COOLING AND MANUFACTURING

BACKGROUND 1. Field

[0001] The present application relates generally to gas turbines, and more particularly to utilizing an insert in rotating turbine blades for cooling air distribution.

2. Description of the Related Art [0002] Hot gas path components, such as blades and vanes of gas turbine engines, are typically exposed to high thermal loads during gas turbine operation. A flow of a hot gas is generated when a mixture of compressed air and a fuel are ignited in a combustor section of the gas turbine. The hot gas flows into the turbine section, which includes the blades and vanes. The temperatures to which the blades and vanes are exposed due to the flow of hot gas may be upwards of 450° C and possibly even as high as 1400-1600° C in the flow path.

[0003] The heat transfer rate and cooling effectiveness between cooling fluids and hot gas path components in a gas turbine engine directly correlates to the overall efficiency of the gas turbine. The more efficiently that heat is removed from the component, the higher the overall efficiency that can be achieved. Additionally, a decrease in the temperature of the hot gas component may increase the life of the component.

[0004] Various conventional methods are used to cool hot gas path components. Cast-in heat transfer features, impingement cooling of backside hot gas path surfaces, and multi-circuit cooling passages are some of the methods in use to improve hot component cooling.

[0005] A gas turbine component’s ability to transfer heat away from itself is particularly important due to the high operating temperatures of the engine. One way to enhance the cooling capability is to increase the component’s surface area through the incorporation of heat transfer features. Another way is to direct cooling air through intricate passages within the rotating turbine blade itself. The incorporation of heat transfer features onto a surface of the turbine blade and intricate passages into interior surface of the turbine blade add considerable cost and complexity to the casting process. Additionally, by casting these features within the structure, the temperature gradient within the structure results in high stresses in areas where cooling features or circuits are attached to the highly heated surfaces.

[0006] Consequently, a more versatile and inexpensive process to incorporate heat transfer features to increase heat transfer onto a surface of the turbine blade and air distribution passages to distribute cooling air within the interior of rotating turbine blades over the current casting process is desired.

SUMMARY

[0007] Briefly described, aspects of the present disclosure relate to a cooled turbine blade assembly and a new method of incorporating heat transfer and cooling air management features within rotating components.

[0008] A first aspect provides a cooled turbine blade assembly. The cooled turbine blade assembly includes a hollow turbine blade airfoil extending span-wise radially outward from a platform. The turbine blade airfoil includes a leading edge and a trailing edge and a suction side with a suction side wall and a pressure side with a pressure side wall. Additionally, the cooled turbine blade assembly includes a root portion extending radially inward from the platform and includes at least one inlet for cooling air. A blade insert is disposed within the cavity, the blade insert including a geometry to distribute a flow of cooling air within the turbine blade airfoil.

[0009] A second aspect provides a method for cooling a turbine blade. The method includes providing a blade body including a hollow turbine blade airfoil extending span-wise radially outward from a platform. The turbine blade airfoil includes a leading edge and a trailing edge and a suction side with a suction side wall and a pressure side with a pressure side wall. Additionally, the cooled turbine blade assembly includes a root portion extending radially inward from the platform and includes at least one inlet for cooling air. The method also includes inserting a blade insert through an opening at a blade tip opposite the root portion and into a cavity of the hollow turbine blade airfoil. The blade is attached to the root portion of the blade body. A flow of cooling air with the cavity distributed by the blade insert during turbine operation when the blade body is exposed to a flow of a hot gas cools the blade body.

[0010] A third aspect provides a method for manufacturing a cooled turbine blade assembly. The method includes casting a blade body of a turbine blade. The body of the turbine blade includes a hollow turbine blade airfoil extending span-wise radially outward from a platform. The turbine blade airfoil includes a leading edge and a trailing edge and a suction side with a suction side wall and a pressure side with a pressure side wall. A cooling component is produced and configured to distribute a flow of cooling air within the turbine blade airfoil. The method also includes inserting the cooling component through an opening at a blade tip opposite the root portion and into a cavity of the hollow turbine blade airfoil. The cooling component is attached to the root portion of the blade body.

BRIEF DESCRIPTION OF THE DRAWINGS [0011] Fig. 1 is a cross sectional suction side view of a prior art turbine blade with two triple pass serpentine flow cooling circuits,

[0012] Fig. 2 illustrates a perspective view of a cooled turbine blade assembly,

[0013] Fig. 3 illustrates a side view of an embodiment of blade inserts,

[0014] Fig. 4 illustrates a perspective view of an embodiment of blade inserts with curved sides,

[0015] Fig. 5 illustrates a cross sectional view of a turbine blade with blade inserts installed, and [0016] Fig. 6 illustrates a partially exploded perspective view of a cooled turbine blade assembly.

DETAILED DESCRIPTION

[0017] To facilitate an understanding of embodiments, principles, and features of the present disclosure, they are explained hereinafter with reference to

implementation in illustrative embodiments. Embodiments of the present disclosure, however, are not limited to use in the described systems or methods.

[0018] The components and materials described hereinafter as making up the various embodiments are intended to be illustrative and not restrictive. Many suitable components and materials that would perform the same or a similar function as the materials described herein are intended to be embraced within the scope of embodiments of the present disclosure.

[0019] Referring now to the figures, where the showings are for purposes of illustrating embodiments of the subject matter herein only and not for limiting the same, Fig. 1 illustrates cross sectional suction side view of a prior art turbine blade 10 with two internal triple pass serpentine flow cooling channels 14, 16. The turbine blade airfoil 22 includes a pressure side (not shown), a suction side (not shown), a leading edge 24 and a trailing edge 26. A flow of cooling air (as shown by arrows) enters the serpentine flow cooling circuits 14, 16 through an inlet 12 in the blade root 20 and follows the serpentine channel as shown by the arrows. Air may flow out of the blade 10 through the trailing edge 26 or the blade tip 30 as shown. Additionally, air may flow out through other locations on the airfoil 22. The serpentine channel 14, 16 is constructed to be as long as possible so that a maximum heat transfer may occur as the cooling air flow winds through the serpentine channel 14, 16. Heat transfer features 32 may also be disposed within the internal structure to increase the heat transfer from the blade 10 to the cooling air. As an example, Fig. 1 shows a spiked heat transfer feature 32 projecting into the serpentine flow channel 14 from a wall of the channel. The serpentine channels 14, 16 as well as the heat transfer features are all intricate internal features that must be cast into the turbine blade 10 adding significant cost and effort to the manufacture of the turbine blade 10.

[0020] By recognizing that the internal cast-in cooling distribution features of a rotating turbine blade are difficult to manufacture as well as adding a significant cost to the production of the turbine blade, the current inventors propose utilizing a separate component, or insert, that contains the primary geometries for distributing the cooling air within the turbine blade.

[0021] Fig. 2 illustrates a perspective view of an embodiment of a cooled turbine blade assembly 100. The turbine blade assembly 100 comprises a hollow turbine blade airfoil 140 including a blade tip 130, the turbine blade airfoil 140 extending radially outward from a platform 160. The hollow turbine blade airfoil 140 includes a leading edge 124 and a trailing edge 126 as well as a pressure side having a pressure side wall and a suction side with a suction side wall. Radially inward from the platform 160 is the root section 120. The turbine blade assembly 100 further comprises at least one blade insert 150 extending at least in a span-wise direction within a cavity of the hollow turbine blade airfoil 140.

[0022] Fig. 3 illustrates a side view of an embodiment of two turbine blade inserts 150. In the shown embodiment, the inserts 150 are embodied as plates such that they extend span-wise from the root section 120 of the rotating turbine blade 10 within the cooling channels of the blade root 120 through the platform 160 and into the hollow airfoil 140. The blade inserts 150 may also extend past the radially inward portion of the blade tip 130 as shown in Fig. 2. In the example as shown in Fig. 2, the blade tip 130 is not integrally cast with the airfoil body 140; it is attached during manufacture by a welding or braze process. However, it may also be integral to the insert 150.

[0023] The thickness of the plates may be as thick as needed and depends on the width of the turbine blade. Depending on the thickness of the plates, a gap (g) may exist between an inner surface of the hollow airfoil 140 and an outer surface of the blade insert 150. Fig. 5 shows a cross sectional view of the turbine airfoil 140 with the blade inserts 150 of Figs.3-4 installed within the cavity. In the shown embodiment, ribs 132 extend from the pressure side wall 128 to the suction side wall 122. On either side of the rib 132 two openings open into a respective cavity within the hollow turbine blade 10 where the blade inserts 150 are disposed. The gap (g) between the inner surface of the airfoil and an outer surface of the insert 150 may be as needed to set the velocity of the cooling air flow. Alternately, the blade insert (150) may slightly touch or abut the inner surface of the hollow airfoil 140.

[0024] In an embodiment, the blade insert 150, embodied as a plate as shown in Fig.4, includes a geometry of a channel so that a flow of cooling air may be distributed throughout the turbine blade airfoil 140. In this embodiment, the insert 150 may include curved edges to create a channel for a fluid flow similarly to the functionality of the serpentine channels 14, 16 shown in Fig. 1. Once the blade insert 150 is installed within the rotating turbine blade 10, during operation, the cooling air may flow into the inlet 12 of the blade root 120, through the channel of the plate and out through trailing edge holes, openings in the blade tip 130, the leading edge 124, and/or out through the side walls. In order to control the degree to which the cooling air is distributed within the interior of the blade airfoil 140, the size of the channels may be increased and/or decreased depending on the particular cooling requirements

[0025] The blade inserts 150 may be retained within the blade airfoil 140 using various techniques. In the embodiment shown in Fig. 2, the blade inserts 150 are attached to the blade root 120 by a mechanical connection. The mechanical connection may include, but is not limited to, a threaded connection (shown in Fig. 2) or a bayonet connection between a section of the cooling geometry inserted into the blade from the tip area of the blade and another part of the cooling geometry inserted from the root area of the blade. In an alternate embodiment, the insert 150 may be retained within the blade airfoil 140 by a brazed a connection. In addition to the techniques described here to retain the blade insert 150 within the blade airfoil 140, other techniques may also be utilized.

[0026] The blade insert 150 may include a heat transfer feature 32 disposed on/incorporated into the outer surface of the blade insert 150 to promote the heat transfer into the cooling air flow. The heat transfer feature 32 may be a variety of shapes according to the cooling requirements of the turbine blade 10. For example, the shape of the heat transfer feature(s) may include pins, waves, chevrons, spikes, ribs, and fins. These listed shapes are for exemplary purposes only and are not meant to be limiting. The embodiment of a heat transfer feature as a spike as illustrated in Fig. 1 is an example of what the heat transfer feature may be. In an embodiment, the spiked heat transfer feature 32 may be disposed on the curved edge of the blade insert 150 so that it projects into the blade insert channel. In an alternate embodiment, the heat transfer feature 32 may be disposed on an inner surface of the turbine blade 10.

[0027] In an embodiment, the blade insert 150 may comprise a material that is the same material or a similar material as that of the turbine blade 10. Cooled turbine components may be formed from a superalloy or nickel-based alloy, such as CM 247, IN939, IN617, IN735, IN718, IN625, Haynes 282, Haynes 230, Hast-X, and Hast-W. For example, the blade insert 150 may comprise a material whose rate of thermal expansion is similar to that of the turbine blade 10 so that upon exposure to the flow of hot gases, the gap g (if one exists) between an inner surface of the hollow airfoil 140 and an outer surface of the blade insert 150 is maintained.

[0028] Referring now to Figs. 1-6, a method for cooling a turbine blade assembly 10 is proposed. The blade insert 150, as described above, is utilized for distributing a flow of cooling air within the interior of a hollow turbine blade airfoil 140. A blade body 130, 140, also described above and shown in Fig. 2, is provided. In an embodiment, the blade insert 150, which is manufactured as a separate component from the blade body, is inserted through an opening at the blade tip 130 opposite the root portion 120 and into a respective cavity of the hollow turbine blade airfoil 140. Fig. 6 shows a partially exploded perspective view of a plurality of blade inserts 150 during insertion into the turbine blade airfoil 140.

[0029] Each blade insert 150 may then be attached to the root section 120. A mechanical attachment is shown in Fig. 2, however, the blade insert 150 may be attached to the root section 120 by various other techniques such as brazing, for example. After the insertion and attachment, the blade tip 130 may then be attached to the blade airfoil 140. The attachment of the blade tip 130 to the blade airfoil 140 may be performed by welding, brazing or may be manufactured as integral with the blade insert 150. The blade insert 150 is effective to distribute a flow of cooling air within the turbine blade airfoil 140 during turbine operation when the blade body is exposed to a flow of hot gases. The distribution of the cooling air within the cavity of the hollow turbine blade airfoil 140 cools the turbine blade 10.

[0030] In an embodiment, the blade body 120, 130, 140, and 160 and the blade insert 150 may be manufactured by separate processes. Separating the manufacture of the blade body from the cooling geometries found in the interior of the blade body allows different manufacturing techniques and materials to be used for each. The requirements of the design of the blade body and the cooling geometries are significantly different. Manufacturing both as one single component significantly reduces the options available for manufacture and limits the geometries which may be incorporated into the design. By separating the manufacture of blade body from the cooling geometries, the cost may be reduced and the complexity of the geometries which may be applied can be increased significantly.

[0031] In an embodiment, according to the proposed method, the blade body 120, 130, 140, and 160 and the blade insert 150 may be cast separately. In another embodiment, the blade body 120, 130, 140, and 160 may be cast and the blade insert 150 may be manufactured using an additive manufacturing process. Alternately to casting the blade body 120, 130, 140, and 160, by simplifying the geometry of the blade body 120, 130, 140, and 160 and including a blade insert 150 , the blade bodyl20, 130, 140, and 160 may be produced by an additive manufacturing process. In a further embodiment, blade insert 150 may also be machined, for instance, in a retrofitting process.

[0032] In summary, the proposed turbine blade assembly and method utilize a separate component for the management of cooling air within a rotating turbine airfoil. The separate component, or insert, is then integrated into the airfoil. Utilizing a separate component for the management of cooling air distribution within a rotating turbine airfoil offers several advantages over prior art solutions. For example, by manufacturing the insert separately from the turbine airfoil, the complexity and precision which may be achieved with the cooling features can be greatly improved. Additionally, by manufacturing of the cooling features separately, the total cost of manufacturing the turbine airfoil can be greatly reduced. Moreover, utilizing an insert, the cooling features of an airfoil may be adjusted or tailored for specific applications or changed during the repair of the component. [0033] 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 spirit and scope of the invention and its equivalents, as set forth in the following claims.