Compressor including an enhanced vaned shroud

• US 20060088412A1
• Filed: 10/27/2004
• Published: 04/27/2006
• Est. Priority Date: 10/27/2004
• Status: Active Grant

• Current Assignee: Honeywell International Inc.
• Original Assignee: Honeywell International Inc.
• Inventors: Barton, Michael T., Nolcheff, Nick A., Palmer, Donald L., Matwey, Mark D., Gunaraj, John A., Mansour, Mahmoud L., Durschmidt, Don F., Slovisky, John A.
• Granted Patent: US 7,189,059 B2
• Time in Patent Office: Days
• US Class Current: 415/206
• CPC Class Codes: F01D 5/26 Antivibration means not restricted to blade form or construction or to blade-to-blade connections or to the use of particular materials
• CPC Class Codes: F04D 29/4213 suction ports
• CPC Class Codes: F04D 29/685 Inducing localised fluid recirculation in the stator-rotor interface

Abstract

A compressor includes an enhanced vaned shroud and is configured such that the flow area ratio is equivalent to that of a conventional, non-vaned shroud. The vaned shroud includes a plurality of airfoils that vary in thickness to obtain desired vibrational mode shapes and natural frequencies. A stiffening ring of limited axial extent is coupled to, and between, the airfoils, and the shroud is manufactured with a section of constant radius.

Claims

1. A compressor, comprising:

a housing;

an impeller rotationally mounted within the housing and having a plurality of impeller blades, a portion of the impeller defining an inducer having an inducer area ratio;

a shroud at least partially surrounding at least a portion of the impeller, the shroud including at least an inner peripheral surface displaced radially outwardly of the impeller; and

a plurality of spaced apart airfoils coupled to, and extending radially inwardly from, the shroud inner peripheral surface,

wherein

the inducer area ratio is substantially equivalent to that of a compressor having a shroud without the plurality of spaced apart airfoils.

2. The compressor of claim 1, wherein:

the impeller blades each include a leading edge and a trailing edge;

each airfoil extends to point of maximum radial extent from the shroud inner peripheral surface; and

the point of maximum radial extent is aligned with the impeller main blade leading edge.

3. The compressor of claim 1, wherein:

each of the airfoils includes a first end and a second end;

each airfoil first end is coupled to the main body inner peripheral surface and has a first thickness;

each airfoil second end has a second thickness; and

the first thickness is greater than the second thickness.

4. The compressor of claim 3, further comprising:

a stiffening ring coupled to each of the airfoils and spaced a predetermined distance from the first end of each of the airfoils.

5. The compressor of claim 4, wherein the stiffening ring is coupled to each of the airfoils between each of the airfoil first and second ends.

6. The compressor of claim 4, wherein the stiffening ring includes:

a faceted leading edge; and

a faceted trailing edge.

7. The compressor of claim 3, wherein the shroud inner surface includes a constant-radius-section of a predetermined axial length, the constant-radius-section having a substantially constant radius along the predetermined axial length,

wherein each of the airfoil first ends is coupled to the constant-radius-section.

8. The compressor of claim 3, wherein each airfoil varies substantially evenly in thickness from the first thickness to the second thickness.

15. The shroud of claim 1, wherein each airfoil varies substantially evenly in thickness from the first thickness to the second thickness.

9. A compressor, comprising:

a housing;

an impeller rotationally mounted within the housing and having a plurality of main blades and a plurality of splitter blades, the main blades and the splitter blades each having at least a leading edge and a trailing edge;

a shroud at least partially surrounding at least a portion of the main blades and the splitter blades, the shroud including at least an inner peripheral surface displaced radially outwardly of each of the main blades and splitter blades; and

a plurality of spaced apart airfoils coupled to, and extending radially inwardly from, the shroud inner peripheral surface,

wherein;

at least a first portion of the shroud inner peripheral surface and each of the splitter blade leading edges define a splitter blade leading edge flow area

each airfoil, at least a second portion of the shroud inner peripheral surface, and each of the main blade leading edges defining a main blade leading edge flow area,

a ratio of the splitter blade leading edge flow area to the main blade leading edge flow area defines an inducer area ratio, and

the inducer area ratio is substantially equivalent to that of a compressor having a shroud without the plurality of spaced apart airfoils.

10. A centrifugal compressor shroud, comprising:

a main body having a first side, a second side, and an inner surface defining a flow passage between the first and second sides;

a plurality of airfoils extending into the main body flow passage, each airfoil having at least a first end and a second end, each airfoil first end coupled to the main body inner surface and having a first thickness, each airfoil second end extending into the main body flow passage and having a second thickness,

wherein the first thickness is greater than the second thickness.

11. The shroud of claim 10, further comprising:

a stiffening ring coupled to each of the airfoils and spaced a predetermined distance from the first end of each of the airfoils.

12. The shroud of claim 11, wherein the stiffening ring is coupled to each of the airfoils between each of the airfoil first and second ends.

13. The shroud of claim 11, wherein the stiffening ring includes:

a faceted leading edge; and

a faceted trailing edge.

14. The shroud of claim 11, wherein the main body inner surface includes a constant-radius-section of a predetermined axial length disposed between the first and second sides, the constant-radius-section having a substantially constant radius along the predetermined axial length,

wherein each of the airfoil first ends is coupled to the main body inner surface airfoil section.

16. A centrifugal compressor shroud, comprising:

a main body having a first side, a second side, and an inner surface defining a flow passage between the first and second sides, the shroud inner surface including a constant-radius-section of a predetermined axial length disposed between the first and second sides, the constant-radius-section having a substantially constant radius along the predetermined axial length; and

a plurality of spaced apart airfoils coupled to, and extending radially inwardly from, the constant-radius-section.

17. The shroud of claim 16, further comprising:

a stiffening ring coupled to each of the airfoils and spaced a predetermined distance from the first end of each of the airfoils.

18. The shroud of claim 17, wherein the stiffening ring is coupled to each of the airfoils between each of the airfoil first and second ends.

19. The shroud of claim 17, wherein the stiffening ring includes:

a faceted leading edge; and

a faceted trailing edge.

20. The shroud of claim 17, wherein each airfoil varies substantially evenly in thickness from the first thickness to the second thickness.

21. A method of designing a vaned shroud for a compressor having an impeller with a plurality of blades, the vaned shroud having a number of airfoils extending from an inner surface thereof, the method comprising the steps of:

determining an inducer area ratio for a conventional, non-vaned shroud compressor;

determining a radial extent for each of the airfoils;

determining the number of airfoils;

determining axial positions for each of the determined number of airfoils radially around the shroud inner surface; and

dimensioning the compressor such that the compressor will have a restored inducer area ratio, the restored inducer area ratio being substantially equivalent to that of the determined inducer area ratio for the conventional, non-vaned shroud compressor.

22. The method of claim 21, wherein the compressor is dimensioned to the restored inducer area ratio by contouring the shroud inner surface at least proximate the splitter blades.

23. The method of claim 21, wherein the impeller blades are coupled to a hub, and wherein the compressor is dimensioned to the restored inducer area ratio by contouring the hub.

24. The method of claim 21, wherein the impeller is dimensioned to the restored inducer area ratio by modifying an angle of the impeller blades.

25. The method of claim 21, wherein the impeller is dimensioned to the restored inducer area ratio by modifying a thickness of the impeller blades.

26. The method of claim 21, further comprising:

determining a shroud inner surface contour that compensates for a reduction in inlet flow area that results from the extension of the airfoils from the shroud inner surface.

27. The method of claim 21, wherein the impeller blades comprise a plurality of main blades and a plurality of splitter blades, the main and impeller blades each having leading edges, the airfoils each include a leading edge, a trailing edge, and a point of maximum radial extent, and wherein the determined axial position is such that:

the point of maximum radial extent is substantially aligned with the main blade leading edges; and

the airfoil trailing edges do not extend beyond the splitter blade leading edges.

28. The method of claim 21, further comprising:

determining a position of a stiffening ring that is coupled to, and between, each of the airfoils.

29. The method of claim 28, wherein the airfoils each include a leading edge, a trailing edge, and a point of maximum radial extent, and the determined stiffening ring position is at least between the shroud inner surface and the point of maximum radial extend.