross section having a curved low aspect ratio.

6. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove has a cross section having a curved high aspect ratio.

7. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove has a cross section having a curved forward tapered aspect ratio.

8. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove has a cross section having a curved, rearward tapered aspect ratio.

9. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove extends radially through the surface.

10. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove extends in a forward sweep through the surface.

11. The impeller or axial stage compressor disk backface shroud of claim 1, wherein the recessed groove extends in a rearward sweep through the surface.

12. A gas turbine engine comprising:

a shaft;

an impeller or axial stage compressor disk affixed to the shaft;

a turbine affixed to the shaft at a location aft of the impeller; and

an impeller or axial stage compressor disk backface shroud comprising;

a substantially funnel shaped body having a surface, the substantially funnel shaped body being statically mounted to the gas turbine engine in a position that is substantially coaxial with the impeller or axial stage compressor disk and axially spaced apart therefrom in an aft direction such that the surface and a backface of the impeller or axial stage compressor disk form a cavity, the cavity being configured to guide an airflow portion from the impeller or axial stage compressor disk to the turbine, the airflow portion having a tangential velocity; and

a recessed groove defined in the surface, the recessed groove oriented generally transversely to the tangential velocity of the airflow portion and configured to at least partially interfere with the airflow portion, whereby a static pressure in the cavity is affected.

13. The gas turbine engine of claim 12, wherein the recessed groove has a cross section having a substantially square aspect ratio.

14. The gas turbine engine of claim 12, wherein the recessed groove has a cross section having a rectangular low aspect ratio.

15. The impeller gas turbine engine of claim 12, wherein the recessed groove has a cross section having a rectangular high aspect ratio.

16. The gas turbine engine of claim 12, wherein the recessed groove has a cross section having one of a curved low aspect ratio and a curved high aspect ratio.

17. The gas turbine engine of claim 12, wherein the recessed groove has a cross section having one of a curved, forward tapered aspect ratio and a curved, aft tapered aspect ratio.

18. The gas turbine engine of claim 12, wherein the recessed groove extends in a forward sweep through the surface.

19. The impeller or axial stage compressor disk backface shroud of claim 12, wherein the recessed groove extends in a rearward sweep through the surface.

20. A method for compensating for an undesirable amount of spool thrust in a gas turbine engine having a shaft, an impeller or axial stage compressor disk affixed to the shaft, a turbine affixed to the shaft at a location aft of the impeller or axial stage compressor disk, and an impeller or axial stage compressor disk backface shroud statically mounted to the gas turbine engine in a position that is substantially coaxial with the impeller and aft thereof such that a surface of the impeller or axial stage compressor disk backface shroud and a backface of the impeller form a cavity configured to guide an airflow portion from the impeller or axial stage compressor disk to the turbine, the airflow portion having a tangential velocity, the method comprising the steps of:

A. determining a target static pressure;

B. performing a computational fluid dynamic analysis using a processor to determine a static pressure in the cavity that would result from defining a recessed groove in the surface of the impeller or axial stage compressor disk backface shroud, the recessed groove having a predetermined configuration;

C. changing the predetermined configuration of the recessed groove if the static pressure in the cavity differs substantially from the target static pressure;

D. repeating steps B and C until a configuration of the recessed groove that yields a static pressure in the cavity that does not differ substantially from the target static pressure is determined;

E. manufacturing a second impeller or axial stage compressor disk backface shroud including the recessed groove having the configuration determined at step D; and

F. assembling the second impeller or axial stage compressor disk backface shroud to the gas turbine engine.