US20150064019A1

US20150064019A1 - Gas Turbine Components with Porous Cooling Features

Info

Publication number: US20150064019A1
Application number: US14/014,528
Authority: US
Inventors: Benjamin Paul Lacy; Srikanth Chandrudu Kottilingam; Brian Brzek; David Edward Schick
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2013-08-30
Filing date: 2013-08-30
Publication date: 2015-03-05
Also published as: JP2015048847A; CN104420893A; DE102014111837A1; CH708570A2; CH708570A8

Abstract

The present application provides a hot gas path component for use with a gas turbine engine. The hot gas path component may include an airfoil, an internal cooling cavity, and a porous section created by a direct metal laser melting technique. The porous section may be built into the airfoil or the airfoil may be built separately and attached to the airfoil.

Description

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with U.S. government support under contract number DE-FC26-05NT42643 awarded by the Department of Energy. The government has certain rights in this invention.

TECHNICAL FIELD

The present application and the resultant patent relate generally to gas turbine engines and more specifically relate to gas turbine components with porous cooling sections created by direct metal laser melting manufacturing techniques and the like.

BACKGROUND OF THE INVENTION

Gas turbine systems are widely utilized in fields such as power generation. Overall gas turbine performance and efficiency generally may be increased by increasing internal combustion temperatures. The components that are subject to the high temperatures in the hot gas path, however, must be cooled. For example, an airfoil and other components of a nozzle and the like may be disposed in the hot gas path and exposed to the relatively high combustion temperatures. A cooling flow therefore may be routed from the compressor or elsewhere and provided to the various components in the hot gas path.
A variety of methods may be used for cooling the airfoils and the other components. These methods may include running a cooling flow on the interior side of the component, running the cooling flow through an impingement sleeve that impinges the flow on the backside of the component so as to increase the heat transfer coefficient therein, running the coolant through cooling holes to the exterior of the component to convectively cool, and exhausting the coolant from the cooling holes as film to provide a layer of cool air over the exterior so as to reduce exterior temperatures. Although the use these methods may provide adequate cooling for the airfoils, a further increase in cooling efficiency is desired. Such an increase in efficiency would allow a reduction in the cooling flows required to cool the airfoils and other components and also may provide a reduction in emissions and/or an increase in firing temperatures.

SUMMARY OF THE INVENTION

The present application and the resultant patent thus provide a hot gas path component for use with a gas turbine engine. The hot gas path component may include an airfoil, an internal cooling cavity, and a porous section created by a direct metal laser melting technique. The porous section may be built into the airfoil or the airfoil may be built separately and attached to the airfoil.
The present application and the resultant patent further provide a method of cooling a hot gas path component for use with a gas turbine engine. The method may include the steps of providing the hot gas path component with an internal cooling cavity, creating a porous section via a direct metal laser melting technique, flowing a cooling medium to the internal cooling cavity, and flowing the cooling medium through the porous section to provide transpiration cooling. The creating step may include building up the porous section on the hot gas path component or building the porous section separately and attaching the porous section to the hot gas path component.
The present application and the resultant patent further provide an airfoil for use with a gas turbine engine. The airfoil may include a pressure side, a suction side, an internal cooling cavity, and a porous section with a porous media created by a direct metal laser melting technique.
These and other features and improvements of the present application and the resultant patent will become apparent to one of ordinary skill in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic diagram of a gas turbine engine showing a compressor, a combustor, and a turbine.

FIG. 2 is a sectional view of a portion of an airfoil.

FIG. 3 is a sectional view of a portion of an airfoil as may be described herein.

FIG. 4 is an expanded view of a portion of the airfoil of FIG. 3.

FIG. 5 is a sectional view of an alternative embodiment of an airfoil as may be described herein.

FIG. 6 is an expanded view of a portion of the airfoil of FIG. 5.

FIG. 7 is a sectional view of an alternative embodiment of an airfoil as may be described herein.

FIG. 8 is an expanded view of a portion of the airfoil of FIG. 7.

FIG. 9 is an expanded view of an alternative embodiment of a portion of the airfoil of FIG. 7.

FIG. 10 is a sectional view of an alternative embodiment of an airfoil as may be described herein.

FIG. 11 is an expanded view of a portion of the airfoil of FIG. 10.

DETAILED DESCRIPTION

Referring now to the drawings, in which like numerals refer to like elements throughout the several views, FIG. 1 shows a schematic view of gas turbine engine 10 as may be used herein. The gas turbine engine 10 may include a compressor 15. The compressor 15 compresses an incoming flow of air 20. The compressor 15 delivers the compressed flow of air 20 to a combustor 25. The combustor 25 mixes the compressed flow of air 20 with a pressurized flow of fuel 30 and ignites the mixture to create a flow of combustion gases 35. Although only a single combustor 25 is shown, the gas turbine engine 10 may include any number of combustors 25. The flow of combustion gases 35 is in turn delivered to a turbine 40. The flow of combustion gases 35 drives the turbine 40 so as to produce mechanical work. The mechanical work produced in the turbine 40 drives the compressor 15 via a shaft 45 and an external load 50 such as an electrical generator and the like.
The gas turbine engine 10 may use natural gas, liquid fuels, various types of syngas, and/or other types of fuels and combinations thereof. The gas turbine engine 10 may be any one of a number of different gas turbine engines offered by General Electric Company of Schenectady, N.Y., including, but not limited to, those such as a 7 or a 9 series heavy duty gas turbine engine and the like. The gas turbine engine 10 may have different configurations and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines, other types of turbines, and other types of power generation equipment also may be used herein together.
FIG. 2 shows a sectional view of an example of a hot gas path component 55. In this example, the hot gas path component 55 may be an airfoil 60. The airfoil 60 may be part of a nozzle, a bucket, or any other type of hot gas path component 55 such as a shroud and the like. The airfoil 60 may include an outer shell 65. The airfoil 60 may extend from a pressure side 70 to a suction side 75. The airfoil 60 also may extend from a leading edge 80 to a trailing edge 85. The airfoil 60 may have an overall aerodynamic shape. The shell 65 may define a number of internal cooling cavities 90 in communication with a number of film cooling holes 92 extending through the shell 65. A number of pin banks 94 also may extend into the internal cooling cavities 90. A portion of the flow of air 20 may be diverted from the compressor 15 so as to cool the airfoil 60. The flow of air 20 may extend through the internal cooling cavities 90 and may exit about the film cooling holes 92 or elsewhere. The pin banks 94 may provide turbulence to the flow of air 20. Many other types of hot gas path components 55 and airfoils 60 may be used. Likewise many different types of cooling schemes and components also may be used.
FIGS. 3 and 4 show a hot gas path component 100 as may be described herein. In this example, the hot gas path component 100 may be an airfoil 110. The airfoil 110 may be part of a nozzle or a bucket. Other types of hot gas path components 100, such as a shroud and the like, also may be used herein. The airfoil 110 may include a shell 120. The shell 120 may have an interior surface 130 and an exterior surface 140. The interior surface 130 may have an impingement sleeve 135, an impingement plate, or a similar type of structure adjacent thereto. The shell 130 may extend from a pressure side 150 to a suction side 160. Likewise, the airfoil 110 may extend from a leading edge 170 to a trailing edge 180 and may define a substantially aerodynamic shape. The shell 120 may define a number of internal cooling cavities 190 about the inner surface 130 thereof. A number of film cooling holes 200 may extend through the shell 120. A number of pin banks 210 also may be positioned within the internal cavities 190. Other components and other configurations also may be used herein.
The airfoil 110 also may have a porous trailing edge section 220. The porous trailing edge section 220 may be filled with a porous media 230. The porous media 230 may be formed from any suitable porous material or materials having a matrix with a number of voids therein. The porous media 230 may be formed from a metal foam, a metal alloy foam, a ceramic foam, such as a ceramic matrix composite foam, a carbon fiber foam, and similar types of porous materials. Non-limiting examples of specific materials may include Rene 142, Rene 195, MarM247, GTD111, GTD444, IN738, H282, H230, IN625 and the like. The foam typically may be formed by mixing a material, such as a metal, a ceramic, a carbon fiber, and the like with another substance and then melting the substance so as to leave the porous foam. The porous media 230 may be “printed” or built up via a direct metal laser melting (“DMLM”) process and the like. Different types of sintering techniques and other types of manufacturing techniques also may be used herein to create the components herein. The porous media may vary in porosity/permeability throughout based on optimizing the cooling flow therethrough. For example, permeability may be lowest in regions of highest heat load so that more coolant flows through these regions as compared to regions where the heat load and the coolant demand may be lower. A cooling medium 240 may flow through the voids in the porous media 230 so as to facilitate cooling in a highly efficient manner.
The porous trailing edge section 220 may be built directly onto the airfoil 110 or the porous trailing edge section 220 may be built separately and attached by any number of different techniques. These techniques may include including brazing, arc welding, high energy density welding such as laser welding and electron beam welding, TLP bonding, diffusion bonding, or different types of mechanical attachment. The buildup of the porous media 230 may be made over an existing component or as part of building a component as a whole. The use of the DMLM process enables high heat transfer through the porous media 230 while providing a high quality joint between the airfoil 110 and the porous trailing edge section 220. The porous trailing edge section 220 may have an external sleeve 250 extending in whole or in part to direct the flow to exit over only a certain section or sections of the trailing edge. The external sleeve 250 may be a metallic component, a thermal barrier coating, and the like. The coating may be an aluminide and the like sprayed thereon. The cooling medium 240 thus flows through the airfoil 110 and exits via the porous trailing edge section 220 so as to cool the trailing edge 180. Other components and other configurations may be used herein.
FIGS. 5 and 6 show a further example of a hot gas path component 100. In this example, the hot gas path component 100 may be an airfoil 260. The airfoil 260 may include a porous side section 270 positioned on the suction side 160. The porous side section 270 may include the porous media 230. Specifically the porous media 230 may be built or attached into the shell 120 of the airfoil 260 along the impingement sleeve 135 or about a grid on the underlying structure. The porous media 230 may be aligned with the shell so as to provide transpiration cooling and the like. The cooling flow 240 thus may leak through the voids in the porous side section 270. Any number of the porous side sections 270 may be used herein in any size or shape. As above, the porosity and the permeability may be varied throughout the porous piece so as to optimize cooling usage. Other components and other configurations may be used herein.
FIGS. 7-9 show a further example of a hot gas path component 100. In this example, the hot gas path component 100 may be an airfoil 280. The airfoil 280 may include a porous external section 290 positioned on the suction side 160 or elsewhere along the airfoil 280. Specifically, the porous external section 290 may include a buildup of the porous media 230 on the shell 120. Alternatively, the porous section may be built up separately and attached by any number of different techniques including those mentioned above. The shell 120 and the porous external section 290 may be in communication with the film cooling holes 200 extending through the shell 120. As is shown in FIG. 8, an external sleeve 300 may be used over the porous media 230. A number of external film cooling holes 310 may be positioned on the external sleeve 300. The external sleeve 300 may be metallic, a thermal barrier coating, and the like. As is shown in FIG. 9, the external sleeve 300 may be optional such that the porous media 230 may not need any type of covering. The external sleeve 300 may cover all, part, or none of the porous media 230. The cooling flow 240 thus may flow through the film cooling holes 200, the porous media 230 and/or the external film cooling holes 310. The film cooling holes may be partially formed in the porous media to improve flow distribution into the porous media. Other components and other configurations also may be used herein.
FIGS. 10 and 11 show a further embodiment of a hot gas path component 100. In this example, the hot gas path component 100 may be an airfoil 320. The airfoil 320 may include a porous internal section 330. The porous internal section 330 may include a buildup of the porous media 230 about the optional impingement sleeve 135 along the film cooling holes 200 or elsewhere within the shell 120 in whole or in part. Alternatively, the porous media may be built separately and attached by a variety of methods such as those described above. The permeability and porosity of the porous media may vary as needed to optimize coolant usage. The cooling flow 240 thus may flow through the impingement sleeve 135, the porous media 230, and the film cooling holes 200. The film cooling holes may be partially formed in the porous media to ensure an optimal film hole shape for maximized film effectiveness. Other components and other configurations may be used herein.
A number of alternative hot gas path components 100 also may be used herein. Specifically, DMLM techniques may be used to build both porous and solid features of the hot gas path component 100. These DMLM techniques may be used to vary the porosities and/or the permeability at different locations within the porous media 230. The DMLM techniques thus can be used to build multiple different discrete porous structures inside or outside thereof. Other methods of making and attaching the porous material may be used as well.
The hot gas path component 100 provides these integral porous features so as to enable better heat transfer as well as providing transpiration cooling. The use of the porous media 230 thus should reduce overall cooling load requirements. Specifically, the porous media has been shown to have a significantly higher heat transfer coefficient as compared to known airfoil materials as well as provides superior control over the distribution of coolant over the part. Using such a process on the hot gas path components in multiple locations may increase heat transfer capability while reducing cooling flow requirements. Moreover, the use of the DMLM process provides the porous foam with an integral joint to the base metal when built directly onto the part or as a whole with the part. The DMLM process also provides control over the porosity and the permeability throughout the part.
It should be apparent that the foregoing relates only to certain embodiments of the present application and the resultant patent. Numerous changes and modifications may be made herein by one of ordinary skill in the art without departing from the general spirit and scope of the invention as defined by the following claims and the equivalents thereof.

Claims

We claim:

1. A hot gas path component for use with a gas turbine engine, comprising:

an airfoil;

an internal cooling cavity; and

a porous section created by a direct metal laser melting technique.

2. The hot gas path component of claim 1, wherein the porous section is built into the airfoil.

3. The hot gas path component of claim 1, wherein the porous section is built separately and attached to the airfoil.

4. The hot gas path component of claim 1, wherein the airfoil comprises a pressure side and a suction side and wherein the porous section is within the suction side.

5. The hot gas path component of claim 1, further comprising an impingement sleeve adjacent to the porous section.

6. The hot gas path component of claim 1, further comprising a plurality of film cooling holes adjacent to the porous section.

7. The hot gas path component of claim 1, wherein the porous section comprises a porous media therein.

8. The hot gas path component of claim 7, wherein the porous media comprises a metal foam, a ceramic foam, and/or a carbon fiber foam.

9. The hot gas path component of claim 1, wherein the porous section comprises a porous trailing edge section with an external sleeve thereon in whole or in part.

10. The hot gas path component of claim 1, wherein the porous section comprises one or more porous side sections.

11. The hot gas path component of claim 1, wherein the porous section comprises a porous external section.

12. The hot gas path component of claim 11, wherein the porous external section comprises an external sleeve with a plurality of external film cooling holes.

13. The hot gas path component of claim 1, wherein the porous section comprises a porous internal section.

14. A method of cooling a hot gas path component for use with a gas turbine engine, comprising:

providing the hot gas path component with an internal cooling cavity;

creating a porous section via a direct metal laser melting technique;

flowing a cooling medium to the internal cooling cavity; and

flowing the cooling medium through the porous section to provide transpiration cooling.

15. The method of claim 14, wherein the creating step comprises building up the porous section on the hot gas path component or building the porous section separately and attaching the porous section to the hot gas path component.

16. An airfoil for use with a gas turbine engine, comprising:

a pressure side;

a suction side;

an internal cooling cavity; and

a porous section created by a direct metal laser melting technique.

17. The airfoil of claim 16, wherein the porous section is built into the airfoil or the porous section is built separately and attached to the airfoil.

18. The airfoil of claim 16, wherein the porous section comprises a porous trailing edge section.

19. The airfoil of claim 16, wherein the porous section comprises one or more porous side sections.

20. The airfoil of claim 16, wherein the porous section comprises a porous external section and/or a porous internal section.