WO2020068109A1

WO2020068109A1 - Modular cooling arrangement for cooling airfoil components in a gas turbine engine

Info

Publication number: WO2020068109A1
Application number: PCT/US2018/053393
Authority: WO
Inventors: Zachary D. Dyer; Moritz Fischle; Jan H. Marsh; Wentao Fu; Allister William James; Andrew Miller; David J. Mitchell
Original assignee: Siemens Aktiengesellschaft; Siemens Energy, Inc.
Priority date: 2018-09-28
Filing date: 2018-09-28
Publication date: 2020-04-02

Abstract

A modular cooling arrangement for cooling an airfoil component (10) in a combustion turbine engine is provided. A body (16) of the airfoil component, and a cooling manifold (18) or a flow sleeve (50, 60) may be disposed in a cavity in the body of the airfoil component. By way of example, the body of the airfoil component may be a ceramic matrix composite body, a MAX phase ternary ceramic body, or a metallic body. Disclosed embodiments are effective to reliably and cost-effectively reduce a level of internal pressurization in the component while providing appropriate cooling to portions of the component subject to a hot flow of gases and accommodating thermal growth differences that may develop between components of the modular cooling arrangement and anchoring structures to which the modular cooling arrangement may be affixed.

Description

MODULAR COOLING ARRANGEMENT FOR COOLING

AIRFOIL COMPONENTS IN A GAS TURBINE ENGINE

BACKGROUND

[0003] The present invention relates generally to the field of turbomachinery, and, more particularly, to a modular cooling arrangement for cooling airfoil components in a turbomachine, such as a gas turbine engine.

2. Description of the Related Art

[0005] As one skilled in the art of gas turbine engine technology would appreciate, the efficiency of the engine may be enhanced by operating a turbine stage of the gas turbine engine at a higher temperature. Another feature that contributes to the efficiency of the engine is the ability to cool components of the gas turbine engine with a lesser amount of cooling air.

[0006] One physical constraint for operating at higher temperatures is the impact of such temperatures on the structural integrity of turbomachinery components disposed in such a high temperature environment. People skilled in the art have attempted to deal with the structural integrity issue by utilizing various cooling techniques, such as may involve internal cooling of such components and/or selecting materials with better resistance properties to the high temperatures involved.

[0007] Challenges associated with internal cooling may be tw_'ofold. Firstly, the cooling air that is utilized for this cooling is obtained from a compressor stage of the gas turbine engine, where energy has been spent to pressurize this air, and consequently air winch is diverted to the cooling process in essence is a deficit in engine efficiency. Secondly, cooling through cooling passages and holes that may be configured in turbomachinery components, such as in the body of an airfoil, can adversely affect the structural integrity of such components.

[0008] While there are materials, such as composite matrix composite (CMC)

materials and other materials, that may he able to withstand the higher temperatures, one issue is how to reliably and cost-effectively arrange these materials to be appropriately cooled while reducing a level of internal pressurization to which such materials may be exposed to during operation. At least in view of the foregoing considerations, further improvements are desired in connection with the cooling of turbomachinery components, such as airfoil components.

DESCRIPTION OF THE DRAWINGS

[0010] FIG. 1 is an isometric view illustrating one non-limiting embodiment of a disclosed modular cooling arrangement, as may be embodied in an airfoil component, such as a vane disposed in a gas turbine engine between anchoring structures, such as a radially-outer shroud and a radially-inner shroud, fragmentarily illustrated in the figure.

FIG. 2 is a cross-sectional view of the disclosed modular cooling arrangement shown in FIG. 1 illustrating one non-limiting embodiment of a cooling manifold, such as may be disposed in a cavity in the body of the airfoil component.

[0012] FIG. 3 is a cross-sectional view of the disclosed modular cooling arrangement shown in FIG. I illustrating another non-limiting embodiment of the cooling manifold within the body of the airfoil component. [0013] FIG. 4 is an isometric view of the disclosed modular cooling arrangement shown in FIG. 1 illustrating respective profiles of the body of the airfoil component and the cooling manifold disposed within the body of the airfoil component.

[0014] FIG. 5 is an isometric view illustrating non-limiting structural details of the anchoring structure where the disclosed modular cooling arrangement shown in FIG. 1 may be affixed.

[0015] FIG. 6-7 are respective isometric views of another non-limiting embodiment of a disclosed modular cooling arrangement involving a flow sleeve, as illustrated in FIG. 8, disposed in the body of the airfoil component.

[0016] FIG. 9 illustrates an alternative embodiment of the flow sleeve.

[0017] FIG. 10 is a cross-sectional view of a disclosed modular cooling arrangement involving a multi-piece cooling manifold arrangement including a cooling manifold piece configured to provide cooling to a trailing edge region of the airfoil component.

[0018] FIG. 1 1 is a cross-sectional view/ illustrating non-limiting structural details of the cooling manifold piece configured to provide cooling to the trailing edge region of the airfoil component.

[0019] DETAILED DESCRIPTION

[0020] Disclosed embodiments are directed to a modular cooling arrangement for cooling airfoil components, such as blades or vanes, In a combustion turbine engine (e.g., a gas turbine engine). Without limitation, disclosed embodiments are effective to reliably and cost-effectively reduce a level of internal pressurization in the component while providing appropriate cooling to portions of the component subject to a hot flow of gases. Disclosed embodiments are designed to accommodate thermal growth differences that may develop between components of the modular cooling arrangement and/or anchoring structures to which the modular cooling arrangement may be affixed.

[0021] In the following detailed description, various specific details are set forth in order to provide a thorough understanding of such embodiments. However, those skilled in the art will understand that disclosed embodiments may be practiced without these specific details that the aspects of the present invention are not limited to the disclosed embodiments, and that aspects of the present invention may be practiced in a variety of alternative embodiments. In other instances, methods, procedures, and components, which would he well- understood by one skilled in the art have not been described in detail to avoid unnecessary and burdensome explanation.

[0022] Furthermore, various operations may be described as multiple discrete steps performed in a manner that is helpful for understanding embodiments of the present invention. However, the order of description should not be construed as to imply that these operations need be performed in the order they are presented, nor that they are even order dependent, unless otherwise indicated. Moreover, repeated usage of the phrase“in one embodiment” does not necessarily refer to the same embodiment, although it may. It is noted that disclosed embodiments need not be construed as mutually exclusive embodiments, since aspects of such disclosed embodiments may be appropriately combined by one skilled in the art depending on the needs of a given application.

[0023] FIG. 1 is an isometric view illustrating one non-limiting embodiment of a disclosed modular cooling arrangement, as may be embodied in an airfoil component 10, such as a vane, disposed in a gas turbine engine between anchoring structures, such as a radially -outer shroud 20a and a radiaily-inner shroud 20b.

[0024] FIG. 2 is a cross-sectional view of the disclosed modular cooling arrangement shown in FIG. I and illustrates one non-limiting embodiment of a cooling manifold 18, such as may he disposed in a cavity in a body 16 of airfoil component 10. In this embodiment, cooling manifold 18 may be formed by a unitary cooling manifold structure that constitutes a load-bearing structure with respect to radially-outer shroud 20a and radially-inner shroud 20b to which cooling manifold 18 and body 16 of airfoil component 10 are affixed.

Body 16 of airfoil component 10, without limitation, may comprise a CMC material; a ternary ceramic, referred to in the art as a MAX phase, or a metal, such as may comprise a relatively less costly material compared to a high temperature metal. Non-limiting examples of relatively less costly metals may include Haste Hoy X, Inconel alloy 625, etc. For readers desirous of background information in connection with the foregoing ternary ceramic, reference is made to article authored by M. Radovic and M. W. Barsoum, titled‘'MAX phases: Bridging the gap between metals and ceramics”,

American Ceramic Society Bulletin, Vol. 92, Nr 3, p. 20-27 (April 2013), which is incorporated herein by reference.

[0026] Non-limiting examples of MAX phases, which may be used to produce the body of the airfoil component, may include a family of ternary ceramics having a M„÷i AX_r. chemical configuration, where n=l, 2, or 3, M is an early transition metal, such as Ti, V, Cr, Zr, Nb, Mo, Hfi Sc, Ta, A is an A-group element, such as Al, Si, P, S, (ta, Ge, As, Cd, In, Sn, T!, Pb; and X is C and/or N. Due to its layered structure. As will be appreciated by those skilled in the art, a MAX phase material can exhibit a unique combination of mechanical and thermal properties. For example, MAX phases can substantially dissipate potentially harmful structural vibrations or acoustic loads, even at relatively high temperatures.

[ 0027 j Cooling manifold 18 may include an arrangement of cooling passageways 22 (FIG. 2) configured to pass a flow of cooling fluid between a cooling manifold inlet 24 and a cooling manifold outlet 26. In one non-limiting embodiment, a segment of manifold inlet 24 and manifold outlet 26, such as may protrude beyond the respective outer sides of radial ly-outer shroud 20a and radiaily- inner shroud 20b, may be threaded and may he affixed using nuts (not shown) or any other suitable affixing structure. Without limitation, the arrangement of cooling passageways may comprise a tortuous (e.g., a serpentine) arrangement of cooling passageways 22. [0028] In one non-limiting embodiment, it is contemplated use of three-dimensional (3D) Printing/ Additive Manufacturing (AM) technologies, such as laser sintering, selective laser melting (SLM). direct rnetal laser sintering (DMLS), electron beam sintering (BBS), electron beam melting (EBM), etc., that may be conducive to cost-effective fabrication of disclosed cooling manifolds that may involve complex geometries and miniaturized cooling features and/or conduits. That is, such miniaturized cooling features and/or conduits can be optimized for the most effective coolant distribution and heat extraction. For readers desirous of general background information in connection with 3D Printing/ Additive Manufacturing (AM) technologies, see, for example, a textbook titled“Additive Manufacturing Technologies, 3D Printing, Rapid Prototyping, and Direct Digital Manufacturing”, by Gibson I., Stucker B., and Rosen D., 2010, published by Springer, which is incorporated herein by reference.

[0029] In one non-limiting embodiment, as may be appreciated in FIG. 2, the

arrangement of cooling passageways 22 may be closed with respect to the cavity in the body of the airfoil component where cooling manifold 18 is disposed. A gap 28 (FIG. 4), such as may comprise a relatively narrow gap (e.g., in a range from approximately 3 mm to approximately 10 mm) may be defined between body 16 of the airfoil component and cooling manifold 18 Without limitation, in this embodiment, gap 28 may facilitate radiative cooling between cooling manifold 18 and body 16 of the airfoil component. As may be appreciated in FIG.4, cooling manifold 18 may have a profile that geometrically matches a profile of body 16 of airfoil component 10.

[0030] As may be appreciated in FIG. 3, the arrangement of cooling passageways 22 in cooling manifold 18 may include apertures 30 arranged to provide metered fluid communication with the cavity in the body 16 of airfoil component 10 where cooling manifold 18 is disposed. In this embodiment the gap between body 16 of the airfoil component and cooling manifold 18 need not be configured as a narrow gap and may comprise a relatively wider gap in the order of a few centimeters (e.g., in a range from approximately 2 cm to approximately 8 cm) to facilitate convective cooling of body 16 of airfoil component 10.

[0031] In one non-limiting embodiment, the anchoring structure (e.g., radiaily-outer shroud 20a and radial ly-inner shroud 20b) to which airfoil component 10 may be affixed includes recessed pockets 84, 35 (FIG. 5) configured to receive corresponding complementary surfaces 38, 36 (FIG. 2) of cooling manifold 18 and body 16 of the airfoil component,

[0032] FIG. 6-7 are respective isometric views of another non-limiting embodiment of a disclosed modular cooling arrangement involving a How sleeve 50, as illustrated in FIG. 8, that may he disposed in the body of airfoil component 10. Flow sleeve 50 includes a flow sleeve inlet 52 and a flow sleeve outlet 54 extending radially between radiaily-outer shroud 20a and radially-inner shroud 20b to pass a flow of cooling fluid between flow sleeve inlet 52 and flow sleeve outlet 54. Pins 40a and 40b may be used to affix the body of airfoil component 10 to radiaily-outer shroud 20a and radialiy-inner shroud 20b.

3] In one non-limiting embodiment, both opposite radial ends of flow sleeve 50 may be affixed onto radial iy-outer shroud 20a and radial!y-inner shroud 20b. Alternatively, just one end of the opposite radial ends of flow sleeve 50 may be affixed to a respective one of radial! y-outer shroud 20a or radiaily-inner shroud 20b and the other radial end of flow sleeve 50 may be a free end (not affixed to the other respective one of shroud radially -outer shroud 20a or radiaily-inner shroud 20b) to accept radial displacement of the flow sleeve with respect to the other respective one of radially-outer shroud 2.0a or radiaily-inner shroud 20b. This approach may also be used in connection with the disclosed modular cooling arrangement involving a cooling manifold, as described above in the context of figures one through five.

[0034] As raay be appreciated in FIG. S, a cooling passageway defined by flow sleeve 50 may be closed with respect to the cavity in the body of the airfoil component where flow sleeve 50 is disposed. Alternatively, as shown in FIG.

9, the cooling passageway defined by another disclosed flow sleeve 60 may include apertures 62 to provide metered fluid communication with the cavity in the body of the airfoil component where flow_'· sleeve 60 is disposed.

[0035] FIG. 10 is a cross-sectional view of a disclosed modular cooling arrangement involving a multi-piece cooling manifold arrangement including cooling manifold pieces 18a. i 8b and including a cooling manifold piece 18c configured to provide cooling to a trailing edge region 70 of airfoil component

10, as may involve flow ejection at the trailing edge of airfoil component 10. Cooling manifold piece 18a may be fluidly and mechanically interconnected to cooling manifold piece 18b by way of a plurality of tubelets 72.

[0036] FIG. 11 is a cross-sectional view illustrating non-limiting structural details of cooling manifold piece 18c configured to provide cooling to trailing edge region 70 of the airfoil component. Cooling manifold piece 18c of the multi- piece cooling manifold arrangement may include a radially staggered array of tubes 74 to convey cooling fluid to the trailing edge region. Depending on the needs of a given application, reinforcement webs 76 may be formed between mutually adjacent tubes 74 for structural reinforcement. [0037] From the foregoing disclosure, it should he appreciated that disclosed embodiments provide a modular cooling arrangement that may be flexibly arranged for cooling airfoil components, such as blades or vanes, in a combustion turbine engine. Without limitation, disclosed embodiments are effective to reliably and cost-effectively reduce a level of internal pressurization in the component while providing appropriate cooling to portions of the component subject to a hot flow of gases and accommodating thermal gro wth differences that may develop between components of the modular cooling arrangement and anchoring structures to which the modular cooling arrangement may be affixed.

[0038] While embodiments of the present disclosure have been disclosed in

exemplary forms, it will be apparent to those skilled in the art that many modifications, additions, and deletions can be made therein without departing from the scope of the invention and its equivalents, as set forth in the following claims.

Claims

What is claimed is:

1. A modular cooling arrangement for cooling an airfoil component (10) in a combustion turbine engine, the modular cooling arrangement comprising: a body (16) of the airfoil component; and

a cooling manifold (18) disposed in a cavity in the body of the airfoil component, the cooling manifold arranged to mechanically support the body of the airfoil component.

2. The modular cooling arrangement of claim 1, wherein the cooling manifold constitutes a load-bearing structure with respect to at least one anchoring structure (20a, 20b) to which the airfoil component is affixed.

3. The modular cooling arrangement of claim 2, wherein the cooling manifold comprises a profile that geometrically matches a profile of the body of the airfoil component.

4. The modular cooling arrangement of claim 3, the cooling manifold including an arrangement of cooling passageways (22) configured to pass a flow of cooling fluid between a cooling manifold inlet (24) and a cooling manifold output (26).

5. The modular cooling arrangement of claim 4, wherein the arrangement of cooling passageways (22) comprises a tortuous arrangement of cooling

passageways.

6. The modular cooling arrangement of claim 4, wherein the arrangement of cooling passageways is closed with respect to the cavity in the body of the airfoil component where the cooling manifold is disposed.

7. The modular cooling arrangement of claim 5, wherein a gap (28) is defined between the body of the airfoil component and the cooling manifold, wherein the gap facilitates radiative cooling between the cooling manifold and the body of the airfoil component.

8. The modular cooling arrangement of claim 4, wherein the arrangement of cooling passageways comprises apertures (30) to provide metered fluid

communication with the cavity in the body of the airfoil component where the cooling manifold is disposed.

9. The modular cooling arrangement of claim 1, wherein the airfoil component is selected from the group consisting of a blade and a vane.

10. The modular cooling arrangement of claim 7, wherein the at least one anchoring structure to which the vane is affixed comprises a radially-inner shroud and a radially-inner shroud, each including recessed pockets 35, 34 configured to receive corresponding complementary surfaces 36, 38 of the body of the airfoil component and the cooling manifold.

11. The modular cooling arrangement of claim 1, wherein the cooling manifold comprises a unitary cooling manifold structure.

12. The modular cooling arrangement of claim 1, wherein the cooling manifold comprises a multi-piece cooling manifold arrangement.

13. The modular cooling arrangement of claim 12, wherein the multi-piece cooling manifold arrangement comprises a cooling manifold piece configured to provide cooling to a trailing edge region (70) of the airfoil component.

14. The modular cooling arrangement of claim 1, wherein the body of the airfoil component is selected from the group consisting of a ceramic matrix composite body, a MAX phase ternary ceramic body, and a metallic body.

15. A modular cooling arrangement for an airfoil component in a combustion turbine engine, the modular cooling arrangement comprising:

a body (16) of the airfoil component; and

at least one flow sleeve (50, 60) disposed in a cavity in the body of the airfoil component,

the at least one flow sleeve extending radially between a radi ally-outer shroud 20a and a radially-inner shroud 20b.

16. The modular cooling arrangement of claim 15, wherein the at least one flow sleeve (50, 60) defines a cooling passageway configured to pass a flow of cooling fluid between a flow sleeve inlet and a flow sleeve outlet.

17. The modular cooling arrangement of claim 16, wherein the cooling passageways is closed with respect to the cavity in the body of the airfoil component where the at least one flow sleeve (50) is disposed.

18. The modular cooling arrangement of claim 16, wherein the cooling passageway comprises apertures (62) to provide metered fluid communication with the cavity in the body of the airfoil component where the at least one flow sleeve (60) is disposed.

19. The modular cooling arrangement of claim 15, wherein one radial end of the at least one flow sleeve (50, 60) is connected to one of the radially-outer shroud 20a or the radially-inner shroud 20b, and the other radial end of flow sleeve (50, 60) is a free end not affixed to the other one of radially-outer shroud 20a or radially-inner shroud 20b to accept radial displacement of the flow sleeve (50, 60) with respect to the other one of radially-outer shroud 20a or radially-inner shroud 20b.

20. The modular cooling arrangement of claim 15, wherein the body of the airfoil component is selected from the group consisting of a ceramic matrix composite body, a MAX phase ternary ceramic body, and a metallic body.