Turbine nozzle or blade with impingement cooling structure having thermal flex elements - Patent Application 20070122997

Abstract

The turbine nozzle or blade includes an airfoil body defined by concave pressure side and convex suction side outer walls that connect along leading and trailing edges and form a radially extending chamber therebetween. The airfoil body has an inner surface facing the radially extending chamber. An impingement cooling structure is within the radially extending chamber. The impingement cooling structure includes a wall spaced from the inner surface of the airfoil body, a plurality of holes defined through the wall, and a plurality of elongated thermal flex elements defined in the wall. Because the nozzle or blade is fabricated by additive manufacturing, the airfoil body and impingement cooling structure include multiple monolithic layers of material.

Description

[Technical Field]

[0001] government contracts This application has been funded in part by U.S. Department of Energy contract DE-FE-0031611. The Government may have certain rights in this invention.

[0002] The present disclosure relates generally to turbine systems and, more particularly, to turbine nozzles or blades that include impingement cooling structures having thermal flex elements. [Background technology]

[0003] Additive manufacturing offers cost reduction opportunities by additively manufacturing parts together that were previously manufactured separately. However, additive manufacturing presents new challenges related to mitigating thermal-induced low-cycle fatigue (LCF) in components that were previously made in multiple parts but are now made as a single piece. For example, a turbine nozzle or blade is typically formed as separate parts, with its airfoil body and impingement insert, including impingement holes, mechanically coupled to each other. During use, the airfoil body is exposed to the hot gas path temperatures of the turbine's working fluid, while the impingement insert is exposed to a lower temperature coolant, e.g., compressor air at the compressor discharge temperature (Tcd). When formed as separate parts, the airfoil body and impingement insert can undergo their respective thermal cycles without incurring significant thermal stresses. However, when a turbine nozzle or blade is formed by additive manufacturing, the airfoil body and impingement insert are a single piece that is exposed to the hot gas path temperatures (Tfire) of the turbine's working fluid and coolant at a much lower temperature. Mitigating thermally induced LCF in such turbine nozzles or blades is a challenge. [Prior art documents] [Patent documents]

[0004] [Patent Document 1] U.S. Patent Application Publication No. 2021 / 0087979 Summary of the Invention

[0005] All aspects, examples, and features described below can be combined in any way technically possible.

[0006] One aspect of the present disclosure provides a turbine nozzle or blade comprising: an airfoil body defined by concave pressure side and convex suction side outer walls connecting along leading and trailing edges and forming a radially extending chamber therebetween, the airfoil body having an inner surface facing the radially extending chamber; and an impingement cooling structure within the radially extending chamber, the impingement cooling structure including a wall spaced from the inner surface of the airfoil body, a plurality of holes defined through the wall, and a plurality of elongated thermal flex elements defined in the wall, wherein the airfoil body and impingement cooling structure include multiple layers of unitary material.

[0007] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the plurality of elongated thermal flex elements extend in a direction perpendicular to the radial length of the impingement cooling structure.

[0008] Another aspect of the present disclosure includes any of the above aspects, wherein the plurality of elongated thermal flex elements extend in a direction at an angle ranging from 30° to 60° relative to a radial length of the impingement cooling structure.

[0009] Another aspect of the present disclosure includes any of the above aspects, wherein the plurality of elongated thermal flex elements extend in a direction at an angle of approximately 45 degrees relative to the radial length of the impingement cooling structure.

[0010] Another aspect of the present disclosure includes any of the above aspects, further comprising a plurality of support members spacing the wall from the inner surface of the airfoil body, the plurality of support members being disposed between the plurality of elongated thermal flex elements.

[0011] Another aspect of the present disclosure includes any of the above aspects, wherein the plurality of elongated thermal flex elements includes a plurality of elongated thermal flex elements between adjacent rows of the plurality of support members.

[0012] Another aspect of the present disclosure includes any of the above aspects, wherein the plurality of elongated thermal flex elements each have a C-shaped cross-section.

[0013] Another aspect of the present disclosure includes any of the above aspects, wherein each of the plurality of elongated thermal flex elements has one of a symmetrical V-shaped cross-section, an asymmetrical V-shaped cross-section, a rounded corner U-shaped cross-section, and a square corner U-shaped cross-section.

[0014] Another aspect of the present disclosure includes any of the above aspects, wherein the plurality of elongated thermal flex elements each have a double cup cross-section.

[0015] Another aspect of the present disclosure includes any of the above aspects, wherein the impingement cooling structure and the airfoil body are integral at their respective first ends.

[0016] Another aspect of the present disclosure includes any of the above aspects, further comprising a curved thermal flex connector coupling the first ends of each of the impingement cooling structure and the airfoil body.

[0017] Another aspect of the present disclosure includes a gas turbine (GT) system including a plurality of nozzles or blades, at least one nozzle or blade comprising: an airfoil body defined by concave pressure side and convex suction side outer walls connecting along leading and trailing edges and forming a radially extending chamber therebetween, the airfoil body having an inner surface facing the radially extending chamber; and an impingement cooling structure within the radially extending chamber, the impingement cooling structure including a wall spaced from the inner surface of the airfoil body, a plurality of holes defined through the wall, and a plurality of elongated thermal flex elements defined in the wall, wherein the airfoil body and impingement cooling structure include multiple layers of unitary material.

[0018] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the plurality of elongated thermal flex elements extend in a direction perpendicular to the radial length of the impingement cooling structure.

[0019] Another aspect of the present disclosure includes any of the above aspects, wherein the plurality of elongated thermal flex elements extend in a direction at an angle ranging from 30° to 60° relative to a radial length of the impingement cooling structure.

[0020] Another aspect of the present disclosure includes any of the above aspects, wherein the plurality of elongated thermal flex elements extend in a direction at an angle of approximately 45 degrees relative to the radial length of the impingement cooling structure.

[0021] Another aspect of the present disclosure includes any of the above aspects, further comprising a plurality of support members spacing the wall from the inner surface of the airfoil body, the plurality of support members being disposed between the plurality of elongated thermal flex elements.

[0022] Another aspect of the present disclosure includes any of the above aspects, wherein the plurality of elongated thermal flex elements includes a plurality of elongated thermal flex elements between adjacent rows of the plurality of support members.

[0023] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the plurality of elongated thermal flex elements each have one of a C-shaped cross-section, a symmetrical V-shaped cross-section, an asymmetrical V-shaped cross-section, a rounded corner U-shaped cross-section, a squared corner U-shaped cross-section, and a double cup-shaped cross-section.

[0024] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the impingement cooling structure and the airfoil body are integral at their respective first ends, and a curved thermal flex connector couples the impingement cooling structure and the airfoil body to each other.

[0025] Another aspect of the present disclosure includes a method of forming a turbine nozzle or blade comprising additively manufacturing a turbine nozzle or blade comprising: an airfoil body defined by concave pressure side outer wall and convex suction side outer wall connecting along leading and trailing edges and forming a radially extending chamber therebetween, the airfoil body having an inner surface facing the radially extending chamber; and an impingement cooling structure within the radially extending chamber, the impingement cooling structure including a wall spaced from the inner surface of the airfoil body, a plurality of holes defined through the wall, and a plurality of elongated thermal flex elements defined in the wall, wherein the airfoil body and impingement cooling structure include a plurality of layers of a unitary material.

[0026] Two or more aspects described in this disclosure, including those described in the Summary section, may be combined to form an embodiment not specifically described herein.

[0027] The details of one or more embodiments are set forth in the accompanying drawings and the description below. Other features, objects, and advantages will become apparent from the description and drawings, and from the claims.

[0028] These and other features of the present disclosure will be more readily understood from the following detailed description of the various aspects of the disclosure, taken in conjunction with the accompanying drawings which illustrate various embodiments of the present disclosure. [Brief explanation of the drawings]

[0029] [Figure 1] 1 illustrates a schematic cross-sectional view of an exemplary turbomachine in the form of a gas turbine system. [Figure 2] 1 illustrates a cross-sectional view of a portion of an exemplary turbine section, according to an embodiment of the present disclosure. [Figure 3] 1 illustrates a perspective view of an exemplary turbine nozzle including an impingement cooling structure in accordance with an embodiment of the present disclosure. [Figure 4] 1 illustrates a perspective view of an exemplary turbine blade including an impingement cooling structure according to an embodiment of the present disclosure. [Figure 5] 1 illustrates a partial cross-sectional view of a turbine nozzle or blade according to an embodiment of the present disclosure. [Figure 6] 2 illustrates a cross-sectional view of a portion of an airfoil body and impingement cooling structure according to an embodiment of the present disclosure. [Figure 7] FIG. 2 illustrates an enlarged cross-sectional view of a portion of an airfoil body and impingement cooling structure according to an embodiment of the present disclosure. [Figure 8] 1 illustrates an internal view of a portion of an impingement cooling structure according to an embodiment of the present disclosure. [Figure 9] 2 illustrates a cross-sectional view of a portion of an airfoil body and impingement cooling structure according to an embodiment of the present disclosure. [Figure 10] 4 illustrates an internal view of a portion of an impingement cooling structure according to another embodiment of the present disclosure. [Figure 11] 1 illustrates a cross-sectional perspective view of a portion of an impingement cooling structure according to an alternative embodiment of the present disclosure. [Figure 12A] 1A-1C show diagrams of cross-sectional shapes of flex elements according to embodiments of the present disclosure. [Figure 12B] 1A-1C show diagrams of cross-sectional shapes of flex elements according to embodiments of the present disclosure. [Figure 12C] 1A-1C show diagrams of cross-sectional shapes of flex elements according to embodiments of the present disclosure. [Figure 13]10A-10C show cross-sectional views of flex elements according to other embodiments of the present disclosure. [Figure 14] 10A-10C show diagrams of cross-sectional shapes of flex elements according to further embodiments of the present disclosure; [Figure 15] 10A-10C show perspective views of cross-sectional shapes of flex elements according to other embodiments of the present disclosure. [Figure 16] 10A-10C show cross-sectional views of flex elements according to yet another embodiment of the present disclosure. [Figure 17] FIG. 1 is a schematic diagram of an additive manufacturing process including a non-transitory computer readable storage medium storing code representing a turbine nozzle or blade according to an embodiment of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION

[0030] It should be noted that the drawings of the present disclosure are not necessarily to scale. The drawings are intended to depict only typical aspects of the disclosure and therefore should not be considered limiting of the scope of the disclosure. In the drawings, like reference numerals represent like elements between the drawings.

[0031] As an initial matter, in order to clearly explain the present disclosure, it is necessary to select specific terminology when referring to and describing relevant machine components in an exemplary turbomachinery application. In doing so, common industry terminology will be used whenever possible and will be employed in a manner consistent with its received meaning. Unless otherwise indicated, such terminology should be given a broad interpretation consistent with the context of this application and the appended claims. Those skilled in the art will understand that in many cases, a particular component may be referred to using several different or overlapping terms. What may be described herein as a single component may also include, and be referred to in other contexts, as consisting of multiple components. Alternatively, what may be described herein as including multiple components may be referred to elsewhere as a single component.

[0032] Additionally, several descriptive terms may be used regularly herein, and it will prove useful to define these terms at the beginning of this section. These terms and their definitions are as follows, unless otherwise stated: As used herein, "downstream" and "upstream" are terms that indicate a direction relative to the flow of a working fluid through a turbomachine, or a fluid, such as, for example, the flow of air through a combustor, or a coolant through one of the component systems of a turbomachine. The term "downstream" corresponds to the direction of fluid flow, and the term "upstream" refers to the direction opposite to the flow. The terms "forward" and "aft," unless otherwise specified, refer to directions, with "forward" referring to the forward or compressor end of the turbomachine and "aft" referring to the aft or turbine end of the turbomachine.

[0033] It is often necessary to describe components at different radial locations relative to a central axis. The term "axial" refers to movement or location parallel to an axis, e.g., the axis of a turbomachine. The term "radial" refers to movement or location perpendicular to an axis, e.g., the axis of a turbomachine. In such cases, if a first component is located closer to the axis than a second component, the first component may be described herein as being "radially inward" or "inward" of the second component. Conversely, if a first component is located farther from the axis than the second component, the first component may be described herein as being "radially outward" or "outward" of the second component. Finally, the term "circumferential" refers to movement or location on the circumferential inner surface of a casing extending around a shaft, e.g., the axis of a turbomachine. As noted above, it will be understood that such terms may be applied with respect to the axis of a turbomachine.

[0034] Additionally, as described below, certain descriptive terms may be used conventionally herein: the terms "first," "second," and "first" may be used interchangeably to distinguish one component from another, and are not intended to denote the location or importance of the individual components.

[0035] The terms used herein are for the purpose of describing particular embodiments only and are not intended to limit the disclosure. As used herein, the singular forms "a," "an," and "the" are intended to include the plural forms unless the context clearly dictates otherwise. It will be further understood that the terms "comprises" and / or "comprising," as used herein, specify the presence of stated features, integers, steps, operations, elements, and / or components, but do not exclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and / or groups thereof. "Optional" or "optionally" means that the subsequently described event may or may not occur, or that the subsequently described feature may or may not be present, and that the description includes instances where the event occurs or the feature is present, as well as instances where the event does not occur or is not present.

[0036] When an element or layer is referred to as being "on," "engaged," "connected," "coupled," or "attached" to another element or layer, it may be directly on, engaged, connected, coupled, or attached to the other element or layer, or there may be intervening elements or layers. Conversely, when an element is referred to as being "directly on," "directly engaged," "directly connected," or "directly coupled" to another element or layer, there are no intervening elements or layers. Other words used to describe relationships between elements should be interpreted in a similar manner (e.g., "between" and "directly between," "adjacent to" and "directly adjacent to," etc.). As used herein, the term "and / or" includes any and all combinations of one or more of the associated listed items. The verb forms of "couple" and "attach" may be used interchangeably herein.

[0037] As described above, the present disclosure provides a turbine nozzle or blade having thermal compliance features. The nozzle or blade may include an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and form a radially extending chamber therebetween for receiving a coolant flow. The airfoil body has an inner surface facing the radially extending chamber. An impingement cooling structure is disposed within the radially extending chamber. The impingement cooling structure includes a wall spaced from the inner surface of the airfoil body, a plurality of holes defined through the wall, and a plurality of elongated thermal flex elements defined in the wall. Because the nozzle or blade is fabricated by additive manufacturing, the airfoil body and impingement cooling structure include multiple monolithic layers of material. The elongated thermal flex elements provide thermal compliance to the integrally formed airfoil body and impingement cooling structure. In particular, the flex elements significantly reduce thermal distortion of components exposed to large thermal differentials. Thus, the nozzle or blade can be shaped to be robust against high cycle fatigue (HCF) and flexible against low cycle fatigue (LCF). The flex element also enables cost-effective additive manufacturing regardless of the temperature gradients the turbine nozzle or blade is expected to be exposed to during use.

[0038] Referring to the drawings, Figure 1 is a cross-sectional view of an exemplary machine including a turbine to which the teachings of the present disclosure may be applied. Shown in Figure 1 is a turbomachine 90 in the form of a combustion turbine or gas turbine (GT) system 100 (hereinafter "GT system 100"). GT system 100 includes a compressor 102 and a combustor 104. Combustor 104 includes a combustion region 105 and a fuel nozzle section 106. GT system 100 also includes a turbine 108 (i.e., expansion turbine) and a common compressor / turbine shaft 110 (hereinafter "rotor 110").

[0039] The GT system 100 may be, for example, a 7HA.03 engine commercially available from General Electric Company of Greenville, South Carolina. The present disclosure is not limited to any particular GT system and may be practiced in connection with other engines, including, for example, other General Electric HA, F, B, LM, GT, TM, and E-class engine models, as well as engine models from other manufacturers. More importantly, the teachings of the present disclosure are not necessarily applicable only to the turbine section of a GT system, but may be applicable to virtually any type of industrial or other turbomachinery, such as steam turbines, jet engines, compressors (as in FIG. 1 ), turbofans, turbochargers, etc. Accordingly, references to the turbine 108 of the GT system 100 are merely illustrative and not limiting.

[0040] FIG. 2 shows a cross-sectional view of an exemplary portion of a turbine 108. In the illustrated example, the turbine 108 includes four stages L0-L3 that may be used with the GT system 100 of FIG. 1. The four stages are designated L0, L1, L2, and L3. Stage L0 is the first stage and is the smallest (radially) of the four stages. Stage L1 is the second stage and is located axially adjacent to the first stage L0. Stage L2 is the third stage and is located axially adjacent to the second stage L1. Stage L3 is the fourth and final stage and is the largest (radially). It should be understood that four stages are shown by way of example only, and each turbine may have more or fewer than four stages.

[0041] A plurality of stationary turbine vanes or nozzles 112 (hereinafter “nozzles 112”) may cooperate with a plurality of rotating turbine blades 114 (hereinafter “blades 114”) to form each stage L0-L3 of the turbine 108 and define a portion of a working fluid path through the turbine 108. The blades 114 of each stage are coupled to the rotor 110 ( FIG. 1 ), for example, by a respective rotor wheel 116 that circumferentially couples them to the rotor 110 ( FIG. 1 ). That is, the blades 114 are circumferentially spaced apart and mechanically coupled to the rotor 110, for example, by the rotor wheels 116. The stationary nozzle section 115 is mounted to the casing 124 and includes a plurality of stationary nozzles 112 circumferentially spaced about the rotor 110 ( FIG. 1 ). It will be appreciated that the blades 114 rotate with the rotor 110 ( FIG. 1 ) and are therefore subject to centrifugal forces, while the nozzles 112 are stationary.

[0042] 1 and 2 , during operation, air flows through the compressor 102, and the compressed air is channeled to the combustor 104. The compressed air is supplied to a fuel nozzle section 106 integral with the combustor 104. The fuel nozzle section 106 is in fluid communication with a combustion zone 105. The fuel nozzle section 106 is also in fluid communication with a fuel source (not shown in FIG. 1 ) and channels fuel and air to the combustion zone 105. The combustor 104 ignites and combusts the fuel to generate combustion gases. The combustor 104 is in fluid communication with a turbine 108, in which thermal energy from the gas stream is converted to mechanical rotational energy by channeling the combusted fuel (e.g., working fluid) through a working fluid passage to rotate blades 114. The turbine 108 is rotatably coupled to and drives a rotor 110. The compressor 102 is rotatably coupled to the rotor 110. At least one end of the rotor 110 may extend axially away from the compressor 102 or the turbine 108 and may be attached to a load or machine (not shown), such as, but not limited to, an electrical generator, a load compressor, and / or another turbine.

[0043] 3 and 4 show perspective views of a (stationary) nozzle 112 and a (rotating) blade 114, respectively, of the type in which an embodiment of the impingement cooling structure 120 of the present disclosure may be employed.

[0044] 3 and 4, each nozzle or blade 112, 114 includes an airfoil 128 having a base end 130, a tip end 132, and an airfoil body 134 extending between the base end 130 and the tip end 132. As shown in FIG. 3, the nozzle 112 includes an outer endwall 136 at the base end 130 and an inner endwall 138 at the tip end 132. The outer endwall 136 couples to the casing 124 (FIG. 2). As shown in FIG. 4, the blade 114 includes a dovetail 140 at the base end 130, thereby attaching the blade 114 to the rotor wheel 116 (FIG. 2) of the rotor 110 (FIG. 2). The base end 130 of the blade 114 may further include a shank 142 extending between the dovetail 140 and a platform 146. A platform 146 is disposed at the junction of the airfoil body 134 and the shank 142 and defines a portion of the inner boundary of the working fluid passage (FIG. 2) through the turbine 108 .

[0045] It will be appreciated that the airfoil bodies 134 of the nozzles 112 and blades 114 are the active components of the nozzles 112 or blades 114 that interrupt the flow of working fluid and, in the case of the blades 114, induce the rotor 110 ( FIG. 1 ) to rotate. The airfoil bodies 134 of the nozzles 112 and blades 114 are seen to include a concave pressure side (PS) outer wall 150 and a circumferentially or laterally opposed convex suction side (SS) outer wall 152, each extending axially between opposite leading and trailing edges 154, 156. The walls 150 and 152 also extend radially from the base end 130 (i.e., the outer endwall 136 of the nozzle 112 and the platform 146 of the blades 114) to the tip end 132 (i.e., the inner endwall 138 of the nozzle 112 and the tip 158 of the blades 114). Walls 150, 152 define a radially extending chamber 160 therebetween, for example, to receive a flow of coolant. As shown in the partial cross-sectional view of Figure 5, airfoil body 134 has an inner surface 162 facing radially extending chamber 160. Coolant may be supplied to radially extending chamber 160 from any now known or later developed source, for example, air from compressor 102.

[0046] In the illustrated example, blade 114 does not include a tip shroud, however, it should be noted that the teachings of this disclosure are equally applicable to blades that include a tip shroud at tip 158. The nozzle 112 and blade 114 shown in Figures 3-4 are exemplary only, and the teachings of this disclosure may be applied to a wide variety of nozzles and blades.

[0047] In accordance with an embodiment of the present disclosure, FIG. 6 illustrates a cross-sectional view of a portion of the airfoil body 134 and the impingement cooling structure 170, FIG. 7 illustrates an enlarged cross-sectional view of a portion of the airfoil body 134 and the impingement cooling structure 170, and FIG. 8 illustrates an internal view of a portion of the impingement cooling structure 170. Referring to FIGS. 5-8 , the nozzle 112 or blade 114 also includes an impingement cooling structure 170 within the radially extending chamber 160. The impingement cooling structure 170 is an integral internal structure formed integrally with the airfoil body 134. More specifically, the airfoil body 134 and the impingement cooling structure 170 are formed together using additive manufacturing to include multiple layers of a single material.

[0048] The impingement cooling structure 170 (hereinafter “structure 170”) includes a wall 172 spaced from the inner surface 162 of the airfoil body 134. A plurality of holes 174 are defined through the wall 172 such that coolant 176 ( FIG. 7 ) supplied to the radially extending chamber 160 can pass through the holes 174 to cool the inner surface 162 of the airfoil body 134. The wall 172 is spaced from the inner surface 162 of the airfoil body 134 to define a post-impingement cavity between the wall 172 and the inner surface 162. The wall 172 is a unitary wall structure, i.e., one piece. Furthermore, the impingement cooling structure 170 and the airfoil body 134 are integral at their respective first ends 184, 186. The impingement cooling structure 170 may also be integral with the airfoil body 134 at their respective radially inner second ends (not shown).

[0049] The spacing S between the wall 172 of the structure 170 and the inner surface 162 of the airfoil body 134 may be user-defined to ensure desired cooling. A plurality of support members 180 may be provided to space the wall 172 from the inner surface 162 of the airfoil body 134. The support members 180 may be, for example, structural columns that can hold the wall 172 in a desired position. The support members 180 may be arranged in rows. In another example, the support members 180 may each be a structural rib that can hold the wall 172 in a desired position. In this case, the support members 180 may be generally parallel to the thermal flex element 190.

[0050] The structure 170 also includes a plurality of elongated thermal flex elements 190 defined in the wall 172. As best seen in FIGS. 6 and 7 , the plurality of elongated thermal flex elements 190 (hereinafter “flex elements 190”) are not solid ribs or supports extending from the surface of the structure 170, but rather hollow curves in the normally flat or sheet-like surface of the wall 172. The flex elements 190 have opposing surfaces 192, 194. Surface 192 faces the radially extending chamber 160, and surface 194 faces the inner surface 162 of the airfoil body 134. The opposing surfaces 192, 194 of the flex element 190 are generally parallel. That is, the opposing surfaces 192, 194 are as parallel as possible using a suitable additive manufacturing process, with slight tolerances for the desired stiffness and / or flexibility of the flex element 190 relative to the remainder of the wall 172. Flex elements 190 extend or protrude inward toward the radially extending chamber 160. As shown in FIG. 7 (and in phantom in FIGS. 6 and 8 ), multiple support members 180 are disposed between the flex elements 190. The flex elements 190 are referred to as “elongated” because they have a generally linear extent with respect to the interior of the wall 172 that is greater than their radial extent (relative to the radial length of the nozzle 112 or blade 114). The impingement cooling holes 174 may be disposed in any manner between adjacent flex elements 190 to accommodate the desired cooling of the interior surface 162 and the location of the flex elements 190 and / or support members 180.

[0051] The flex element 190 provides thermal compliance to the integrally formed airfoil body 134 and impingement cooling structure 170. More specifically, the flex element 190 significantly reduces thermal distortion of the components when exposed to large thermal differentials between the hot combustion gases and the impingement coolant (e.g., the coolant 176). Thus, the nozzle 112 or blade 114 can be shaped to be robust against high-cycle fatigue (HCF) and flexible against low-cycle fatigue (LCF). The flex element 190 also enables cost-effective additive manufacturing regardless of the thermal gradients to which the turbine nozzle 112 or blade 114 is expected to be exposed during use. The flex element 190 also allows for the maintenance of normal hole 174 spacing, preventing fracture of the support member 180 even when thermal gradients become large.

[0052] The arrangement and shape of flex elements 190 can take a variety of forms depending on the embodiment of the present disclosure. In Figures 6-8, flex elements 190 extend in a direction perpendicular to the radial length L of structure 170, i.e., approximately 90° relative to radial length L. Furthermore, flex elements 190 have a symmetrical trapezoidal or open C-shaped cross-section in Figures 6-8. The C-shaped cross-section can vary from a perfect partial circular arrangement, as shown in Figure 7, for example.

[0053] FIG. 9 illustrates a cross-sectional view of a portion of the airfoil body 134 and structure 170, and FIG. 10 illustrates an internal view of a portion of the structure 170, according to another embodiment of the present disclosure. In FIGS. 9-10, the flex element 190 extends at an angle α ranging from 30° to 60° relative to the radial length L of the structure 170 (and possibly the radial length L of the turbine nozzle 112 or blade 114). In certain embodiments, the flex element 190 may extend at an angle α of approximately 45° relative to the radial length L of the structure 170. The flex element 190 may have an angled surface (angle α) joined to a short mid-surface parallel to the wall 172. The mid-surface is coupled to a short connecting surface opposite the angled surface and joined to the wall 172. Such a flex element 190 has an asymmetric cross-sectional shape.

[0054] 11 illustrates a cross-sectional perspective view of a portion of structure 170 according to an alternative embodiment. In FIG. 11, a plurality of elongated thermal flex elements 190 are included between rows of a plurality of support members 180. Any number of flex elements 190 may be disposed along the radial length of wall 172, and any number of flex elements 190 may be disposed between rows of support members 180.

[0055] 11 also illustrates an optional embodiment in which a curved thermal flex connector 200 couples the respective first ends 184, 186 of the airfoil body 134 and the structure 170. The curved thermal flex connector 200 can have any shape desirable to provide additional thermal flexibility between the structure 170 and the airfoil body 134.

[0056] Regarding the shape of the flex element 190, FIGS. 12-16 show illustrations of various exemplary cross-sectional shapes of the flex element 190, according to embodiments of the present disclosure. As previously described and shown in FIGS. 6-11, the flex element 190 can have a trapezoidal or C-shaped cross-section, which can be symmetrical or asymmetrical. The length and / or angle of the C-shaped cross-section can be user-defined to provide the desired flexibility. In FIGS. 12A-12C, the flex element 190 has a generally V-shaped cross-section. The apex of the V-shape can be located anywhere desired to provide the necessary stress relief. FIG. 12A shows the apex 198 of the V-shape approximately centered with respect to the side of the V-shape (e.g., a symmetric V-shaped cross-section), FIG. 12B shows the apex 198 to one side (upward in the figure) of the center position with respect to the side of the V-shape, and FIG. 12C shows the apex 198 to one side (downward in the figure) of the center position with respect to the side of the V-shape. 12B-12C therefore show an asymmetric V-shaped cross section.

[0057] In FIG. 13, the flex element 190 has a rounded U-shaped cross-section. In FIG. 14, the flex element 190 has a squared U-shaped cross-section, e.g., with approximately 90° corners. In FIG. 15, the flex element 190 has a sloped U-shaped cross-section, e.g., with one rounded corner and one approximately 90° corner (along the radially inner surface of the wall 170). In FIG. 16, the flex element 190 has a double cup-shaped cross-section. More specifically, a "double cup-shaped cross-section" can include a U-shaped or C-shaped cross-section 210 set into a curved portion 212 of another U-shaped portion 214.

[0058] Although particular cross-sectional shapes are shown separately in the figures, the different embodiments can be used together in any desired manner, i.e., with different flex elements 190 having different cross-sectional shapes on the same turbine nozzle 112 or blade 114. Additionally, although particular cross-sectional shapes are shown, the flex elements 190 can have any cross-sectional shape formable using additive manufacturing.

[0059] Although flex element 190 is illustrated as being generally straight in the drawings, it may also have some degree of curvature.

[0060] Embodiments of the present disclosure may also include a GT system 100 including a plurality of nozzles 112 or blades 114, at least one of the nozzles or blades including an airfoil body 134 and structure 170 therein, as previously described.

[0061] Regardless of location and / or cross-sectional shape, during operation of the GT system 100, the structure 170 provides the necessary impingement cooling to the inner surface 162 of the airfoil body 134 by means of the flex elements 190 that provide sufficient thermal expansion and contraction to reduce stress, and the impingement holes 174 located between the flex elements 190, as described herein.

[0062] The nozzle 112 or blade 114 may include any metal or metal compound capable of withstanding the environment in which it will be used. The nozzle 112 or blade 114 may be advantageously made using additive manufacturing. Through additive manufacturing, the airfoil body 134 and impingement cooling structure 170 may be formed to include multiple monolithic layers of material. The multiple monolithic layers of material may also include multiple internal supports 180. Thus, a method of forming the turbine nozzle 112 or blade 114 may include additively manufacturing the turbine nozzle 112 or blade 114 to include the airfoil body 134 and structure 170, as described above.

[0063] As used herein, additive manufacturing (AM) may include any process that produces an object by successive layering of material rather than by removal of material as in traditional processes. Additive manufacturing can create complex geometric shapes without the use of tools, molds, or fixtures of any kind, and with little or no waste material. Unlike machining a component from a solid billet of metal, where most of the material is cut away and discarded, additive manufacturing uses only the material needed to form the part. Additive manufacturing processes may include, but are not limited to, 3D printing, rapid prototyping (RP), direct digital manufacturing (DDM), binder jetting, selective laser melting (SLM), and direct metal laser melting (DMLM). In the current context, DMLM has proven advantageous.

[0064] To illustrate an example additive manufacturing process, FIG. 17 shows a schematic / block diagram of an exemplary computerized additive manufacturing system 900 for generating an object 902. In this example, the system 900 is configured for DMLM. It is understood that the overall teachings of this disclosure are equally applicable to other forms of additive manufacturing. While the object 902 is illustrated as a blade 114 as described herein, the system is equally applicable to a nozzle 112. The AM system 900 generally includes a computerized additive manufacturing (AM) control system 904 and an AM printer 906. The AM system 900 executes code 920, including a set of computer-executable instructions defining the nozzle 112 or blade 114, as described below, and physically generates the object using the AM printer 906. Each AM process can use different raw materials in the form of, for example, a fine powder, a liquid (e.g., polymer), a sheet, etc., the stock of which can be held in a chamber 910 of the AM printer 906. In this case, the nozzle 112 or blade 114 can be made of a metal or metal compound.

[0065] As shown, the applicator 912 can create a thin layer of raw material 914, which is laid out as a blank canvas from which each successive slice of the final object 902 is created. In other cases, the applicator 912 can apply or print the next layer directly on top of the previous layer, as defined by code 920, for example, if the material is a polymer or if a metal binder jetting process is used. In the illustrated example, the laser or electron beam 916 melts particles for each slice, as defined by code 920, although this may not be necessary if a fast-setting liquid plastic / liquid polymer is employed. Various parts of the AM printer 906 can move to accommodate the addition of each new layer; for example, after each layer, the build platform 918 can be lowered and / or the chamber 910 and / or applicator 912 can be raised.

[0066] The AM control system 904 is shown implemented as computer program code on a computer 930. In this regard, the computer 930 is shown to include a memory 932, a processor (PU) 934, an input / output (I / O) interface 936, and a bus 938. Additionally, the computer 930 is shown in communication with external I / O devices / resources 940 and a storage system 942. Generally, the processor 934 executes computer program code, such as the AM control system 904, stored in the memory 932 and / or the storage system 942 under instructions from code 920 representing the nozzles 112 or blades 114 described herein. When executing the computer program code, the processor 934 can read and / or write data from the memory 932, the storage system 942, the I / O device 940, and / or the AM printer 906. The bus 938 provides a communication link between each of the components of the computer 930, and the I / O devices 940 can include any device that allows a user to interact with the computer 930 (e.g., a keyboard, pointing device, display, touch screen, etc.).

[0067] Computer 930 is merely representative of various possible combinations of hardware and software. For example, processor 934 may comprise a single processing unit or may be distributed across one or more processing units in one or more locations, e.g., a client and a server. Similarly, memory 932 and / or storage system 942 may reside in one or more physical locations. Memory 932 and / or storage system 942 may comprise any combination of various types of non-transitory computer-readable storage media, including magnetic media, optical media, random access memory (RAM), read-only memory (ROM), etc. Computer 930 may comprise any type of computing device, such as a network server, desktop computer, laptop, handheld device, mobile smartphone, personal data assistant, etc.

[0068] The additive manufacturing process begins with a non-transitory computer-readable storage medium (e.g., memory 932, storage system 942, etc.) storing code 920 representing the nozzle 112 or blade 114. As described above, the code 920 includes a set of computer-executable instructions defining the nozzle 112 or blade 114 that, upon execution of the code by the system 900, can be used to, among other things, physically generate the corrugated surface(s) of the impingement wall 172. For example, the code 920 may include a precisely defined 3D model of the nozzle 112 or blade 114 and can be generated from any of a wide variety of well-known computer-aided design (CAD) software systems, such as AutoCAD®, TurboCAD®, DesignCAD 3D Max, etc. In this regard, the code 920 can take any now known or later developed file format. For example, code 920 may be a Standard Tessellation Language (STL) created for a 3D system's stereolithography CAD program, or it may be an Additive Manufacturing File (AMF), which is an American Society of Mechanical Engineers (ASME) standard that is an Extensible Markup Language (XML)-based format designed to allow any CAD software to represent the shape and configuration of any three-dimensional object to be produced on an AM printer. Code 920 may be converted between different formats, converted into a set of data signals, transmitted, received as a set of data signals, converted into code, and stored, for example, as needed.

[0069] The code 920 may be input to the system 900 and may come from a part designer, an intellectual property (IP) provider, a design firm, the operator or owner of the system 900, or from other sources. In either case, the AM control system 904 executes the code 920 and divides the nozzle 112 or blade 114 into a series of thin slices, which are assembled with successive layers of liquid, powder, sheet, or other material using the AM printer 906. In the DMLM example, each layer is melted to the precise geometry defined by the code 920 and fused to the previous layer. The nozzle 112 or blade 114 may then be subjected to any of a variety of finishing processes, such as, for example, light machining, sealing, polishing, assembly to other portions of the blade, etc.

[0070] Embodiments of the present disclosure provide various technical and commercial advantages, examples of which are described herein. As described above, the elongated thermal flex elements provide thermal compliance to the integrally formed airfoil body and impingement cooling structure. More specifically, the flex elements significantly reduce thermal distortion of components exposed to large thermal differentials. As such, the nozzle or blade can be shaped to be robust against high cycle fatigue (HCF) and flexible against low cycle fatigue (LCF). The flex elements also enable cost-effective additive manufacturing regardless of the thermal gradients that the turbine nozzle or blade is expected to experience during use.

[0071] As used herein throughout this specification and claims, approximation language may be applied to modify any quantitative expression that may vary within acceptable limits without resulting in a change in the relevant basic function. Thus, values ​​modified by terms such as "about," "approximately," and "substantially" are not limited to the exact value specified. In at least some instances, approximation language may correspond to the precision of the instrument used to measure the value. Here, and throughout this specification and claims, range limitations are combinable and / or interchangeable, and unless the context or language dictates otherwise, such ranges are identified and include all subranges encompassed therein. "Approximately" or "about," as applied to specific values ​​in a range, may apply to both endpoints and indicate + / - 5% of the stated value(s), unless specifically dependent on the precision of the instrument used to measure the value.

[0072] The corresponding structure, material, operations, and equivalents of all means-plus-function or step-plus-function elements in the following claims are intended to include any structure, material, or operations for performing the function as specifically claimed in combination with other claimed elements. The description of the present disclosure has been presented for purposes of illustration and description, but is not intended to be exhaustive or to limit the disclosure to the form disclosed. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of the disclosure. The embodiments were chosen and described to best explain the principles and practical applications of the disclosure and to enable others skilled in the art to understand the disclosure in various embodiments with various modifications suitable for the particular use contemplated.

Claims

1. A turbine nozzle (112) or blade (114), an airfoil body (134) defined by a concave pressure side outer wall (150) and a convex suction side outer wall (152) connecting along a leading edge (154) and a trailing edge (156) and forming a radially extending chamber (160) therebetween, said airfoil body (134) having an inner surface (162) facing said radially extending chamber (160); an impingement cooling structure (170) within the radially extending chamber (160), the impingement cooling structure (170) comprising: a wall (172) spaced from the inner surface (162) of the airfoil body (134); a plurality of holes (174) defined through said wall (172); a plurality of elongated thermal flex elements (190) defined in said wall (172); an impingement cooling structure (170) including: Equipped with A turbine nozzle (112) or blade (114), wherein the airfoil body (134) and the impingement cooling structure (170) comprise multiple monolithic layers of material.

2. The turbine nozzle (112) or blade (114) of claim 1, wherein the plurality of elongated thermal flex elements (190) extend in a direction perpendicular to a radial length of the impingement cooling structure (170).

3. 2. The turbine nozzle or blade of claim 1, wherein the plurality of elongated thermal flex elements extend in a direction at an angle in a range of 30° to 60° relative to a radial length of the impingement cooling structure.

4. 4. The turbine nozzle or blade of claim 3, wherein the plurality of elongated thermal flex elements extend in a direction that is at an angle of approximately 45 degrees relative to the radial length of the impingement cooling structure.

5. 2. The turbine nozzle or blade of claim 1, further comprising a plurality of support members spacing the wall from the inner surface of the airfoil body, the plurality of support members being disposed between the plurality of elongated thermal flex elements.

6. 6. The turbine nozzle or blade of claim 5, wherein the plurality of elongated thermal flex elements comprises a plurality of elongated thermal flex elements between adjacent rows of the plurality of support members.

7. The turbine nozzle (112) or blade (114) of claim 1, wherein each of the plurality of elongated thermal flex elements (190) has a C-shaped cross section.

8. 2. The turbine nozzle or blade of claim 1, wherein each of the plurality of elongated thermal flex elements has one of a symmetrical V-shaped cross-section, an asymmetrical V-shaped cross-section, a rounded U-shaped cross-section, and a squared U-shaped cross-section.

9. The turbine nozzle (112) or blade (114) of claim 1, wherein each of the plurality of elongated thermal flex elements (190) has a double cup cross section.

10. The turbine nozzle (112) or blade (114) of claim 1, wherein the impingement cooling structure (170) and the airfoil body (134) are integral at their respective first ends (184, 186).

11. 11. The turbine nozzle or blade of claim 10, further comprising a curved thermal flex connector coupling the impingement cooling structure and the respective first ends of the airfoil body.

12. A gas turbine (GT) system (100) including a plurality of nozzles (112) or blades (114), at least one nozzle (112) or blade (114) comprising: an airfoil body (134) defined by a concave pressure side outer wall (150) and a convex suction side outer wall (152) connecting along a leading edge (154) and a trailing edge (156) and forming a radially extending chamber (160) therebetween, said airfoil body (134) having an inner surface (162) facing said radially extending chamber (160); an impingement cooling structure (170) within the radially extending chamber (160), the impingement cooling structure (170) comprising: a wall (172) spaced from the inner surface (162) of the airfoil body (134); a plurality of holes (174) defined through said wall (172); a plurality of elongated thermal flex elements (190) defined in said wall (172); an impingement cooling structure (170) including: Equipped with A gas turbine (GT) system (100), wherein the airfoil body (134) and the impingement cooling structure (170) comprise multiple monolithic layers of material.

13. The GT system (100) of claim 12, wherein the plurality of elongated thermal flex elements (190) extend in a direction perpendicular to a radial length of the impingement cooling structure (170).

14. The GT system (100) of claim 12, wherein the plurality of elongated thermal flex elements (190) extend in a direction at an angle in a range of 30° to 60° relative to a radial length of the impingement cooling structure (170).

15. The GT system (100) of claim 14, wherein the plurality of elongated thermal flex elements (190) extend in a direction that is at an angle of approximately 45 degrees relative to the radial length of the impingement cooling structure (170).

16. 13. The GT system of claim 12, further comprising a plurality of support members spacing the wall from the inner surface of the airfoil body, the plurality of support members being disposed between the plurality of elongated thermal flex elements.

17. The GT system (100) of claim 16, wherein the plurality of elongated thermal flex elements (190) comprises a plurality of elongated thermal flex elements (190) between adjacent rows of the plurality of support members (180).

18. 13. The GT system of claim 12, wherein each of the plurality of elongated thermal flex elements has one of a C-shaped cross section, a symmetrical V-shaped cross section, an asymmetrical V-shaped cross section, a rounded corner U-shaped cross section, a squared corner U-shaped cross section, and a double cup-shaped cross section.

19. 13. The GT system of claim 12, wherein the impingement cooling structure and the airfoil body are integral at their respective first ends, and a curved thermal flex connector couples the impingement cooling structure and the respective first ends of the airfoil body.

20. A method of forming a turbine nozzle (112) or blade (114), comprising: an airfoil body (134) defined by a concave pressure side outer wall (150) and a convex suction side outer wall (152) connecting along a leading edge (154) and a trailing edge (156) and forming a radially extending chamber (160) therebetween, said airfoil body (134) having an inner surface (162) facing said radially extending chamber (160); an impingement cooling structure (170) within the radially extending chamber (160), the impingement cooling structure (170) comprising: a wall (172) spaced from the inner surface (162) of the airfoil body (134); a plurality of holes (174) defined through said wall (172); a plurality of elongated thermal flex elements (190) defined in said wall (172); an impingement cooling structure (170) including: Equipped with the airfoil body (134) and the impingement cooling structure (170) include a plurality of monolithic layers of material; Additively manufacturing said turbine nozzle (112) or blade (114). A method of forming a turbine nozzle (112) or blade (114), comprising:

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