Additively manufactured turbine rotor blade root with hollow mounting frame with grid support structure

Abstract

The present invention provides a turbine rotor blade root (144) that is additively manufactured and includes a shank (148, 148A) having a radially extending chamber (134) defined therein. A blade mount (146) is located at a radially inner end of the shank. The blade mount (146) has a hollow interior (330) defined therein, wherein the hollow interior (330) is in fluid communication with the radially extending chamber (134). A grid support structure (340) is disposed within the hollow interior (330) of the blade mount (146).

Description

Background Art

[0001] The present disclosure relates generally to turbines, and more particularly, to turbine rotor blades having multiple additively manufactured integral features, such as integral impingement sleeves, non-linear cooling channels in the platform, angel wings with coolant delivery channels, and hollow dovetails with integral grid support structures.

[0002] A turbine includes a plurality of turbine rotor blades coupled to a rotor. A working fluid such as steam or burning fuel is forced against the blades to force them to turn the rotor. The turbine rotor blades operate under extremely hot conditions and require cooling. Cooling features can be provided in a variety of ways.

[0003] One mechanism for providing cooling is an impact insert. An impact insert or sleeve includes a hollow body having cooling channels in its walls that allow coolant to be delivered through the cooling channels to impinge or impinge on the surface to be cooled. Impact inserts are used in various hot gas path (HGP) components such as turbine rotor blades in, for example, turbomachinery to improve the cooling performance of the cooling circuit therein. One problem faced by impact inserts is positioning the impact insert within a tapered or curved cavity in the HGP component in a manner close enough to achieve high cooling performance, but not so close that the cooling is ineffective. One indicator of the cooling performance of an impact insert is the Z / D parameter, which is the ratio of the standoff distance Z between the insert and the inner surface of the HGP component to the diameter D of the cooling channel (hole) in the impact insert. The Z / D parameter value of the insert is typically designed to be within a desired range that results in better cooling performance.

[0004] Impingement cooling is typically not provided by the insert in situations where the necessary standoff distance cannot be formed. For example, where the cavity curvature in the HGP component is so significant that the impingement insert cannot be made thin enough or curved to the corresponding necessary standoff distance, impingement cooling cannot be provided. One approach to addressing this problem is to provide the impingement insert in multiple flexible longitudinal sections to make it easier to insert them into the HGP component. However, having to position and couple multiple insert sections together or couple them to the HGP component in sequence increases manufacturing complexity, time, and cost. The flexible impingement insert section also does not provide adjacent elements around its perimeter (i.e., transverse (cross-section)), which can detract from its cooling performance at discontinuities.

[0005] Impingement cooling is applied in a limited manner to rotating turbine rotor blades in turbines, for example to their leading edges. However, impingement cooling has not been applied more widely to the entire inner surface of the turbine rotor blades because the centrifugal forces experienced by the rotating blades force the coolant to the radially outer tip of the blades as they rotate, making impingement cooling less effective.

[0006] Another cooling feature includes a cooling passage through a portion of a turbine rotor blade to be cooled. For example, a turbine rotor blade includes a platform that extends laterally to cooperate with a platform of an adjacent turbine rotor blade to form a portion of a working fluid path through the turbine. Due to the high temperature of the working fluid, the platform typically includes a cooling circuit therein that feeds a plurality of cooling passages that exit through an inclined surface of the platform. Some platforms include a damping pin seat in the inclined surface that receives an axially extending pin therein that cooperates with an adjacent damping pin seat in an adjacent platform to seal the working fluid path. The cooling passage is typically drilled into the inclined surface to connect the passage fluid to the cooling circuit. Therefore, the cooling passage has a linear configuration that may not adequately cool all of the platforms. For example, a cooling passage may pass through an extension that forms a damping pin seat, but does not adequately cool other portions of the inclined surface.

[0007] Cooling features may also be used with the angel wings. In this regard, another cooling feature includes cooling passages that deliver coolant into or radially around the angel wings. Mounting brackets for turbine rotor blades may also include cooling features therein. Summary of the invention

[0008] A first aspect of the present disclosure provides a turbine rotor blade, the turbine rotor blade comprising: an airfoil body, the airfoil body comprising a concave pressure side outer wall and a convex suction side outer wall connected along a leading edge and a trailing edge, the outer wall having an airfoil inner surface, the airfoil inner surface defining a radially extending chamber for receiving a coolant flow; a tip, the tip being located at a radially outer end of the airfoil body; a shank, the shank being located at a radially inner end of the airfoil body, the radially extending chamber at least partially extending into the shank to define the shank inner surface; and an impingement cooling structure, the impingement cooling structure The impingement cooling structure is located in a radially extending chamber, the impingement cooling structure comprising: a hollow body including a first end, a second end, an inner surface and an outer surface; a plurality of cooling channels passing through the hollow body and in fluid communication with the radially extending chamber to allow a coolant flow to pass from the inner surface of the hollow body to impinge on at least the airfoil inner surface, wherein the first end of the hollow body is integrally formed with the shank inner surface, and wherein the outer surface of the hollow body is evenly spaced apart from the airfoil inner surface between the first end and the second end of the hollow body.

[0009] A second aspect of the present disclosure provides an additively manufactured turbine rotor blade, the turbine rotor blade comprising: an airfoil body, the airfoil body comprising a concave pressure side outer wall and a convex suction side outer wall connected along a leading edge and a trailing edge, the outer wall having an airfoil inner surface, the airfoil inner surface defining a radially extending chamber for receiving a coolant flow; and an integral impingement cooling structure, the integral impingement cooling structure being located within the radially extending chamber, the integral impingement cooling structure comprising: a hollow body, the hollow body comprising a first end, a second end, an inner surface and an outer surface; and a plurality of cooling channels, the plurality of cooling channels passing through the hollow body and being fluidly connected to the radially extending chamber to allow a coolant flow to pass from the inner surface of the hollow body to impinge on at least the inner surface of the airfoil, wherein the outer surface of the hollow body is evenly spaced apart from the inner surface of the airfoil between the first end and the second end of the hollow body.

[0010] A third aspect of the present disclosure provides a method, which includes: sequentially forming material layers and applying a heat source to sinter the material layers to form: an airfoil body, the airfoil body including a concave pressure side outer wall and a convex suction side outer wall connected along a leading edge and a trailing edge, the outer wall having an airfoil inner surface, the airfoil inner surface defining a radially extending chamber for receiving a coolant flow; and an impingement cooling structure, the impingement cooling structure being located in the radially extending chamber, the overall impingement cooling structure including: a hollow body, the hollow body including a first end, a second end, an inner surface and an outer surface; and a plurality of cooling channels, the plurality of cooling channels passing through the hollow body and being fluidly connected to the radially extending chamber to allow a coolant flow to pass from the inner surface of the hollow body to impinge on at least the inner surface of the airfoil, wherein the outer surface of the hollow body is evenly spaced apart from the inner surface of the airfoil between the first end and the second end of the hollow body.

[0011] A fourth aspect of the present disclosure provides a turbine rotor blade, comprising: an airfoil body, the airfoil body comprising a concave pressure-side outer wall and a convex suction-side outer wall connected along a leading edge and a trailing edge, the outer walls defining a radially extending chamber for receiving a coolant flow; a platform, the platform extending laterally outward relative to the airfoil body and terminating at at least one inclined surface; a cooling circuit defined within the platform and fluidly connected to a source of the coolant flow; and at least one cooling channel defined in the platform and fluidly connected to the cooling circuit, the at least one cooling channel extending in a nonlinear configuration from the cooling circuit through the at least one inclined surface of the platform to an outlet.

[0012] A fifth aspect of the present disclosure provides an additively manufactured turbine rotor blade, the turbine rotor blade comprising: an airfoil body, the airfoil body comprising a concave pressure side outer wall and a convex suction side outer wall connected along a leading edge and a trailing edge, the outer walls defining a radially extending chamber for receiving a coolant flow; a platform, the platform extending laterally outward relative to the airfoil body and terminating at at least one inclined surface; a cooling circuit defined within the platform and fluidly connected to a source of the coolant flow; and at least one cooling channel, the at least one cooling channel defined in the platform and fluidly connected to the cooling circuit, the at least one cooling channel extending in a nonlinear configuration from the cooling circuit through the inclined surface of the platform to an outlet.

[0013] A sixth aspect includes a turbine rotor blade comprising: an airfoil body including a concave pressure side outer wall and a convex suction side outer wall connected along a leading edge and a trailing edge; a shank at a radially inner end of the airfoil body; at least one angel wing extending laterally from at least one side of the shank; and a coolant transfer passage defined through the at least one angel wing, the coolant transfer passage fluidly connecting a first wheel space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade and a second wheel space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade.

[0014] A seventh aspect of the present disclosure relates to an additively manufactured turbine rotor blade, the turbine rotor blade comprising: an airfoil body comprising a concave pressure-side outer wall and a convex suction-side outer wall connected along a leading edge and a trailing edge; a shank at a radially inner end of the airfoil body; at least one angel wing extending laterally from at least one side of the shank; and a coolant transfer passage defined through the at least one angel wing, the coolant transfer passage fluidly connecting a first wheel space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade and a second wheel space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade.

[0015] An eighth aspect relates to a group of turbine rotor blades, the group of turbine rotor blades comprising: a first turbine rotor blade, a second turbine rotor blade and a third turbine rotor blade, the first turbine rotor blade being positioned between the second turbine rotor blade and the third turbine rotor blade, each turbine rotor blade comprising: an airfoil body comprising a concave pressure side outer wall and a convex suction side outer wall connected along a leading edge and a trailing edge; a shank portion, the shank portion being located at a radially inner end of the airfoil body; at least one angel wing, the at least one angel wing extending laterally from at least one side of the shank portion, and wherein the shanks of the first turbine rotor blade and the second turbine rotor blade define a first wheel space portion therebetween, and the shanks of the first turbine rotor blade and the third turbine rotor blade define a second wheel space portion therebetween; and a coolant transfer passage defined through at least one angel wing in the first turbine rotor blade, the coolant transfer passage fluidly coupling the first wheel space portion and the second wheel space portion.

[0016] A ninth aspect relates to a turbine rotor blade root, which includes: a shank having a radially extending chamber defined therein; a blade mounting frame located at the radially inner end of the shank, the blade mounting frame having a hollow interior defined therein, the hollow interior being fluidically connected to the radially extending chamber; and a grille support structure disposed within the hollow interior of the blade mounting frame.

[0017] The tenth aspect includes a turbine rotor blade root, the turbine rotor blade root comprising: a shank having a radially extending chamber defined therein; a blade mounting frame located at a radially inner end of the shank, the blade mounting frame having a hollow interior defined therein, the hollow interior being fluidically connected to the radially extending chamber; a grating support structure disposed within the hollow interior of the blade mounting frame; at least one angel wing extending laterally from at least one side of the shank; and a coolant transfer passage defined through the at least one angel wing, the coolant transfer passage fluidly connecting a first wheel space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade root and a second wheel space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade root.

[0018] An eleventh aspect relates to a turbine rotor blade root, the turbine rotor blade root comprising: a shank having a radially extending chamber defined therein; a blade mounting frame located at a radially inner end of the shank, the blade mounting frame having a hollow interior defined therein, the hollow interior being fluidically connected to the radially extending chamber; a grating support structure disposed within the hollow interior of the blade mounting frame; at least one angel wing extending laterally from at least one side of the shank; and a coolant transfer passage defined through the at least one angel wing, the coolant transfer passage fluidly connecting a first wheel space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade root and a second wheel space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade root.

[0019] The exemplary aspects of the present disclosure are designed to solve the problems described herein and / or other problems not discussed. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] These and other features of the present disclosure will be more readily understood from the following detailed description of various aspects of the present disclosure taken in conjunction with the accompanying drawings which depict various embodiments of the present disclosure, wherein:

[0021] Figure 1 A schematic diagram of an exemplary gas turbine (GT) system is shown.

[0022] Figure 2 Shows that it can be Figure 1 A cross-sectional view of an exemplary gas turbine assembly for use with the GT system in FIG.

[0023] Figure 3 A perspective view of a turbine rotor blade of the type in which embodiments of the present disclosure may be employed is shown.

[0024] Figure 4 An axial cross-sectional view of a turbine rotor blade including an integral impingement sleeve is shown according to one embodiment of the present disclosure.

[0025] Figure 5 A radial circumferential cross-sectional view of a turbine rotor blade including an integral impingement sleeve according to one embodiment of the present disclosure is shown.

[0026] Figure 6 An enlarged cross-sectional view showing the location where an impingement cooling structure of a turbine rotor blade meets a shank according to another embodiment of the present disclosure.

[0027] Figure 7 An enlarged cross-sectional view of an impingement cooling structure with variable wall thickness according to another embodiment of the present disclosure is shown.

[0028] Figure 8 An enlarged cross-sectional view of an impingement cooling structure having reinforcement members surrounding its cooling channels according to an embodiment of the present disclosure is shown.

[0029] Fig. 9 A plan cross-sectional view of an impingement cooling structure according to an alternative embodiment of the present disclosure is shown.

[0030] Fig.10 A first partial axial cross-sectional view of a turbine rotor blade including an integral impingement sleeve is shown according to an embodiment of the present disclosure.

[0031] Fig.11 A second partial axial cross-sectional view of a turbine rotor blade including an integral impingement sleeve is shown according to an embodiment of the present disclosure.

[0032] Fig.12 A perspective view of a cooling passage in a platform of a turbine rotor blade is shown according to an embodiment of the present disclosure.

[0033] Fig.13 A transparent plan view of a cooling passage in one side of a platform of a turbine rotor blade is shown according to an embodiment of the present disclosure.

[0034] Fig.14 An enlarged cross-sectional view of a cooling passage in a platform of a turbine rotor blade is shown according to an embodiment of the present disclosure.

[0035] Fig.15 An enlarged cross-sectional view of a cooling passage in a platform of a turbine rotor blade is shown according to an embodiment of the present disclosure.

[0036] Fig.16 An enlarged cross-sectional view of a cooling passage in a platform of a turbine rotor blade is shown according to an embodiment of the present disclosure.

[0037] Fig.17 An enlarged cross-sectional view of a cooling passage in a platform of a turbine rotor blade is shown according to an embodiment of the present disclosure.

[0038] Fig.18 An enlarged cross-sectional view of a cooling passage in a platform of a turbine rotor blade is shown according to an embodiment of the present disclosure.

[0039] Fig.19 A cross-sectional view of a turbine rotor blade including an angel wing according to an embodiment of the present disclosure is shown.

[0040] Fig. 20 A transparent perspective view of a turbine rotor blade including an angel wing according to an embodiment of the present disclosure is shown.

[0041] Fig.21A schematic axial view of a set of turbine rotor blades including angel wings is shown according to an embodiment of the present disclosure.

[0042] Fig. 22 A plan view of a turbine rotor blade including an angel wing according to an embodiment of the present disclosure is shown.

[0043] Fig.23 A side view of a turbine rotor blade including an angel wing according to an embodiment of the present disclosure is shown.

[0044] Fig.24 A side view of a turbine rotor blade including an angel wing according to an embodiment of the present disclosure is shown.

[0045] Fig.25 A cross-sectional view of a hollow blade mount for a turbine rotor blade including a lattice support structure is shown according to an embodiment of the present disclosure.

[0046] Fig.26 A perspective cutaway view of a root portion of a turbine rotor blade including a lattice support structure is shown according to an embodiment of the present disclosure.

[0047] It should be noted that the drawings of the present disclosure are not necessarily drawn to scale. The drawings are intended to depict only typical aspects of the present disclosure and therefore should not be considered to limit the scope of the present disclosure. In the drawings, like numbers represent like elements between drawings. DETAILED DESCRIPTION

[0048] As an initial matter, in order to clearly describe the present disclosure, it will be necessary to select certain terms when referring to and describing relevant machine components within, for example, a turbine. In doing so, common industry terms will be used and adopted in a manner consistent with their accepted meanings, if possible. Unless otherwise indicated, such terms should be given a broad interpretation consistent with the context of the present application and the scope of the appended claims. A person of ordinary skill in the art will understand that several different or overlapping terms may be used to refer to a particular component. An object that may be described herein as a single part may include multiple components and be referenced in another context as consisting of multiple components. Alternatively, an object that may be described herein as comprising multiple components may be referred to elsewhere as a single part.

[0049] In addition, several descriptive terms may be used regularly in this article, and it should prove helpful to define these terms at the beginning of this section. Unless otherwise specified, these terms and their definitions are as follows. As used herein, "downstream" and "upstream" are terms indicating the direction relative to the flow of a fluid, such as a working fluid through a turbine, or, for example, an air flow through a combustor or a coolant through one of the components of a turbine. The term "downstream" corresponds to the direction of fluid flow, and the term "upstream" refers to the direction opposite to the flow. In the absence of any additional particularity, the terms "front" and "rear" refer to directions, where "front" refers to the front end of the turbine or its components (i.e., the compressor end), and "rear" refers to the rear end of the turbine or its components (i.e., the turbine end). In the accompanying drawings, the front and rear are generally represented by the X direction. It is usually necessary to describe parts that are in different radial positions relative to the central axis. The term "radial" refers to a movement or position perpendicular to an axis, such as the axis of a turbine rotor. In cases such as this, if the first component resides closer to the axis than the second component, this article will describe that the first component is "radially inward" or "inside" of the second component. On the other hand, if a first component resides farther from the axis than a second component, it may be stated herein that the first component is "radially outward" or "outboard" of the second component. In the drawings, the radial direction is generally represented by the Z direction. The term "axial" refers to movement or position parallel to an axis, such as the axis of a turbine rotor. Finally, the term "circumferential" refers to movement or position around an axis. Although not shown as curved in the legend of the drawings, the circumferential direction is generally represented by the Y direction in the drawings. It should be understood that such terms can be applied relative to the rotor axis of the turbine.

[0050] In addition, several descriptive terms may be used regularly herein, as described below. The terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of individual components.

[0051] The terms used herein are only used for the purpose of describing specific embodiments and are not intended to limit the present disclosure. As used herein, the singular forms "one", "a kind of" and "the" are intended to also include plural forms, unless the context clearly states otherwise. It will be further understood that when used in the specification, the terms "include" and / or "comprising" specify the presence of stated features, integers, steps, operations, elements and / or parts, but do not exclude the presence or addition of one or more other features, integers, steps, operations, elements, parts and / or their groups. "Optional" or "optionally" means that the event or situation described subsequently may or may not occur, and the description includes instances where the event occurs and instances where the event does not occur.

[0052] Where an element or layer is referred to as being "on," "engaged to," "connected to," or "coupled to" another element or layer, it may be directly on, engaged to, connected to, or coupled to the other element or layer, or there may be intervening elements or layers. In contrast, when an element is referred to as being "integral," "directly on," "directly engaged to," "directly connected to," or "directly coupled to" another element or layer, there may be no intervening elements or layers. Other words used to describe the relationship between elements should be interpreted in a similar manner (e.g., "between" versus "directly between," "adjacent" versus "directly adjacent," etc.). As used herein, the term "and / or" includes any and all combinations of one or more of the associated listed items.

[0053] As described above, the present disclosure provides a turbine rotor blade or turbine rotor blade root that includes a plurality of integral features made possible by additive manufacturing of the blade and / or root. Additive manufacturing allows for the formation of structures that provide cooling where previously not possible, improve cooling compared to conventional systems, provide additional structural strength, and / or reduce the weight of the blade.

[0054] A. Introduction

[0055] Figure 1 A schematic diagram of an exemplary turbine 90 that may include a turbine rotor blade according to various embodiments of the present disclosure is shown, the turbine rotor blade including an integral feature. In the example shown, the turbine 90 includes a gas turbine (GT) system 100, the GT system including a compressor 102 and a combustor 104. The combustor 104 includes a combustion region 105 and a fuel nozzle assembly 106. The GT system 100 also includes a turbine 108 and a common compressor / turbine shaft (sometimes referred to as a rotor) 110. In one embodiment, the GT system 100 is a 7HA or 9HA GT system, commercially available from General Electric Company, Greenville, SC, in Greenville, South Carolina, USA. The present disclosure is not limited to any particular GT system, and can be used with other engines, including, for example, General Electric's HA, F, B, LM, GT, TM and E-class engine models, as well as engine models of other companies. In addition, as described herein, the turbine rotor blades can be applied to other forms of turbines, such as steam turbines, jet engines, compressors, etc.

[0056] In operation, air flows through compressor 102 and the compressed air is supplied to combustor 104. Specifically, the compressed air is supplied to fuel nozzle assembly 106, which is integral with combustor 104. Assembly 106 is in fluid communication with combustion zone 105. Fuel nozzle assembly 106 is also connected to a fuel source ( Figure 1 106. The compressor 102 is connected to the rotor 110 (not shown) and directs fuel and air to the combustion area 105. The burner 104 is ignited and the fuel is burned. The burner 104 is in fluid communication with the turbine 108 so that the gas flow thermal energy is converted into mechanical rotational energy. The turbine 108 is rotatably coupled to the rotor 110 and drives the rotor. The compressor 102 is also rotatably coupled to the rotor 110. In the exemplary embodiment, there are multiple burners 104 and fuel nozzle assemblies 106.

[0057] Figure 2 Shows that Figure 1 1. A cross-sectional view of an exemplary turbine 108 having three stages for use with the GT system 100 in FIG. Each stage includes a set of stationary vanes or nozzles 112 and turbine rotor blades 120. The stationary nozzles 112 may be retained in the turbine 108 by radially outer platforms 114 and radially inner platforms 116. The stationary nozzles 112 may include one or more circumferentially spaced airfoils 118 ( Figure 4 ). Turbine rotor blades 120 are coupled to rotor 110 and extend between rows of stationary nozzles 112. Combustion gases are directed by stationary nozzles 112 against turbine rotor blades 120 to cause rotor 110 ( Figure 1 ) to rotate.

[0058] Figure 3 A perspective view of an exemplary turbine rotor blade 120 of a GT system 100 is shown in which integral features according to various embodiments of the present disclosure may be employed. Figure 4 An axial cross-sectional view of a turbine rotor blade 120 including an integral feature in the form of an integral impingement cooling structure 160 is shown according to various embodiments of the present disclosure. The turbine rotor blade 120 includes an airfoil body 122 including a concave pressure side outer wall 124 and a convex suction side outer wall 126 connected along a leading edge 128 and a trailing edge 130. Figure 4 As shown, the outer walls 124, 126 have an airfoil inner surface 132 that defines a radially extending chamber 134 for receiving a coolant flow 136. Figure 3 The turbine rotor blade 120 may also include a tip 140 located at a radially outer end 142 of the airfoil body 122. The turbine rotor blade 120 may also include a turbine rotor blade root 144 (hereinafter referred to as "root 144"), through which the turbine rotor blade 120 is connected, for example, via a rotor wheel 147 ( Fig.21 ) is attached to the rotor 110 ( Figure 1 ). For purposes of this disclosure, root 144 may include any portion of turbine rotor blade 120 including platform 150 and radially inward of the platform. Root 144 may include blade mounting bracket 146 configured to be mounted on rotor wheel 147 ( Fig.21 ) in corresponding slots around the periphery of the blade mount 146. The blade mount 146 may have a blade mount for mounting to the rotor disk 147 ( Fig.21 ) of any now known or later developed external configuration, such as, but not limited to, a dovetail or fir tree configuration. The turbine rotor blade 120, i.e., its root portion 144, may also include a shank 148 extending between a blade mounting frame 146 and a platform 150. The platform 150 is disposed at the junction of the airfoil body 122 and the shank 148 and defines a portion of the turbine 108 ( Figure 1 to Figure 2 ) of the inner boundary of the flow path of the airfoil. The shank 148 is thus located at the radially inner end 152 of the airfoil body 122, and the blade mount 146 is located radially inward of the shank 148. The platform 150 extends laterally outward relative to the shank 148. As will be further described herein, the radially extending chamber 134 may extend at least partially into the shank 148 to define a shank inner surface 154 ( Figure 5 ). The outer walls 124 and 126 of the airfoil body 122 extend in the radial (Z) direction from the platform 150 to the tip 140. It should be appreciated that the airfoil body 122 is the active component of the turbine rotor blade 120 that intercepts the flow of the working fluid and causes the rotor to rotate.

[0059] B. Overall impact cooling structure

[0060] In certain embodiments, the turbine rotor blade 120 may include, among other things, an airfoil body 122 and an integral feature in the form of an integral impingement cooling structure 160 therein. The impingement cooling structure is not an insert, but is made integrally with the rest of the blade, for example, by additive manufacturing. As described herein, the airfoil body 122 may include a concave pressure side outer wall 124 and a convex suction side outer wall 126 connected along a leading edge 128 and a trailing edge 130. The outer walls 124, 126 have an airfoil inner surface 132 that defines a radially extending chamber 134 for receiving a coolant flow 136. The turbine rotor blade 120 may also include a tip 140 located at a radially outer end 142 of the airfoil body 122, and a shank 148 located at a radially inner end 152 of the airfoil body 122. The radially extending chamber 134 may extend at least partially into the shank 148 to define a shank inner surface 154. The integral impingement cooling structure 160 is located within the radially extending chamber 134 and may include a hollow body 162 including a first end 164, a second end 166, an interior surface 168, and an exterior surface 170. A plurality of cooling passages 172 extend through the hollow body 162 and are in fluid communication with the radially extending chamber 134 to allow coolant flow to pass through the interior surface 168 of the hollow body 162 to impinge on at least the interior surface of the airfoil body 122. In contrast to conventional impingement inserts, the first end 164 of the hollow body 162 is integrally formed with the shank inner surface 154, i.e., formed via additive manufacturing. Thus, the exterior surface 170 of the hollow body 162 may be made uniformly spaced from the airfoil inner surface 132 between the first end 164 and the second end 166 of the hollow body 162, regardless of the curvature of the airfoil inner surface 132. In another embodiment, non-uniform but customized spacing may be employed to provide different impingement cooling, heat absorption, and / or reuse. For example, a closer standoff spacing may be employed where increased impingement cooling is desired, and a wider standoff spacing may be used where less impingement cooling is desired. Additionally, the hollow body 162 may have cooling passages around its entire perimeter and radial span to provide impingement cooling throughout the blade, not just at its leading edge. Thus, the integral impingement cooling structure allows for maximum impingement coverage with limited sacrifices typically associated with impingement inserts, and may have multiple variable cooling features. For example, a turbine rotor blade may have: a variable chord-wise width of the impingement cooling structure or its rearward pin set, an impingement cooling structure wall thickness tailored for different cooling loads, and different supports to account for different coefficients of thermal expansion (CTE) between the airfoil body and the impingement cooling structure.

[0061] like Figure 4 as well as Figure 5As shown in the radial circumferential cross-sectional view of FIG. 1 , the turbine rotor blade 120 may include an impingement cooling structure 160 located within the radially extending chamber 134. Impingement cooling is generally provided by one or more impingement inserts that are inserted into the radially extending chamber 134 and coupled to the airfoil body 122, for example, by fasteners or welding. The impingement insert is generally linear but may include some curvature. The radially extending chamber 134 has a Figure 5 With the curved airfoil inner surface 132 shown, it is not possible to have impingement inserts evenly spaced from the inner surface along the entire radial span of the blade. To address this challenge, an impingement cooling structure 160 according to an embodiment of the present disclosure is integrally formed with the remainder of the turbine rotor blade 120 via additive manufacturing.

[0062] like Figure 5 As shown, the impingement cooling structure 160 includes a hollow body 162 including a first end 164, a second end 166, an interior surface 168, and an exterior surface 170. The impingement cooling structure 160 also includes a plurality of cooling passages 172 that pass through the hollow body 162 and are in fluid communication with the radially extending chambers 134 to allow the coolant flow 136 to pass from the interior surface 168 of the hollow body to impinge on at least the airfoil inner surface 132, such as the airfoil body 122, the tip 140, the shank 148, and / or the platform 150. Compared to conventional turbine rotor blades, and as shown in FIG. Figure 4 and especially Figure 5As shown, the outer surface 170 of the hollow body 162 is uniformly spaced from the airfoil inner surface 132 between the first end 164 and the second end 166 of the hollow body 162. That is, using additive manufacturing rather than mechanical insertion, the impingement cooling structure 160 can be formed (at the same time as, for example, the airfoil inner surface 132) to have the same curvature, bend, twist, and any other shape or size to match the shape or size of its adjacent inner surface. Notably, the impingement cooling structure 160 can be uniformly spaced from the airfoil inner surface 132 along the entire radial span it covers, thereby ensuring the desired Z / D parameter on all turbine rotor blades 120. The Z / D parameter is the ratio of the separation distance Z between the outer surface 170 and the inner surface (e.g., the airfoil inner surface 132, the shank inner surface 154, etc.) of the turbine rotor blade 120 to the diameter D of the cooling channel 172 (hole) in the impingement cooling structure 160. In one example, Z / D is in the range of about 1 to about 10. In another example, Z / D may be in the range of about 2 to about 6. The separation distance Z may be less than a conventionally available distance for castings, such as less than about 1.27 mm (0.05 inches). Cooling channels 172 with a smaller diameter D than conventional castings may also be employed, for example, based on debris blockage. Advantageously, the cooling channels 172 may extend around the entire circumferential extent of the hollow body 162 so that the coolant flow 136 exits the hollow body 162 in all directions to provide impingement cooling to all of the airfoil inner surfaces 132 of the airfoil body 122. Alternatively, the cooling channels 172 may be omitted in areas where impingement cooling of the inner surfaces 132, 154 is not desired or required. The cooling channels 172 may extend along any desired radial extent of the hollow body 162.

[0063] like Figure 5 As shown, the first end 164 of the hollow body 162 is integrally formed onto the shank inner surface 154. The first end 164 meets the shank inner surface 154 at a meeting location 174, which extends around the entire periphery of the first end 164, that is, at the meeting location, there is no opening between the first end 164 and the shank inner surface 154 (except for the possible cooling channel 172). In certain embodiments, the first end 164 of the hollow body 162 is integrally formed onto the shank inner surface 154 radially inward of the platform 150. However, this particular meeting location 174 may not be required in all cases. For example, in some cases, the meeting location 174 may be located radially outward of the platform 150. As shown in FIG. Figure 5 As shown, although not required, the second end 166 of the hollow body 162 may also be integrally formed to the inner surface 176 of the tip 140. The cooling passage 172 may optionally provide impingement cooling of the tip 140, or deliver coolant to the tip 140 for other forms of cooling.

[0064] Figure 6 FIG. 1 shows a junction 174 of the impingement cooling structure 160 and the shank 148 of the turbine rotor blade 120 according to various embodiments of the present disclosure. Figure 5 ). Figure 5 and Figure 6 As shown, the first end 164 of the hollow body 162 may extend substantially in a radial direction (arrow Z) relative to a meeting position 174 of the first end 164 of the hollow body 162 and the shank inner surface 154. As used herein, "substantially in a radial direction" indicates that the first end 164 is away from the rotor 110 ( Figure 1 ) radially extends. In contrast, at least a portion of the shank inner surface 154 extends at an angle α relative to the radial direction Z from the junction 174 ( Figure 6 ) extends. In another embodiment, the handle inner surface 154 is substantially aligned in a radial direction, and the first end 164 of the hollow body gradually bends or transitions toward the meeting position 174 to maintain the angle α, for example, <30°. In another embodiment, both the handle inner surface 154 and the first end 164 of the hollow body 162 gradually bend or transition toward the meeting position 174. The angle α can be any desired angle and is within the scope of additive manufacturing, for example, <45° from the vertical direction. In order to maintain structural integrity, it is desirable that the angle α is as small as possible, for example, <10°, <20°, or <30°. Only if Figure 6 As shown, in certain embodiments, the support structure 180 may be positioned between the first end 164 of the hollow body 162 and the shank inner surface 154, for example, radially outward of the junction location 174 and radially inward of the platform 150. In other embodiments, the support structure 180 may be positioned anywhere between the outer surface 170 of the hollow body 162 and the airfoil inner surface 132, the shank inner surface 154, etc. In other embodiments, at least a portion of the support structure 180 includes a hollow support element (e.g., a grille) to enable cooling flow to flow directly from the chamber 134 to the outer wall 124 and / or 126 ( Figure 4 ). For example, it is desirable to provide film cooling directly from the radially extending chambers 134 to certain areas of the airfoil body 122, such as the leading edge 128 and / or the trailing edge 130. The support structure 180 may include any now known or later developed element capable of positioning the first end 164 of the hollow body 162 relative to the shank inner surface 154. The support structure 180 may include, but is not limited to: a lattice structure, straight or curved rods, etc. The support structure 180 may also be integrally formed via additive manufacturing.

[0065] The impingement cooling structure 160 may also include a variety of optional alternative integral cooling features. In one example, the impingement cooling structure 160 may be optionally formed to have varying wall thicknesses. The different wall thicknesses may be useful for accommodating, for example, different CTEs between the impingement cooling structure 160 and the hotter airfoil body 122, shank 148, and / or platform 150. Figure 5 As shown, the airfoil body 122 , shank 148 , and / or platform 150 may have a variety of wall thicknesses and may have varying thicknesses within their ranges. Figure 7 An enlarged partial cross-sectional view of a portion of the impingement cooling structure 160 adjacent to the airfoil body 122, the platform 150, or the shank 148 is shown. As described above, in certain embodiments, such as Figure 7 As shown, the hollow body 162 may include at least one first portion 182 having a first wall thickness W1 between its interior surface 168 and the exterior surface 170, and at least one second portion 184 having a second wall thickness W2 between its interior surface 168 and the exterior surface 170. In the example shown, the first wall thickness W1 is greater than the second wall thickness W2. Any number of thicker portions 182 and / or thinner portions 184 may be provided in the impingement cooling structure 160. The thickness of the portions 182, 184 may be any size required to meet the structural requirements and / or thermal requirements of the location.

[0066] In another exemplary optional configuration, additional supports may be desired and / or required to support the overall impingement cooling structure 160 relative to the inner surfaces 132, 154. For example, at the thinner wall thickness portion 184 ( Figure 7 ) may be desirable and / or required. To this end, Figure 7 As shown, the turbine rotor blade 120 may also include a support member 186 located on the outer surface 170 of the hollow body 162 in at least one portion 184 having a thinner wall thickness W2. The support member 186 may be integrally formed with the hollow body 162 (and the rest of the turbine rotor blade 120) to be disposed at the first end 164 ( Figure 5 ) and the second end 166 to space the exterior surface 170 of the hollow body 162 from, for example, the airfoil inner surface 132. Any number of supports 186 may be provided in the thinner wall portion 184. The support 186 may include a channel 188 therethrough that is in fluid communication with one of the plurality of cooling channels 172, i.e., to allow the coolant flow 136 to pass therethrough and impinge on the inner surface 132, 154. In certain embodiments, regardless of the wall thickness, the turbine rotor blade 120 may include a support 189 located on the exterior surface 170 of the hollow body 162. The support 189 may be integrally formed with the hollow body 162 (and the remainder of the turbine rotor blade 120) to be disposed at the first end 164 ( Figure 5 ) and the second end 166 to space the exterior surface 170 of the hollow body 162 from, for example, the airfoil inner surface 132. The supports 186, 189 may take any form that allows for the following: reducing stresses between the hotter exterior walls 124, 126 of the airfoil body 122 and the cooler impingement cooling structure 160, providing any necessary thermal expansion, providing structural support, and / or desired spacing of the hollow body 162 from the inner surface 132, 154. The supports 186, 189 may have any desired size and / or shape, such as, but not limited to, tubes, rods, etc.

[0067] Figure 8 An enlarged cross-sectional view of another alternative embodiment including a reinforcement member 190 surrounding at least one of the cooling channels 172 is shown. The reinforcement member 190 may include any structural strengthening member, such as a thicker wall, etc. Figure 4 As shown, certain embodiments may also include reinforcing ribs 192 integrally formed to the interior surface 168 of the hollow body 162. Any number of reinforcing ribs 192 may be provided, and each reinforcing rib may extend any desired radial extent of the hollow body 162. The supports 186, 189, the reinforcing member 190, and / or the reinforcing ribs 192 may be integrally formed with the remainder of the turbine rotor blade 120 via additive manufacturing.

[0068] Fig. 9 A cross-sectional view of a turbine rotor blade 120 is shown, which includes an integral impingement cooling structure 160 and additional optional integral cooling features. In an alternative embodiment, the impingement cooling structure 160 may be optionally formed to have a varying spacing Z from the inner surface 132, 154. The spacing Z can be customized to provide the desired Z / D parameters and the desired cooling at various locations. For example, the turbine rotor blade 120 may have multiple high heat load areas 195, that is, areas that experience higher temperatures and require more cooling than other areas of the blade. In the example shown, the high heat load areas 195 include the following areas: an area (195A) near the leading edge 128, an area (195B) of the pressure side outer wall 124 near the trailing edge 130, and an area (195C) of the suction side wall 126 downstream of the leading edge 128. At high heat load areas 195, a first spacing Z1 may be used between the integral impingement cooling structure 160 and the inner surface 132, 154 at the high heat load areas 195, while a larger second spacing Z2 may be used at other locations that do not have such high heat loads. In this way, the first spacing Z1 may be used to provide more cooling when necessary (i.e., at high heat load areas 195), with the spacing increasing to a second larger spacing Z2 between the impingement cooling structure 160 and the inner surface 132, 154 for use in lower heat load areas. Fig. 9As shown, the larger second spacing Z2 can allow the coolant flow 136 to limit or reduce heat absorption when moving downstream toward the trailing edge 130, thereby making the coolant flow 136 cooler and having greater heat absorption capacity for downstream areas such as the spiral cooling channel 200 and / or the pin group 206 (described herein). The transition between the spacing Z1 and Z2 can be performed at any desired rate, such as gradually over a relatively long distance, suddenly at a specific location, or at any rate in between. The second spacing Z2 can be, for example, any value between 1.01 times and 3.00 times the first spacing Z1. The Z / D parameter can be customized for each area of ​​interest. As described above, in one example, Z / D is in the range of about 1 to about 10. In another example, Z / D can be in the range of about 2 to about 6. The diameter D of the cooling channel 172 can also be configured to customize the Z / D parameter for different areas.

[0069] Fig. 9 Also shown is a turbine rotor blade 120 including one or more post-impact target features 196 located on the inner surface 132. The post-impact target feature 196 may include any now known or later developed structure located on the inner surface 132 to promote cooling. In the example shown, the impact target feature 196 includes a bump, but it may also include any structure. In one case, the hollow body 162 may include a local bulge 198 to match the contour of the impact target feature 196 and thus maintain the spacing Z (i.e., Z1 as shown in the figure). Although two target features 196 and bulge 198 pairs are shown, any number may also be used. In one embodiment, the post-impact target feature 196 may also optionally include additional integral cooling features, such as but not limited to film cooling holes 199. After impacting with the post-impact target feature 196 (i.e., its inner surface 132), the film cooling holes 199 guide the coolant flow 136 to form a cooling film 201 above the sidewalls 124, 126. Any number of film-cooling holes 199 may be implemented within each post-impingement cooling feature 196 .

[0070] Fig.10 Shown along Figure 4 A first radial cross-sectional view of the sight line 10-10, and Fig.11 Shown along Figure 4 The second radial section view of the line of sight 11-11, the latter is in the same Fig.10 In slightly different planes and in opposite directions. Fig.10 As shown, in some embodiments, the hollow body 162 has a smaller chord-wise width WC1 near the end 140 than the shank 148. Most conventional impact inserts have an opposite chord-wise width configuration to allow them to be inserted through the open end of the airfoil body. In addition, the hollow body 162 may have a smaller chord-wise width WC1 near the end 140 than the shank 148. Figure 10 to Figure 11The impingement cooling structure 160 may have an alternating wider chord-wise width and a narrower chord-wise width WC1 on the upper and lower pages in the hollow body 162. Therefore, the axial rearward end 194 of the hollow body 162 may vary in chord-wise position along the radial span of the hollow body 162. In this way, the impingement cooling structure 160 may have a curvature in its radial span to align with the airfoil inner surface 132 and / or the shank inner surface 154 ( Figure 5 ) are evenly spaced apart in shape (chord-wise width WC1) regardless of the shape of the latter. Conventional impact inserts cannot provide this feature.

[0071] like Figure 4 , Fig.10 and Fig.11 As shown, the airfoil body 122 also includes at least one chord-wise extending spiral cooling passage 200 extending from the airfoil inner surface 132 behind the hollow body 162 toward the trailing edge 130. Fig.10 and Fig.11 As best shown, each chord-wise extending spiral cooling passage 200 may have the same chord-wise width WC2, e.g., shorter than the chord-wise width WC1 of the hollow body 162. The airfoil body 122 may also include a plurality of radially spaced trailing edge cooling passages 202 extending through the trailing edge 130 (i.e., from the spiral cooling passage 200). Each of the plurality of trailing edge cooling passages 202 has the same chord-wise width WC3, i.e., a radial span along the turbine rotor blade 120. As shown in FIG. Fig.10 and Fig.11 As shown, the space 204 between the trailing edge cooling passage 202 and the spiral cooling passage 200 has a varying chord-wise width WC4, that is, a radial span along the turbine rotor blade 120. The turbine rotor blade 120 may also include a front end ( Fig.10 On the right side, Fig.11 The left side of the cooling channel 200 extends in the chord direction. Fig.10 On the left side, Fig.11 The pin group 206 between the right side of the Figure 10 to Figure 11 As shown, the pin set 206 may have a varying chord-wise width WC4 along its radial span. Fig.10 and Fig.11 Various cooling features are shown including, but not limited to, film cooling via openings 208 , main chord impingement cooling via impingement cooling structures 160 , cooling near the trailing edge 130 via helical cooling passages 200 , and trailing edge 130 pin group cooling via pin groups 206 .

[0072] Additive manufacturing (AM) includes a variety of processes for producing parts by continuous stacking of materials rather than removing materials. Therefore, additive manufacturing can form complex geometries, such as those described herein with respect to turbine rotor blades 120, without the use of any kind of tools, molds or fixtures, and with little or no waste of material. Instead of machining parts from solid material blanks (many of which are cut away and discarded), the only material used in additive manufacturing is the material required to form the part. Additive manufacturing technology generally includes taking a three-dimensional computer-aided design (CAD) file of a part to be formed (e.g., a turbine rotor blade 120), electronically slicing the part into layers (e.g., 18-102 microns thick) and creating a file with a two-dimensional image (including vectors, images or coordinates) of each layer. The file can then be loaded into a preparation software system that interprets the file so that the part can be built by different types of additive manufacturing systems. In the 3D printing, rapid prototyping (RP) and direct digital manufacturing (DDM) forms of additive manufacturing, material layers are selectively distributed, sintered, formed, deposited, etc. to form the part. The turbine rotor blades 120 may advantageously be fabricated via additive manufacturing, although other manufacturing processes such as casting may also be employed.

[0073] In metal powder additive manufacturing techniques such as direct metal laser melting (DMLM) (also known as selective laser melting (SLM)), direct metal laser sintering (DMLS), selective laser sintering (SLS), electron beam melting (EBM), and possibly other forms of additive manufacturing, layers of metal powder are sequentially melted together to form the part. More specifically, layers of fine metal powder are sequentially melted after being evenly distributed on a bed of metal powder using an applicator. Each applicator includes an applicator element in the form of a die lip, brush, scraper, or roller made of metal, plastic, ceramic, carbon fiber, or rubber that spreads the metal powder evenly over the build platform. The bed of metal powder can move in a vertical axis. The process occurs in a processing chamber with a precisely controlled atmosphere. Once each layer is formed, each two-dimensional slice of the part geometry can be fused by selectively melting the metal powder. The melting can be performed by a high-powered melting beam (such as a 100-watt ytterbium laser) to completely weld (melt) the metal powder to form a solid metal. The melting beam is moved in the XY direction using a scanning mirror and is strong enough to completely weld (melt) the metal powder to form a solid metal. The metal powder bed can be lowered for each subsequent two-dimensional layer, and the process is repeated until the part is fully formed. In order to form certain larger blades faster, some metal additive manufacturing systems use a pair of high-power lasers that work together to form the blades. Here, a method of preparing a turbine rotor blade 120 may include sequentially forming material layers and applying a heat source to sinter the material layers to form the structure described herein. Thus, additive manufacturing produces an airfoil body 122, a tip 140, a shank 148, and an impingement cooling structure 160 comprising multiple integral material layers.

[0074] The turbine rotor blades 120 may be made of a metal that can withstand the environment in which it is employed, which may include pure metals or alloys. In one example, the metal may include virtually any non-reactive metal powder, i.e., non-explosive or non-conductive powder, such as, but not limited to: cobalt-chromium-molybdenum (CoCrMo) alloy, stainless steel, austenitic nickel-chromium-based alloys such as nickel-chromium-molybdenum-niobium alloy (NiCrMoNb) (e.g., Inconel 625 or Inconel 718), nickel-chromium-iron-molybdenum alloy (NiCrFeMo) (e.g., available from Haynes International, Inc.), and nickel-chromium-iron-molybdenum alloy (NiCrFeMo). X) or nickel-chromium-cobalt-molybdenum alloy (NiCrCoMo) (e.g., Haynes 282 available from Haynes International, Inc.), etc. In another example, the metal may include almost any metal, such as, but not limited to: tool steel (e.g., H13), titanium alloy (e.g., Ti6Al4V), stainless steel (e.g., 316L), cobalt-chromium alloy (e.g., CoCrMo), and aluminum alloy (e.g., AlSi 10 Mg).

[0075] In contrast to conventional impingement inserts, the first end 164 of the hollow body 162 is integrally formed to the shank inner surface 154, i.e., formed via additive manufacturing. Thus, the outer surface 170 of the hollow body 162 can be made uniformly spaced from the inner surface between the first end 164 and the second end 166 of the hollow body 162, regardless of, for example, the curvature of the airfoil inner surface 132 and / or the shank inner surface 154. Additionally, the hollow body 162 can have cooling channels 172 around its entire circumference and radial span to provide impingement cooling throughout the blade, not just at its leading edge. Thus, the integral impingement cooling structure 160 allows for maximum impingement range without the sacrifices typically associated with impingement inserts, and can have multiple variable cooling features. For example, the turbine rotor blade 120 may have a variable chord-wise width of the impingement cooling structure 160 (i.e., width WC1) or the pin group 206 behind the structure 160 (i.e., WC4), a customized impingement cooling structure 160, different wall thicknesses for different cooling and / or structural loads, and different supports 186, 189 to account for different coefficients of thermal expansion (CTE) between the airfoil body 122 and the impingement cooling structure 160. Additional cooling features, such as turbulators (not shown), may also be provided and customized around each cooling channel 172 to optimize impingement cooling. The turbine rotor blade 120 may also include axial ventilation through the trailing edge 130, such as relative to the impingement cooling structure 160. Figure 10 to Figure 11 described.

[0076] C. Platform with nonlinear cooling channels

[0077] See also Figures 12 to 18 , showing another integral feature according to an embodiment of the present disclosure. Similar to the previous embodiments, a turbine rotor blade 120 may include an airfoil body 122 having a radially extending chamber 134 for receiving a coolant flow 136. Figure 3 and Figures 10 to 13 As best shown, the platform 150 extends laterally outward relative to the airfoil body 122 and terminates at at least one ramp 230 (eg, Fig.11 , Fig.12 ). Fig.12A perspective transparent view of the pressure side 232 of the platform 150 is shown, and Fig.13 A top-down transparent view of the suction side 231 of the platform 150 is shown. Fig.12 and Fig.13 As shown, the cooling circuit 234 is located within the platform 150 and is in fluid communication with a source of coolant 236. The source of coolant 236 may take any of a variety of forms. In one example, where the turbine rotor blade 120 includes an impingement cooling structure 160 located in a radially extending chamber 134, the source of coolant 236 may provide the coolant to the cooling circuit 234 after passing through the impingement cooling structure 160, i.e., the coolant is a post-impingement coolant. In another embodiment, the source of coolant 236 may be the radially extending chamber 134. For example, where the impingement cooling structure 160 is not provided or the impingement cooling structure is provided radially outside the platform 150, the source of coolant 236 may provide the coolant to the cooling circuit 234 directly from the radially extending chamber 134. Other sources of coolant 236 may also be used, such as a wheel space portion between the shanks 148 of adjacent turbine rotor blades 120. The cooling circuit 234 may take any form now known or later developed. Fig.13 In the example shown, the cooling circuit 234 includes a sinusoidal path through the platform 150. In contrast, in Fig.12 In the embodiment, the cooling circuit 234 includes a curved path. The cooling circuit 234 can have a less complex path or a more complex path and can extend where cooling of the platform 150 is needed.

[0078] The turbine rotor blade 120 also includes a cooling passage 240 extending from the cooling circuit 234 through the surface 242 of the ramp 230, i.e., to cool the ramp 230 and other structures. The cooling passage 240 is located in the platform 150 and is in fluid communication with the cooling circuit 234. Compared to conventional linear cooling passages, the cooling passage 240 extends from the cooling circuit 234 through at least one ramp 230 of the platform 150 to the outlet in a non-linear configuration, thereby providing improved cooling compared to linear cooling passages. For example, in Fig.12 In the embodiment of the present invention, the cooling passage 240 has a (slightly) curved shape. Any number of cooling passages 240 may be used to provide the desired cooling. In addition, the cooling passages may have any uniform or non-uniform cross-sectional shape desired, and may be spaced uniformly or non-uniformly to provide the desired cooling. Non-linear configurations are made possible by, for example, additive manufacturing. As described above, the airfoil body 122 and the platform 150 (including the components defining the cooling passages 240) may include multiple integral material layers.

[0079] Fig.14 An enlarged cross-sectional view of the inclined surface 230 is shown. In some embodiments, as Fig.14As shown, the ramp 230 may include an extension member 244. The extension member 244 may define a damping pin seat 246 that is configured to receive a damping pin 248 (shown in phantom) that seals with a damping pin seat of an adjacent turbine rotor blade (not shown). Where provided, the cooling passage 240 may extend through the extension member 244. In this regard, the cooling passage 240 may have a non-linear configuration that extends radially outward from the cooling circuit 234 and extends around the damping pin seat 246 to the outer surface 242 of the ramp 230, i.e., in a non-linear configuration relative to the outer surface 242 of the ramp 230. Fig.12 A sharper or more turned curved shape.

[0080] The cooling passages 240 may adopt any of a variety of nonlinear configurations to provide the desired cooling. The nonlinear configuration (e.g., a curved shape) may extend in any desired direction within the platform 150, such as radially (inward or outward), axially (backward or forward), or circumferentially (clockwise or counterclockwise), or a combination of these directions. The cooling passages 240 may all have the same shape to provide the same cooling properties at each location where the cooling passages are provided, or the shape of the cooling passages may vary within the platform 150 to provide customized cooling for each location where the cooling passages are provided. In addition to Fig.12 and Fig.13 In addition to the curved shape shown, Fig.15 In another embodiment shown, the cooling channel 240 may have a spiral (screw) shape, that is, have a plurality of spiral coils 250. Any number of spiral coils 250 may be used for each cooling channel 240. Fig.13 and Fig.16 As shown, the cooling channel 240 may have at least one first turn 252 ( Fig.16 ) and at least one second turn 254 ( Fig.16 ), thereby forming a generally zigzag path. Any number of first turns 252 and second turns 254 ( Fig.16 ) can be used for each cooling channel 240. Fig.16 As shown on the left side of FIG. , the amplitude A of each turn 252, 254 may be consistent so as to form a sinusoidal shape with at least one first turn and a second turn having the same amplitude A. Alternatively, as Fig.16 As shown on the right side of FIG. , the magnitude of each turn 252, 254 may be inconsistent so as to form a more random zigzag path with the turns 252, 254. Fig.16 As shown on the right side of FIG, the input 260 and the outlet 262 of each cooling channel 240 do not need to be aligned. Fig.17 In another embodiment shown, the cooling channel 240 may have a plurality of branches 264, such as a tree. Any branching configuration may be used.

[0081] Fig.18 An embodiment is shown in which the cooling channel 240 has a curved shape, for example, within the platform 150 Fig.12 Flatter. Fig.18 It is also shown that the outlet 262 of the cooling channel 240 can meet the inclined surface 230 of the platform 150 at an angle α other than 90°. In an embodiment, the angle α is less than 15°. The angle α can be customized to provide the desired cooling to the platform 150 and / or to provide film cooling to the inclined surface 230. Although shown separately, any of the cooling channel examples or aspects thereof can be combined with other examples.

[0082] The cooling circuit 234 and the cooling passage 240 may be provided separately in the pressure side 232 of the platform (only Fig.13 ), which is separately arranged in the suction side 231 of the platform 150 (only Fig.12 ), or disposed on both sides 231, 232 of the platform 150 ( Figure 12 to Figure 13 ). If provided on only one side of the platform 150, any other conventional structure may be provided on the other side of the platform. In the latter case, Fig.12 and Fig.13 As shown together, the cooling circuit 234 may include a first portion 234SS located in the suction side 231 of the platform 150 and a second portion 234PS located on the pressure side 232 of the platform 150. The portions 234SS and 234PS may be separate or fluidly coupled. In this regard, the ramp 230 includes a suction side ramp 230SS and a pressure side ramp 230PS. Here, the cooling passage 240 in the platform 150 may include: at least one first cooling passage 240 in fluid communication with the first portion 234SS of the cooling circuit and exiting the suction side ramp 230SS, and at least one second cooling passage 240 in fluid communication with the second portion 234PS of the cooling circuit 234 and exiting the pressure side ramp 230PS.

[0083] Compared to conventionally drilled linear coolant channels, non-linear cooling channels 240 allow coolant to be directed where needed in the platform 150. Additive manufacturing of coolant channels 140 enables coolant channels to have a wide variety of non-linear configurations that direct cooling where necessary and provide enhanced cooling through their shape.

[0084] D. Angel wings with coolant delivery channels

[0085] See also Figure 2 and Figures 19 to 22 Another integral feature according to an embodiment of the present disclosure includes an angel wing 280 having a coolant delivery passage therein. Fig.19 shows a radial section through a turbine rotor blade 120 including an angel wing 280, Fig. 20 A transparent perspective view of a turbine rotor blade 120 including an angel wing 280 is shown. Fig.21 An axial view of a set of turbine rotor blades 120A-C including an angel wing 280 is shown, and Fig. 22 A top-down view of a turbine rotor blade 120 including an angel wing 280 is shown. Fig.21 , a set of turbine rotor blades includes: a first turbine rotor blade 120A, a second turbine rotor blade 120B, and a third turbine rotor blade 120C (together or individually, turbine rotor blades 120). The first turbine rotor blade 120A is positioned between the second turbine rotor blade 120B and the third turbine rotor blade 120C. In this embodiment, as Figure 19 to Figure 20 As shown, each turbine rotor blade 120 may include an airfoil body 122 including a front edge 128 and a rear edge 130 ( Figure 3 ) connected to the concave pressure side outer wall 124 ( Figure 3 ) and the convex suction side outer wall 126 ( Figure 3 The turbine rotor blade 120 may further include a shank portion 148 at the radially inner end 152 of the airfoil body 122. In addition, the turbine rotor blade 120 includes at least one angel wing 280 extending laterally from at least one side 282, 284 of the shank portion 148.

[0086] like Figure 2 As shown in FIG. 1 , there is an opening 286 at the interface between the adjacent nozzle 112 and the turbine rotor blade 120 that allows hot process fluid to leave the hot gas path and enter the wheel space 300 of the turbine 108. To limit this leakage of hot gases, the turbine rotor blade 120 typically includes an axially protruding angel wing seal 280, also referred to as an angel wing for short. The angel wing 280 cooperates with a protruding section or "obstruction" 288 extending from the nozzle 112. The angel wing 280 and the obstruction 288 overlap (or nearly overlap), but do not contact each other, thereby restricting fluid flow.

[0087] Go to Figures 19 to 22 According to an embodiment of the present disclosure, the turbine rotor blade 120 may further include a coolant delivery passage 290 defined by at least one angel wing 280. For example, for the first turbine rotor blade 120A ( Fig.21 ), the coolant delivery passage 290 is fluidly coupled to the handle 148A ( Fig.21 ) and the first adjacent shank 148B of the first adjacent turbine rotor blade 120B ( Fig.21 ) and the first wheel space portion 292 defined between the handle 148A ( Fig.21 ) and the second adjacent turbine rotor blade 120C ( Fig.21) between the second adjacent handles 148C of the second wheel space portion 294. As can be seen by Figure 2 , Fig.21 and Fig. 22 As best shown, each wheel space portion 292, 294 is a portion of a wheel space 300. The wheel space 300 is defined as a space circumferentially defined between the shanks 148A-C ( Fig.21 ), axially defined between the shank 148 and the adjacent nozzle 112, and radially defined by the platform 150 and the rotor disk 147. The wheel space portions 292, 294 are portions of the wheel space 300 axially alongside the shank 148 of a particular blade.

[0088] like Fig. 22 As best shown, the coolant transfer passage 290 includes a first open end 310 in fluid communication with the first wheel space portion 292 and a second open end 312 in fluid communication with the second wheel space portion 294. Thus, the coolant transfer passage 290 allows wheel space coolant 316 to pass between the wheel space portions 292, 294 on circumferentially opposite sides of the shank 148. The first open end 310 and the second open end 312 may face in an axial-circumferential direction relative to the airfoil body 122, or any direction that will allow the wheel space coolant 316 to pass between the wheel space portions 292, 294. The wheel space coolant 316 may be any now known or later developed coolant, such as a refrigerant from the compressor 102 ( Figure 1 As previously described, the outer walls 124, 126 of the airfoil body 122 define a radially extending chamber 134 that may extend into the shank 148. Fig. 22 As shown, the coolant transfer passage 290 is fluidly isolated from the radially extending chamber 134, that is, the coolant flow 136 ( Figure 4 ) does not mix with the wheel space coolant 316.

[0089] Any number of angel wings 280 may be used. In one example, Fig.19 , Fig. 20 and Fig. 22 As shown, the first angel wing 280 extends laterally from a first side 282 of the handle 148, and the second angel wing 280 extends laterally from a second opposite side 284 of the handle 148. In another example, as shown in FIG. Fig.23 As shown, the first pair of radially spaced angel wings 280 may extend laterally from the first side 282 of the handle 148, and neither of them extends from the side 284 of the handle 148. In another embodiment, as shown in FIG. Fig.24As shown, a first pair of radially spaced angel wings 280 may extend laterally from a first side 282 of the shank 148, and a second pair of radially spaced angel wings 280 may extend laterally from a second, opposite side 284 of the shank 148. In any case, each angel wing 280 may include a corresponding coolant delivery passage 290. Alternatively, while each angel wing 280 is shown as including a coolant delivery passage 290, selective angel wings may not include a coolant delivery passage.

[0090] The coolant transfer passage 290 allows the wheel space coolant 316 ( Fig. 22 ) moves between the wheel space portions 292 , 294 , allowing cooling of the angel wing 280 and reducing the weight of the turbine rotor blade 120 .

[0091] E. Hollow blade mounting frame with grille support structure

[0092] See also Fig.19 , Fig. 20 , Fig.25 and Fig.26 , another integral feature according to an embodiment of the present disclosure includes a hollow blade mount 146. In this embodiment, a root 144 is provided, which includes a shank 148 having a radially extending chamber 134 defined therein. The blade mount 146 is located at the radially inner end of the shank 148. Compared to many conventional blade mounts, the blade mount 146 has a hollow interior 330 defined therein, for example, by an inner wall surface 344 of the blade mount 146. The hollow interior 330 is in fluid communication with the radially extending chamber 134. The hollow interior 330 can have any desired interior shape, for example, such as Fig.25 The blade mounting frame 146 may have a structure configured to be mounted to the rotor 110 ( Figure 2 ) of the rotor wheel 147 ( Fig.21 ) any now known or later developed external shape, such as a dovetail or fir tree shape.

[0093] The turbine rotor blade root 144 may also include a lattice support structure 340 disposed within the hollow interior 330 of the blade mount 146. The lattice support structure 340 may take the form of a variety of hollow support structures. In one example, the lattice support structure 340 may include a plurality of radially extending V-shaped portions 342. The V-shaped portions 342 may be integral with the inner wall surface 344 of the blade mount 146. The root 144 including the shank 148 and the blade mount 146 and including the lattice support structure 340 may be made by additive manufacturing. The shank 148 and the blade mount 146 may therefore include a plurality of integral material layers.

[0094] According to this embodiment, the root 144, i.e., the root having the lattice support structure 340, may also include a platform 150, as described herein with respect to Figures 12 to 18 As described above, the platform 150 is positioned radially outward of the shank 148 and extends laterally outward relative to the shank, terminating in at least one ramp 230. The platform 150 may include a cooling circuit 234 defined within the platform and in fluid communication with a source of coolant flow (e.g., the radially extending chamber 134). The cooling passage 240 ( Figures 12 to 18 ) may be defined in the platform 150 and in fluid communication with the cooling circuit 234. As described above, the cooling passage 240 extends from the cooling circuit 234 through the inclined surface 230 of the platform to the outlet in a non-linear configuration. The inclined surface 230 may include an extension member 244 through which the cooling passage 240 extends. The cooling passage 240 may have: a spiral shape ( Fig.15 ); at least one first turn in a first direction and at least one second turn in an opposite second direction ( Fig.16 ); multiple branches ( Fig.17 ); or curved shapes (e.g. Fig.12 , Fig.14 , Fig.18 ).

[0095] According to this embodiment, the root 144, i.e., the root having the lattice support structure 340, may also include an angel wing 280 extending laterally from at least one side of the shank 148, as described herein with respect to Figures 19 to 24 As described above, the coolant delivery channel 290 may be defined by the angel wing 280. Fig.21 As shown, the coolant delivery passage 290 fluidly couples a first wheel space portion 292 defined between the shank portion 148A and a first adjacent shank portion 148B of a first adjacent turbine rotor blade root 144B and a second wheel space portion 294 defined between the shank portion 148A and a second adjacent shank portion 148C of a second adjacent turbine rotor blade root 144C. The coolant delivery passage 290 includes a first open end 310 in fluid communication with the first wheel space portion 292 and a second open end 312 in fluid communication with the second wheel space portion 294. Fig. 22 As shown, the first open end 310 and the second open end 312 may face in a circumferential direction relative to the shank 148. The coolant delivery passage 290 may be fluidly isolated from the radially extending chamber 134 in the shank 148.

[0096] The root 144 according to this embodiment, i.e., the root having the lattice support structure 340, may also include both the platform 150 and the angel wing 280, as described herein. Additive manufacturing allows the root 144 having the shank 148, the hollow blade mount 146, the lattice support structure 340, and the platform 150 and / or the angel wing 280 to be formed to form multiple integral material layers for any features provided.

[0097] The root portion 144 including an integral grid support structure 340 within the hollow interior 330 of the blade mount 146 provides additional cooling of the lighter turbine rotor blade 120 and the blade mount 146 .

[0098] Although various embodiments have been described and shown herein as being used together, it should be understood that the various embodiments may be used alone or in combination.

[0099] As used throughout the specification and claims, approximate language may be used to modify any quantitative representation that may be permissibly varied without resulting in a change in the basic function to which it is related. Thus, values ​​modified by one or more terms, such as "about," "approximately," and "substantially," are not limited to the precise values ​​specified. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value. Here and throughout the specification and claims, range limitations may be combined and / or interchanged; unless the context or language indicates otherwise, these ranges are identified and include all subranges contained therein. "Approximately" applied to a particular value of a range applies to both end values ​​and may indicate + / -10% of the value unless otherwise dependent on the precision of the instrument for measuring the value.

[0100] The corresponding structures, materials, actions and equivalents of all means or steps plus function elements in the following claims are intended to include any structure, material or action for performing functions in conjunction with other claimed elements specifically claimed. A description of the present disclosure has been given for the purpose of illustration and description, but it is not intended to be exhaustive or to limit the present disclosure to the disclosed form. Many modifications and variations will be apparent to those of ordinary skill in the art without departing from the scope and essence of the present disclosure. Embodiments have been selected and described in order to best explain the principles and practical applications of the present disclosure, and to enable other persons of ordinary skill in the art to understand the various embodiments of the present disclosure with various modifications suitable for the intended specific use.

Claims

1. A turbine rotor blade root (144), comprising: a shank (148, 148A) having a radially extending chamber (134) defined therein; a blade mount (146) located at a radially inner end (152) of the shank (148, 148A), the blade mount (146) having a hollow interior (330) defined therein, the hollow interior (330) being in fluid communication with the radially extending chamber (134); a grille support structure (340) disposed within the hollow interior (330) of the blade mounting frame (146); a platform (150) positioned radially outward of the shank (148, 148A) and extending laterally outward relative to the shank (148, 148A) and terminating at at least one bevel (230, 230PS, 230SS); a cooling circuit (234) defined within the platform (150) and in fluid communication with a source of a coolant (236) flow; and at least one cooling passage (172, 240) defined in the platform (150) and in fluid communication with the cooling circuit (234), the at least one cooling passage (172, 240) extending in a non-linear configuration from the cooling circuit (234) through the at least one inclined surface (230, 230PS, 230SS) of the platform (150) to an outlet (262); wherein the at least one sloped surface (230, 230PS, 230SS) includes an extension member (244), and wherein the at least one cooling channel (172, 240) extends through the extension member (244).

2. The turbine rotor blade root (144) of claim 1, wherein the lattice support structure (340) includes a plurality of radially extending V-shaped portions (342).

3. The turbine rotor blade root (144) of claim 1, wherein the at least one cooling passage (172, 240) has one of: a spiral shape; at least one first turn (252) in a first direction and at least one second turn (254) in an opposite second direction; a plurality of branches (264); and a curved shape.

4. The turbine rotor blade root (144) of claim 1, wherein the shank (148, 148A) and the blade mounting bracket (146) comprise a plurality of integral material layers.

5. The turbine rotor blade root (144) of claim 1, further comprising: at least one angel wing (280) extending laterally from at least one side (282, 284) of the handle (148, 148A); and A coolant (236) delivery passage is defined through the at least one angel wing (280), the coolant (236) delivery passage fluidly coupling a first wheel space portion (292) (294) defined between the shank (148, 148A) and a first adjacent shank (148B) of a first adjacent turbine rotor blade root (144) and a second wheel space portion (292) (294) defined between the shank (148, 148A) and a second adjacent shank (148C) of a second adjacent turbine rotor blade root (144).

6. The turbine rotor blade root (144) of claim 5, wherein the coolant delivery passage (290) includes a first open end (310) in fluid communication with the first wheel space portion (292) (294) and a second open end (312) in fluid communication with the second wheel space (292) (294), wherein the first open end (310) and the second open end (312) face in a circumferential direction relative to the shank (148, 148A).

7. The turbine rotor blade root (144) of claim 6, wherein the coolant delivery passage (290) is fluidly isolated from the radially extending chamber (134) in the shank (148, 148A).

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