Additively Manufactured Turbine Rotor Blade with a Platform Having Nonlinear Cooling Channels

Abstract

The present invention is entitled a turbine rotor blade with a platform having non-linear cooling channels for additive manufacturing. The present invention provides a turbine rotor blade (120) that is additively manufactured and includes an airfoil body (122) having a radially extending chamber (134) for receiving a coolant (236) flow. A platform (150) extends laterally outward relative to the airfoil body (122) and terminates at at least one bevel (230, 230PS, 230SS). A cooling circuit (234) is located within the platform (150) and is in fluid communication with a source of coolant (236) flow. Cooling channels (172, 240) are located in the platform (150) and are in fluid communication with the cooling circuit (234). The cooling channels (172, 240) extend from the cooling circuit (234) through at least one bevel (230, 230PS, 230SS) of the platform (150) to an outlet in a non-linear configuration, thereby providing improved cooling compared to linear cooling channels (172, 240).

Description

BACKGROUND OF THE DISCLOSURE

[0001] The present disclosure relates generally to turbines and, more specifically, to turbine rotor blades having multiple additively manufactured integral features such as integral impingement sleeves, non-linear cooling channels in platforms, airfoils having coolant delivery channels, and hollow tails 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 combusted fuel is forced against the blades to cause them to rotate 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 impingement insert. The impingement 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 strike a surface to be cooled. Impingement inserts are used, for example, in various hot gas path (HGP) components in turbomachinery such as turbine rotor blades to improve the cooling performance of the cooling circuits therein. One problem faced by impingement inserts is positioning the impingement insert in a tapered or curved cavity in an HGP component close enough to achieve high cooling performance but not so close that the cooling is ineffective. One metric of the cooling performance of an impingement insert is the Z / D parameter, which is the ratio of the separation distance Z between the insert and the inner surface of the HGP component to the diameter D of the cooling channels (holes) in the impingement 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] In cases where the necessary separation distance cannot be formed, impingement cooling is typically not provided by the insert. For example, if the cavity in the HGP component is curved too significantly such that the impingement insert cannot be made thin enough or curved to the corresponding necessary separation distance, impingement cooling cannot be provided. One way to address 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 in sequence or couple them to the HGP component increases the manufacturing complexity, time, and cost. The flexible impingement insert sections also do not provide adjacent elements around their perimeter (i.e., laterally (in cross-section)), which can detract from the cooling performance at their discontinuities.

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

[0006] Another cooling feature includes a cooling passage that extends through a portion of the turbine rotor blade to be cooled. For example, the turbine rotor blade includes a platform that extends laterally to cooperate with the platforms of adjacent turbine rotor blades to form a portion of the 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 the beveled surface of the platform. Some platforms include damping pin receptacles in the beveled surface that receive axially extending pins therein, and the axially extending pins cooperate with adjacent damping pin receptacles in adjacent platforms to seal the working fluid path. The cooling passages are typically drilled into the beveled surface to fluidly couple the passage fluid to the cooling circuit. Thus, the cooling passages have a linear configuration that may not adequately cool all of the platforms. For example, the cooling passages may extend through the extensions that form the damping pin receptacles but do not adequately cool other portions of the beveled surface.

[0007] The cooling feature may also be used with the airfoil. In this regard, another cooling feature includes cooling passages that deliver coolant into or radially around the airfoil. The mount for the turbine rotor blade may also include a cooling feature therein. Summary of the Invention

[0008] A first aspect of the present disclosure provides a turbine rotor blade including: 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, the outer walls having an airfoil inner surface that defines a radially extending chamber for receiving a coolant flow; a tip located at a radially outer end of the airfoil body; a shank located at a radially inner end of the airfoil body, the radially extending chamber at least partially extending into the shank to define a shank inner surface; and an impingement cooling structure located within the radially extending chamber, the impingement cooling structure including: a hollow body including a first end, a second end, an inner surface, and an outer surface; a plurality of cooling passages that pass through the hollow body and are 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 from the airfoil inner surface between the first end and the second end of the hollow body.

[0009] The second aspect of the present disclosure provides an additive manufactured turbine rotor blade, the 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, the outer walls having an airfoil inner surface that defines a radially extending chamber for receiving a coolant flow; and an integral impingement cooling structure located within the radially extending chamber, the integral impingement cooling structure comprising: a hollow body including a first end, a second end, an inner surface, and an outer surface; and a plurality of cooling channels passing through the hollow body and in fluid communication with the radially extending chamber to allow coolant flow to pass from the inner surface of the hollow body to impinge on at least the airfoil inner surface, wherein the outer surface of the hollow body is uniformly spaced from the airfoil inner surface between the first end and the second end of the hollow body.

[0010] The third aspect of the present disclosure provides a method comprising: sequentially forming material layers and applying a heat source to sinter the material layers to form: 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, the outer walls having an airfoil inner surface that defines a radially extending chamber for receiving a coolant flow; and an impingement cooling structure located within the radially extending chamber, the integral impingement cooling structure comprising: a hollow body including a first end, a second end, an inner surface, and an outer surface; and a plurality of cooling channels passing through the hollow body and in fluid communication with the radially extending chamber to allow coolant flow to pass from the inner surface of the hollow body to impinge on at least the airfoil inner surface, wherein the outer surface of the hollow body is uniformly spaced from the airfoil inner surface between the first end and the second end of the hollow body.

[0011] The fourth aspect of the present disclosure provides a turbine rotor blade, the 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, the outer walls defining a radially extending chamber for receiving a coolant flow; a platform that extends laterally outward relative to the airfoil body and terminates at at least one bevel; a cooling circuit defined within the platform and in fluid communication with a source of coolant flow; and at least one cooling channel defined within the platform and in fluid communication with the cooling circuit, the at least one cooling channel extending from the cooling circuit through at least one bevel of the platform to an outlet in a non-linear configuration.

[0012] A fifth aspect of the present disclosure provides an additively manufactured turbine rotor blade, the 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, the outer walls defining a radially extending chamber for receiving a coolant flow; a platform extending laterally outward relative to the airfoil body and terminating at at least one bevel; a cooling circuit defined within the platform and in fluid communication with a source of the coolant flow; and at least one cooling channel defined within the platform and in fluid communication with the cooling circuit, the at least one cooling channel extending from the cooling circuit through the bevel of the platform to an outlet in a non-linear configuration.

[0013] A sixth aspect includes a turbine rotor blade, the 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 located at a radially inner end of the airfoil body; at least one fin extending laterally from at least one side of the shank; and a coolant transfer channel defined through the at least one fin, the coolant transfer channel fluidly coupling a first annulus space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade and a second annulus 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 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 located at a radially inner end of the airfoil body; at least one fin extending laterally from at least one side of the shank; and a coolant transfer channel defined through the at least one fin, the coolant transfer channel fluidly coupling a first annulus space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade and a second annulus space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade.

[0015] The eighth aspect relates to a set of turbine rotor blades, the set of turbine rotor blades including: 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 including: 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 located at a radially inner end of the airfoil body; at least one fin extending laterally from at least one side of the shank, and wherein the shanks of the first turbine rotor blade and the second turbine rotor blade define a first annular space portion therebetween, and the shanks of the first turbine rotor blade and the third turbine rotor blade define a second annular space portion therebetween; and a coolant delivery passage defined through at least one fin in the first turbine rotor blade, the coolant delivery passage fluidly coupling the first annular space portion and the second annular space portion.

[0016] The ninth aspect relates to a turbine rotor blade root, the turbine rotor blade root including: a shank having a radially extending chamber defined therein; a blade mount located at a radially inner end of the shank, the blade mount having a hollow interior defined therein, the hollow interior being in fluid communication with the radially extending chamber; and a grid support structure disposed within the hollow interior of the blade mount.

[0017] The tenth aspect includes a turbine rotor blade root, the turbine rotor blade root including: a shank having a radially extending chamber defined therein; a blade mount located at a radially inner end of the shank, the blade mount having a hollow interior defined therein, the hollow interior being in fluid communication with the radially extending chamber; a grid support structure disposed within the hollow interior of the blade mount; at least one fin extending laterally from at least one side of the shank; and a coolant delivery passage defined through the at least one fin, the coolant delivery passage fluidly coupling a first annular space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade root and a second annular space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade root.

[0018] The eleventh aspect relates to a root of a turbine rotor blade, the root of the turbine rotor blade comprising: a shank having a radially extending chamber defined therein; a blade mounting bracket located at a radially inner end of the shank, the blade mounting bracket having a hollow interior defined therein, the hollow interior being in fluid communication with the radially extending chamber; a grid support structure disposed within the hollow interior of the blade mounting bracket; at least one wing extending laterally from at least one side of the shank; and a coolant transfer passage defined through the at least one wing, the coolant transfer passage being fluidly coupled to 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] Exemplary aspects of the present disclosure are designed to address 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 in conjunction with the drawings depicting the 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 an exemplary Figure 1 cross-sectional view of an exemplary gas turbine assembly that can be used with the GT system in.

[0023] Figure 3 A perspective view of a turbine rotor blade of a type that can employ embodiments of the present disclosure is shown.

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

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

[0026] Figure 6 An enlarged cross-sectional view of the meeting location of the impingement cooling structure of a turbine rotor blade and the shank according to another embodiment of the present disclosure is shown.

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

[0028] Figure 8 Shows an enlarged cross-sectional view of an impingement cooling structure according to an embodiment of the present disclosure, the impingement cooling structure having a reinforcement member surrounding its cooling channel.

[0029] Figure 9 Shows a plan cross-sectional view of an impingement cooling structure according to an alternative embodiment of the present disclosure.

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

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

[0032] Figure 12 Shows a perspective view of a cooling channel in a platform of a turbine rotor blade according to an embodiment of the present disclosure.

[0033] Figure 13 Shows a transparent plan view of a cooling channel in one side of a platform of a turbine rotor blade according to an embodiment of the present disclosure.

[0034] Figure 14 Shows an enlarged cross-sectional view of a cooling channel in a platform of a turbine rotor blade according to an embodiment of the present disclosure.

[0035] Figure 15 Shows an enlarged cross-sectional view of a cooling channel in a platform of a turbine rotor blade according to an embodiment of the present disclosure.

[0036] Figure 16 Shows an enlarged cross-sectional view of a cooling channel in a platform of a turbine rotor blade according to an embodiment of the present disclosure.

[0037] Figure 17 Shows an enlarged cross-sectional view of a cooling channel in a platform of a turbine rotor blade according to an embodiment of the present disclosure.

[0038] Figure 18 Shows an enlarged cross-sectional view of a cooling channel in a platform of a turbine rotor blade according to an embodiment of the present disclosure.

[0039] Figure 19 Shows a cross-sectional view of a turbine rotor blade including a wing portion according to an embodiment of the present disclosure.

[0040] Figure 20 Shows a transparent perspective view of a turbine rotor blade including a wing portion according to an embodiment of the present disclosure.

[0041] Figure 21Shows a schematic axial view of a set of turbine rotor blades including airfoils according to an embodiment of the present disclosure.

[0042] Figure 22 Shows a plan view of a turbine rotor blade including an airfoil according to an embodiment of the present disclosure.

[0043] Figure 23 Shows a side view of a turbine rotor blade including an airfoil according to an embodiment of the present disclosure.

[0044] Figure 24 Shows a side view of a turbine rotor blade including an airfoil according to an embodiment of the present disclosure.

[0045] Figure 25 Shows a cross-sectional view of a hollow blade mount of a turbine rotor blade including a grid support structure according to an embodiment of the present disclosure.

[0046] Figure 26 Shows a perspective cross-sectional view of the root of a turbine rotor blade including a grid support structure 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 should not therefore be considered to limit the scope of the present disclosure. In the drawings, like numbers represent like elements between the 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, if possible, common industry terms will be used and adopted in a manner consistent with their accepted meanings. 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. Those of ordinary skill in the art will appreciate that several different or overlapping terms may typically 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 referred to in another context as being composed of multiple components. Alternatively, an object that may be described herein as including multiple components may elsewhere be referred to as a single part.

[0049] In addition, several descriptive terms may be regularly used herein, 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 directions relative to the direction of fluid flow, such as the working fluid passing through a turbine, or for example the air flow through a burner or the 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 component (i.e., the compressor end), and "rear" refers to the rear end of the turbine or its component (i.e., the turbine end). In the drawings, the front and rear are typically represented by the X direction. It is often necessary to describe parts at different radial positions relative to the central axis. The term "radial" refers to movement or position perpendicular to the axis, such as the axis of a turbine rotor. In such cases, if a first component resides closer to the axis than a second component, it will be stated herein that the first component is "radially inward" or "inner" 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 "outer" of the second component. In the drawings, the radial direction is typically represented by the Z direction. The term "axial" refers to movement or position parallel to the axis, such as the axis of a turbine rotor. Finally, the term "circumferential" refers to movement or position around the axis. Although not shown as curved in the legends of the drawings, the circumferential direction is typically represented by the Y direction in the drawings. It should be understood that such terms may be applied relative to the rotor axis of the turbine.

[0050] In addition, several descriptive terms may be regularly used herein as follows. 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 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 as well, unless the context clearly indicates otherwise. It will be further understood that when used in the specification, the terms "comprises" and / or "comprising" specify the presence of stated features, integers, steps, operations, elements, and / or components, but do not preclude 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 subsequent described event or circumstance may or may not occur, and the description includes instances where the event occurs and instances where the event does not occur.

[0052] In cases where an element or layer is referred to as being "on", "engaged to", "connected to", or "coupled to" another element or layer, it can be directly on, engaged to, connected to, or coupled to the other element or layer, or intervening elements or layers may be present. In contrast, when an element is referred to as being "directly on", "directly engaged to", "directly connected to", or "directly coupled to" another element or layer, intervening elements or layers may not be present. 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 a turbine rotor blade root that includes a plurality of integral features made possible by additive manufacturing of the blade and / or the root. Additive manufacturing allows for the formation of a structure that provides cooling where it was not previously possible, improves cooling compared to conventional systems, provides additional structural strength and / or reduces 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 in accordance with various embodiments of the present disclosure is shown. The turbine rotor blade includes integral features. In the example shown, the turbine 90 includes a gas turbine (GT) system 100 that includes a compressor 102 and a combustor 104. The combustor 104 includes a combustion zone 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 that is commercially available from General Electric Company, Greenville, S.C., USA. The present disclosure is not limited to any particular GT system and can be employed 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 from other companies. Additionally, as described herein, the turbine rotor blade 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 compressed air is supplied to burner 104. Specifically, compressed air is supplied to fuel nozzle assembly 106, which is integral with burner 104. Assembly 106 is in fluid communication with combustion zone 105. Fuel nozzle assembly 106 is also in fluid communication with a fuel source ( Figure 1 not shown) and directs fuel and air to combustion zone 105. Burner 104 is ignited and fuel is burned. Burner 104 is in fluid communication with turbine 108 so that the thermal energy of the gas flow is converted into mechanical rotational energy. Turbine 108 is rotatably coupled to rotor 110 and drives the rotor. Compressor 102 is also rotatably coupled to rotor 110. In an exemplary embodiment, there are multiple burners 104 and fuel nozzle assemblies 106.

[0057] Figure 2 Shown is a cross-sectional view of an exemplary three-stage turbine 108 that can be used with Figure 1 the GT system 100 in. Each stage includes stationary vanes or nozzle sets 112 and turbine rotor blades 120. The stationary nozzles 112 can be held in turbine 108 by a radially outer platform 114 and a radially inner platform 116. The stationary nozzles 112 can include one or more circumferentially spaced airfoils 118 ( Figure 4 ). The turbine rotor blades 120 are coupled to rotor 110 and extend between rows of the stationary nozzles 112. Combustion gases are directed by the stationary nozzles 112 against the turbine rotor blades 120 to rotate the rotor 110 ( Figure 1 ).

[0058] Figure 3 Shown is a perspective view of an exemplary turbine rotor blade 120 of the GT system 100, in which integral features according to various embodiments of the present disclosure can be employed. Figure 4 Shown is an axial cross-sectional view of a turbine rotor blade 120 according to various embodiments of the present disclosure, the turbine rotor blade including an integral feature in the form of an integral impingement cooling structure 160. The turbine rotor blade 120 includes an airfoil body 122 that includes 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. As Figure 4 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. Referring again to Figure 3 , the turbine rotor blade 120 can also include a tip 140 located at the radially outer end 142 of the airfoil body 122. The turbine rotor blade 120 can also include a turbine rotor blade root 144 (hereinafter referred to as "root 144"), through which the turbine rotor blade 120 is attached to the rotor 110, for example via a rotor wheel 147 ( Figure 21 ). ) attached to the rotor 110 (Figure 1 ) For the purposes of this disclosure, the root 144 may include any part of the turbine rotor blade 120, including the platform 150 and radially inward of the platform. The root 144 may include a blade mount 146 that is configured to be mounted in a corresponding slot in the periphery of the rotor wheel 147( Figure 21 ). The blade mount 146 may have any currently known or later developed external configuration for mounting to the rotor disk 147( Figure 21 ), such as but not limited to a dovetail or fir tree arrangement. The turbine rotor blade 120, i.e., its root 144, may also include a shank 148 that extends between the blade mount 146 and the platform 150. The platform 150 is disposed at the junction of the airfoil body 122 and the shank 148 and defines a part of the inner boundary of the flow path through the turbine 108( Figures 1 to 2 ). 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 an inner surface 154 of the shank( Figure 5 ). The outer walls 124 and 126 of the airfoil body 122 extend from the platform 150 to the tip 140 in the radial (Z) direction. It should be understood that the airfoil body 122 is the active part of the turbine rotor blade 120 that intercepts the flow of the working fluid and causes the rotor to rotate.

[0059] B. Overall Impingement Cooling Structure

[0060] In some embodiments, the turbine rotor blade 120 may particularly include an airfoil body 122 and integral features in the form of an integral impingement cooling structure 160 therein. The impingement cooling structure is not an insert, but is integrally formed with the remainder 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 further include a tip 140 at a radially outer end 142 of the airfoil body 122 and a shank 148 at a radially inner end 152 of the airfoil body 122. The radially extending chamber 134 may at least partially extend 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 that includes a first end 164, a second end 166, an inner surface 168, and an outer surface 170. A plurality of cooling channels 172 extend through the hollow body 162 and are in fluid communication with the radially extending chamber 134 to allow the coolant flow to pass from the inner surface 168 of the hollow body 162 to impinge on at least the inner surface of the airfoil body 122. Compared to a conventional impingement insert, 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 outer surface 170 of the hollow body 162 can be made to be 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, to provide different impingement cooling, heat absorption, and / or reuse, non-uniform but customized spacing can be employed. For example, a closer spacing can be employed in cases where increased impingement cooling is needed, and a wider spacing can be used in cases where less impingement cooling is needed. Additionally, the hollow body 162 may have cooling channels around its entire perimeter and radial span to provide impingement cooling over the entire blade and not just at its leading edge. Thus, the integral impingement cooling structure allows for a maximum impingement range to be achieved with a limited sacrifice typically associated with impingement inserts and may have multiple variable cooling features. For example, the turbine rotor blade may have: a variable chordwise width of the impingement cooling structure or its rear pin set, a wall thickness of the impingement cooling structure customized for different cooling loads, and different supports to address different coefficients of thermal expansion (CTE) between the airfoil body and the impingement cooling structure.

[0061] As Figure 4 and Figure 5As shown in the radial and circumferential cross-sectional view, the turbine rotor blade 120 may include an impingement cooling structure 160 located within a radially extending chamber 134. Impingement cooling is typically 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 inserts are typically linear but may include some curvature. In the case where the radially extending chamber 134 has a curved airfoil inner surface 132 as shown in Figure 5 , it is not possible to have impingement inserts that are uniformly spaced from the inner surface along the entire radial span of the blade. To address this challenge, the 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] As shown in Figure 5 , the impingement cooling structure 160 includes a hollow body 162 that includes a first end 164, a second end 166, an inner surface 168, and an outer surface 170. The impingement cooling structure 160 also includes a plurality of cooling channels 172 that pass through the hollow body 162 and are in fluid communication with the radially extending chamber 134 to allow a coolant flow 136 to pass from the inner surface 168 of the hollow body to impinge on at least the airfoil inner surface 132, for example, may impinge on the inner surfaces of, in particular, the airfoil body 122, the tip 140, the shank 148, and / or the platform 150. Compared to a conventional turbine rotor blade, and as shown in Figure 4 and in particular 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 (simultaneously with, for example, the airfoil inner surface 132) to have the same curvature, bend, twist, and any other shape or dimension to match the shape or dimension of the 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 of the turbine rotor blade 120 and the inner surface (e.g., the airfoil inner surface 132, the shank inner surface 154, etc.) to the diameter D of the cooling channels 172 (holes) 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 can be in the range of about 2 to about 6. The separation distance Z can be less than the conventionally available distance for castings, such as less than about 1.27 millimeters (0.05 inches). For example, depending on debris clogging, cooling channels 172 with a smaller diameter D than conventional castings can also be employed. Advantageously, the cooling channels 172 can extend around the entire peripheral extent of the hollow body 162 such that the coolant flow 136 exits the hollow body 162 in all directions to provide impingement cooling to all airfoil inner surfaces 132 of the airfoil body 122. Alternatively, in regions where impingement cooling of the inner surfaces 132, 154 is not desired or required, the cooling channels 172 can be omitted. The cooling channels 172 can extend along any desired radial extent of the hollow body 162.

[0063] As Figure 5 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 position 174 that extends around the entire periphery of the first end 164, i.e., at the meeting position, there is no opening (except possibly the cooling channels 172) between the first end 164 and the shank inner surface 154. In some 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, in all cases, this particular meeting position 174 may not be necessary. For example, in some cases, the meeting position 174 can be located radially outside the platform 150. As Figure 5 shown, although not necessary in all cases, the second end 166 of the hollow body 162 can also be integrally formed onto the inner surface 176 of the tip 140. The cooling channels 172 can optionally provide impingement cooling to the tip 140 or transfer coolant to the tip 140 for other forms of cooling.

[0064] Figure 6 Shows an enlarged cross-sectional view of the meeting position 174 of the impingement cooling structure 160 of the turbine rotor blade 120 and the shank 148 in accordance with various embodiments of the present disclosure. As Figure 5 shown. As Figure 5 and Figure 6 shown, the first end 164 of the hollow body 162 may extend substantially in the radial direction (arrow Z) relative to the meeting position 174 of the first end 164 of the hollow body 162 and the shank inner surface 154. As used herein, "substantially in the radial direction" indicates that the first end 164 extends radially away from the rotor 110 ( Figure 1 ) with a certain degree of tolerance (e.g., + / -5°). 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 meeting position 174 of the first end 164 of the hollow body 162 and the shank 148 ( Figure 6 ). In another embodiment, the shank inner surface 154 is substantially aligned in the radial direction, and the first end 164 of the hollow body is gradually curved or transitioned toward the meeting position 174 to maintain the angle α, e.g., <30°. In another embodiment, both the shank inner surface 154 and the first end 164 of the hollow body 162 are gradually curved or transitioned toward the meeting position 174. The angle α can be any desired angle and within the range of additive manufacturing, e.g., <45° from the vertical direction. To maintain structural integrity, it is desirable for the angle α to be as small as possible, e.g., <10°, <20° or <30°. Only as Figure 6 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, e.g., radially outside the meeting position 174 and radially inside the platform 150. In additional embodiments, the support structure 180 may be positioned at any location between the outer surface 170 of the hollow body 162 and the airfoil inner surface 132, the shank inner surface 154, etc. In additional embodiments, at least a portion of the support structure 180 includes a hollow support element (e.g., a grid) such that cooling flow can flow directly from the chamber 134 to the outer walls 124 and / or 126 of the airfoil body 122 ( Figure 4 ). For example, it is desirable to provide film cooling directly from the radially extending chamber 134 to certain regions of the airfoil body 122 such as the leading edge 128 and / or the trailing edge 130. The support structure 180 may include any currently 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: grid structures, 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 optionally be formed with a varying wall thickness. Different wall thicknesses may be beneficial to accommodate different CTEs, for example, between the impingement cooling structure 160 and the hotter airfoil body 122, shank 148, and / or platform 150. As Figure 5 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, platform 150, or shank 148 is shown. As described above, in certain embodiments, as Figure 7 shown, the hollow body 162 may include at least one first portion 182 having a first wall thickness W1 between its inner surface 168 and outer surface 170, and at least one second portion 184 having a second wall thickness W2 between its inner surface 168 and outer 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 thicknesses of portions 182, 184 may be any size required to meet the structural and / or thermal requirements of the location.

[0066] In another exemplary optional structure, additional supports may be desirable and / or necessary to support the integral impingement cooling structure 160 relative to the inner surfaces 132, 154. For example, additional supports may be desirable and / or necessary at the thinner wall thickness portions 184 of the impingement cooling structure 160 ( Figure 7 ). To this end, as Figure 7 shown, the turbine rotor blade 120 may also include a support 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 186 may be integrally formed with the hollow body 162 (and the remainder of the turbine rotor blade 120) to space the outer surface 170 of the hollow body 162 from, for example, the airfoil inner surface 132 between the first end 164 ( Figure 5 ) and the second end 166 of the hollow body 162. Any number of supports 186 may be provided in the thinner wall portion 184. The support 186 may include a passage 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 surfaces 132, 154. In certain embodiments, regardless of the wall thickness, the turbine rotor blade 120 may include a support 189 located on the outer 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 space the outer surface 170 of the hollow body 162 from, for example, the airfoil inner surface 132 between the first end 164 (Figure 5 )spaces the outer surface 170 of the hollow body 162 from, for example, the airfoil inner surface 132 between it and the second end portion 166. The support members 186, 189 can take any form that allows for: reducing the stress between the hotter outer 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 the desired spacing between the hollow body 162 and the inner surfaces 132, 154. The support members 186, 189 can 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 is shown that includes a reinforcement member 190 surrounding at least one of the cooling channels 172. The reinforcement member 190 can include any structural strengthening member, such as a thicker wall, etc. As Figure 4 shown, certain embodiments can also include reinforcing ribs 192 integrally formed on the inner surface 168 of the hollow body 162. Any number of reinforcing ribs 192 can be provided, and each reinforcing rib can extend any desired radial extent of the hollow body 162. The support members 186, 189, the reinforcement member 190, and / or the reinforcing ribs 192 can be integrally formed with the remainder of the turbine rotor blade 120 via additive manufacturing.

[0068] Figure 9 A cross-sectional view of a turbine rotor blade 120 is shown that includes an integral impingement cooling structure 160 and additional optional alternative integral cooling features. In one alternative embodiment, the impingement cooling structure 160 can optionally be formed with a varying spacing Z from the inner surfaces 132, 154. The spacing Z can be customized to provide a desired Z / D parameter and desired cooling at various locations. For example, the turbine rotor blade 120 can have a plurality of high heat load regions 195, i.e., regions that experience higher temperatures and require more cooling compared to other regions of the blade. In the example shown, the high heat load regions 195 include the regions: near the leading edge 128 (195A), near the trailing edge 130 on the pressure side outer wall 124 (195B), and on the suction side outer wall 126 downstream of the leading edge 128 (195C). At the high heat load regions 195, a first spacing Z1 can be employed between the integral impingement cooling structure 160 and the inner surfaces 132, 154 at the high heat load regions 195, while a larger second spacing Z2 is employed at other locations without such high heat loads. Thus, when necessary (i.e., at the high heat load regions 195), the first spacing Z1 can be used to provide more cooling, where the spacing is increased to the second larger spacing Z2 between the impingement cooling structure 160 and the inner surfaces 132, 154 for lower heat load regions. As Figure 9As shown, the larger second spacing Z2 may allow the coolant flow 136 to limit or reduce heat absorption as it moves downstream toward the trailing edge 130, thereby making the coolant flow 136 cooler and having a greater heat absorption capacity for downstream regions such as the spiral cooling channels 200 and / or the pin arrays 206 (described herein). The transition between the spacings Z1 and Z2 can occur at any desired rate, such as gradually over a relatively long distance, suddenly at a particular location, or at any rate in between. The second spacing Z2 can be any value between, for example, 1.01 times and 3.00 times the first spacing Z1. The Z / D parameter can be customized for each region 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 regions.

[0069] Figure 9 Also shown is a turbine rotor blade 120 including one or more post-impingement target features 196 located on the inner surface 132. The post-impingement target features 196 can include any currently known or later developed structure located on the inner surface 132 to facilitate cooling. In the example shown, the impingement target feature 192 includes bumps, but it can also include any structure. In one case, the hollow body 162 can include local bulges 198 to match the profile of the post-impingement target features 196 and thus maintain the spacing Z (i.e., Z1 as shown). Although two pairs of target features 196 and bulges 198 are shown, any number can be employed. In one embodiment, the post-impingement target features 196 can also optionally include additional integral cooling features, such as but not limited to film cooling holes 199. After impingement using the post-impingement target features 196 (i.e., its inner surface 132), the film cooling holes 199 direct the coolant flow 136 to form a cooling film 201 above the sidewalls 124, 126. Any number of film cooling holes 199 can be applied within each post-impingement cooling feature 196.

[0070] Figure 10 A first radial cross-sectional view along Figure 4 the line of sight 10-10 in Figure 11 is shown, and Figure 4 a second radial cross-sectional view along Figure 10 the line of sight 11-11 in Figure 10 is shown, the latter being in a slightly different plane and in the opposite direction. As Figures 10 to 11 Figure 10 shown, in certain embodiments, the hollow body 162 has a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to be inserted through the open tip of the airfoil body. Additionally, the hollow body 162 can have a smaller chordal width WC1 near the tip 140 than at the shank 148. Most conventional impingement inserts have the opposite chordal width arrangement to allow them to behas alternating wider chordal widths and narrower chordal widths WC1 on the upper and lower pages). Thus, the axial rear end 194 of the hollow body 162 can vary in chordal position along the radial span of the hollow body 162. In this way, the impingement cooling structure 160 can have a shape (chordal width WC1) that bends along its radial span to be evenly spaced from the airfoil inner surface 132 and / or the shank inner surface 154( Figure 5 ), regardless of the shape of the latter. Conventional impingement inserts do not provide such features.

[0071] As Figure 4 , Figure 10 and Figure 11 shown, the airfoil body 122 also includes at least one chordally extending spiral cooling channel 200 that extends from the airfoil inner surface 132 behind the hollow body 162 toward the trailing edge 130. As Figure 10 and Figure 11 best shown, each chordally extending spiral cooling channel 200 can have the same chordal width WC2, for example shorter than the chordal width WC1 of the hollow body 162. The airfoil body 122 can also include a plurality of radially spaced trailing edge cooling channels 202 that extend through the trailing edge 130 (i.e., from the spiral cooling channels 200). Each of the plurality of trailing edge cooling channels 202 has the same chordal width WC3, that is, along the radial span of the turbine rotor blade 120. As Figure 10 and Figure 11 shown, the space 204 between the trailing edge cooling channels 202 and the spiral cooling channels 200 has a varying chordal width WC4, that is, along the radial span of the turbine rotor blade 120. The turbine rotor blade 120 can also include a pin group 206 located at the front end of the plurality of trailing edge cooling channels 202( Figure 10 its right side in Figure 11 ), its left side in Figure 10 ), and the rear end of the chordally extending spiral cooling channels 200( Figure 11 its left side in Figures 10 to 11 shown, the pin group 206 can have a varying chordal width WC4 along its radial span. Figure 10 and Figure 11 show various cooling features, including but not limited to: film cooling via the openings 208, primary chord impingement cooling via the impingement cooling structure 160, cooling near the trailing edge 130 via the spiral cooling channels 200, and trailing edge pin group cooling via the pin group 206.

[0072] Additive manufacturing (AM) includes a variety of processes for producing components by successive layering of materials rather than by removal of materials. Thus, additive manufacturing can form complex geometries such as those described herein with respect to turbine rotor blade 120 without the use of any kind of tooling, dies, or fixtures and with little or no waste of material. Instead of machining components from solid material blanks, many of which are cut away and discarded, the only materials used in additive manufacturing are those needed to form the component. Additive manufacturing techniques generally involve taking a three-dimensional computer-aided design (CAD) file of the component to be formed (e.g., turbine rotor blade 120), electronically slicing the component into layers (e.g., 18 - 102 microns thick), and creating a file with two-dimensional images (including vectors, images, or coordinates) of each layer. The file can then be loaded into a preparation software system that interprets the file such that the component 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 dispensed, sintered, formed, deposited, etc. to form the component. Although other manufacturing processes such as casting can also be employed, turbine rotor blade 120 can advantageously be made by additive manufacturing.

[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, metal powder layers are sequentially melted together to form the component. More specifically, after being evenly distributed over a metal powder bed using an applicator, fine metal powder layers are sequentially melted. Each applicator includes an applicator element in the form of a die lip, brush, doctor blade, or roller made of metal, plastic, ceramic, carbon fiber, or rubber, which spreads the metal powder evenly above the build platform. The metal powder bed can be moved in the vertical axis. The process takes place in a processing chamber with precisely controlled atmosphere. Once each layer is formed, each two-dimensional slice of the component geometry can be fused by selectively melting the metal powder. The melting can be performed by a high-power melting beam (such as a 100-watt ytterbium laser) to fully weld (melt) the metal powder to form solid metal. The melting beam is moved in the X-Y direction using scanning mirrors, and its intensity is sufficient to fully weld (melt) the metal powder to form solid metal. The metal powder bed can be lowered for each subsequent two-dimensional layer, and the process is repeated until the component is fully formed. To form some larger blades more quickly, some metal additive manufacturing systems employ a pair of high-power lasers that work together to form the blade. Here, a method of fabricating a turbine rotor blade 120 can 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 that includes a plurality of integral material layers.

[0074] The turbine rotor blade 120 can be made of a metal capable of withstanding the environment employed, which can include pure metals or alloys. In one example, the metal can include almost 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., X) available from Haynes International, Inc. or nickel-chromium-cobalt-molybdenum alloy (NiCrCoMo) (e.g., Haynes282 available from Haynes International, Inc.), etc. In another example, the metal can include almost any metal, such as but not limited to: tool steel (e.g., H13), titanium alloy (e.g., Ti 6 Al 4V), stainless steel (e.g., 316L), cobalt-chromium alloy (e.g., CoCrMo), and aluminum alloy (e.g., AlSi 10 Mg).

[0075] Compared to conventional impact inserts, the first end 164 of the hollow body 162 is integrally formed onto the inner surface 154 of the shank, i.e., formed via additive manufacturing. Thus, the outer surface 170 of the hollow body 162 can be made to be uniformly spaced from the inner surface between the first end 164 and the second end 166 of the hollow body 162, regardless of the curvature of, for example, 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 perimeter and radial span to provide impingement cooling over the entire blade and not just at its leading edge. Thus, the overall impingement cooling structure 160 allows for maximum impingement coverage to be achieved without the sacrifice typically associated with impact inserts and can have multiple variable cooling features. For example, the turbine rotor blade 120 can have: a variable chordwise 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 address 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), can also be provided and customized around each cooling channel 172 to optimize impingement cooling. The turbine rotor blade 120 can also include axial ventilation through the trailing edge 130, as described with respect to Figures 10 to 11 above.

[0076] C. Platform with Cooling Channels of Nonlinear Configuration

[0077] See Figures 12 to 18 , which shows another overall feature according to an embodiment of the present disclosure. Similar to the previous embodiment, the turbine rotor blade 120 can include an airfoil body 122 that has a radially extending chamber 134 for receiving a coolant flow 136. As Figure 3 and Figures 10 to 13 best shown, the platform 150 extends laterally outward relative to the airfoil body 122 and terminates at at least one bevel 230 (e.g., Figure 11 , Figure 12 ). Figure 12 Shows a perspective transparent view of the pressure side 232 of the platform 150, and Figure 13 shows a top-down transparent view of the suction side 231 of the platform 150. As Figure 12 and Figure 13As 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 can take any of a variety of forms. In one example, where the turbine rotor blade 120 includes an impingement cooling structure 160 located within a radially extending chamber 134, the source of coolant 236 leading to the cooling circuit 234 can provide coolant after passing through the impingement cooling structure 160, i.e., the coolant is post-impingement coolant. In another embodiment, the source of coolant 236 can be the radially extending chamber 134. For example, where no impingement cooling structure 160 is provided or an impingement cooling structure is provided radially outside of the platform 150, the source of coolant 236 leading to the cooling circuit 234 can provide coolant directly from the radially extending chamber 134. Other sources of coolant 236 can also be used, such as the annulus space portion between the shanks 148 of adjacent turbine rotor blades 120. The cooling circuit 234 can take any form now known or later developed. In Figure 13 the example shown, the cooling circuit 234 includes a sinusoidal path through the platform 150. In contrast, in Figure 12 it, the cooling circuit 234 includes a bent tube 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 desired.

[0078] The turbine rotor blade 120 also includes cooling channels 240 that extend from the cooling circuit 234 through the surface 242 of the bevel 230, i.e., to cool the bevel 230 and other structures. The cooling channels 240 are located within the platform 150 and are in fluid communication with the cooling circuit 234. Compared to conventional linear cooling channels, the cooling channels 240 extend from the cooling circuit 234 through at least one bevel 230 of the platform 150 to an outlet in a non-linear configuration, thereby providing improved cooling compared to linear cooling channels. For example, in Figure 12 it, the cooling channels 240 have a (slightly) curved shape. Any number of cooling channels 240 can be employed to provide the desired cooling. Additionally, the cooling channels can have any desired uniform or non-uniform cross-sectional shape and can be spaced uniformly or non-uniformly to provide the desired cooling. The non-linear configuration is made possible, for example, by additive manufacturing. As described above, the airfoil body 122 and the platform 150 (including the components defining the cooling channels 240) can include multiple monolithic material layers.

[0079] Figure 14 An enlarged cross-sectional view of the bevel 230 is shown. In some embodiments, as Figure 14As shown, the inclined surface 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 dashed lines), which seals with a damping pin seat of an adjacent turbine rotor blade (not shown). When provided, the cooling channel 240 may extend through the extension member 244. In this regard, the cooling channel 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 inclined surface 230, i.e., in a more sharply curved or more turning bent shape than Figure 12 more sharply or more turning bent shape.

[0080] The cooling channel 240 may adopt any one of a variety of non-linear configurations to provide desired cooling. The non-linear configuration (e.g., a bent 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 channels 240 may all have the same shape to provide the same cooling properties at each location where the cooling channels are provided, or the shape of the cooling channels may vary within the platform 150 to provide customized cooling for each location where the cooling channels are provided. In addition to Figure 12 and Figure 13 the bent shapes shown, in Figure 15 another embodiment shown, the cooling channel 240 may have a spiral (screw) shape, i.e., having a plurality of spiral coils 250. Any number of spiral coils 250 may be used for each cooling channel 240. As Figure 13 and Figure 16 shown, the cooling channel 240 may have at least one first turn 252 ( Figure 16 ) in a first direction FD and at least one second turn 254 ( Figure 16 ) in an opposite second direction SD, thereby forming a generally zigzag path. Any number of first turns 252 and second turns 254 ( Figure 16 ) may be used for each cooling channel 240. As Figure 16 shown on the left side, the amplitude A of each turn 252, 254 may be consistent so as to form a sine shape with at least one first turn and second turn having the same amplitude A. Alternatively, as Figure 16 shown on the right side, the amplitude of each turn 252, 254 may be inconsistent so as to form a more random zigzag path with the turns 252, 254. Also as Figure 16 shown on the right side, the inputs 260 and outlets 262 of each cooling channel 240 do not need to be aligned. In Figure 17 another embodiment shown, the cooling channel 240 may have a plurality of branches 264, such as a tree-like. Any branch configuration may be adopted.

[0081] Figure 18 shows an embodiment in which the cooling channel 240 has a curved shape, e.g., flatter within the platform 150 than Figure 12 more flat. Figure 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 desired cooling to the platform 150 and / or film cooling to the inclined surface 230. Although shown separately, any one of the cooling channel examples or aspects thereof can be combined with other examples.

[0082] The cooling circuit 234 and the cooling channel 240 can be separately disposed in the pressure side 232 of the platform (only Figure 13 ), separately disposed in the suction side 231 of the platform 150 (only Figure 12 ), or disposed in both sides 231, 232 of the platform 150 ( Figures 12 to 13 ). If disposed on only one side of the platform 150, any other conventional structure can be disposed on the other side of the platform. In the latter case, as Figure 12 and Figure 13 collectively shown, the cooling circuit 234 can 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 can be separate or fluidly coupled. In this regard, the inclined surface 230 includes a suction side inclined surface 230SS and a pressure side inclined surface 230PS. Here, the cooling channel 240 in the platform 150 can include: at least one first cooling channel 240 that is fluidly connected to the first portion 234SS of the cooling circuit and exits the suction side inclined surface 230SS, and at least one second cooling channel 240 that is fluidly connected to the second portion 234PS of the cooling circuit 234 and exits the pressure side inclined surface 230PS.

[0083] Compared with the linear coolant channels of conventional drilling, the non-linear cooling channels 240 allow the coolant to be directed where needed in the platform 150. The additive manufacturing of the coolant channels 140 enables the coolant channels to have a variety of non-linear configurations that directly cool when necessary and provide enhanced cooling through their shape.

[0084] D. Airfoil with Coolant Delivery Channels

[0085] See Figure 2 and Figures 19 to 22 , another overall feature portion of an embodiment according to the present disclosure includes a wing portion 280 having a coolant delivery channel therein. Figure 19 shows a radial cross-section through the turbine rotor blade 120 including the wing portion 280,Figure 20 A transparent perspective view of a turbine rotor blade 120 including an airfoil section 280 is shown, Figure 21 an axial view of a set of turbine rotor blades 120A-C including the airfoil section 280 is shown, and Figure 22 a top view of the turbine rotor blade 120 including the airfoil section 280 is shown. Further referring to Figure 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 (collectively or individually the turbine rotor blade 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 Figures 19 to 20 shown, each turbine rotor blade 120 may include an airfoil body 122 that includes a concave pressure side outer wall 124 ( Figure 3 ) and a convex suction side outer wall 126 ( Figure 3 ) connected along a leading edge 128 and a trailing edge 130 ( Figure 3 ). The turbine rotor blade 120 may also include a shank 148 located at a radially inner end 152 of the airfoil body 122. Additionally, the turbine rotor blade 120 includes at least one airfoil section 280 that extends laterally from at least one side surface 282, 284 of the shank 148.

[0086] As Figure 2 shown in one of the nozzle - blade junctions, there is an opening 286 at the junction between an adjacent nozzle 112 and the turbine rotor blade 120, and this opening may allow the hot working fluid to leave the hot gas path and enter the wheel space 300 of the turbine 108. To limit such leakage of the hot gas, the turbine rotor blade 120 typically includes an axially projecting airfoil seal 280, also simply referred to as an "airfoil section". The airfoil section 280 mates with a projecting segment or "blocking portion" 288 extending from the nozzle 112. The airfoil section 280 and the blocking portion 288 overlap (or nearly overlap), but do not contact each other, thereby restricting fluid flow.

[0087] Turning to Figures 19 to 22 , according to an embodiment of the present disclosure, the turbine rotor blade 120 may further include a coolant transfer passage 290 defined by at least one airfoil section 280. For example, for the first turbine rotor blade 120A ( Figure 21 ), the coolant transfer passage 290 is fluidly coupled to a first wheel space portion 292 defined between a shank 148A ( Figure 21 ) and a first adjacent shank 148B of a first adjacent turbine rotor blade 120B ( Figure 21 ) and to a second wheel space portion 292 defined between the shank 148A ( Figure 21 ) and a second adjacent turbine rotor blade 120C ( Figure 21) the second round space portion 294 between the adjacent shanks 148C. As can be seen by observing Figure 2 、 Figure 21 and Figure 22 As best shown, each of the round space portions 292, 294 is part of the round space 300. The round space 300 is defined as: circumferentially defined between the shanks 148A-C of adjacent turbine rotor blades 120A-C ( Figure 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 round space portions 292, 294 are portions of the round space 300 axially beside the shank 148 of a particular blade.

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

[0089] Any number of fins 280 can be employed. In one example, as Figure 19 、 Figure 20 and Figure 22 shown, a first fin 280 extends laterally from the first side 282 of the shank 148, and a second fin 280 extends laterally from the second opposite side 284 of the shank 148. In another example, as Figure 23 shown, a first pair of radially spaced fins 280 can extend laterally from the first side 282 of the shank 148 and none extend from the side 284 of the shank 148. In another embodiment, as Figure 24As shown, a first pair of radially spaced wings 280 may extend laterally from a first side 282 of the shank 148, and a second pair of radially spaced wings 280 may extend laterally from a second opposite side 284 of the shank 148. In any case, each wing 280 may include a corresponding coolant transfer passage 290. Alternatively, although each wing 280 is shown as including a coolant transfer passage 290, select wings may not include a coolant transfer passage.

[0090] The coolant transfer passage 290 allows wheel space coolant 316 ( Figure 22 ) to move between wheel space portions 292, 294, allows the wings 280 to be cooled, and reduces the weight of the turbine rotor blade 120.

[0091] E. Hollow Vane Mount with Grille Support Structure

[0092] See Figure 19 、 Figure 20 、 Figure 25 and Figure 26 According to another overall feature of embodiments of the present disclosure, a hollow blade mount 146 is included. In this embodiment, a root 144 is provided that 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, for example, defined 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 may have any desired internal shape, for example, radially expanded as Figure 25 shown. The blade mount 146 may have any presently known or later developed external shape configured for mounting to a rotor wheel 147 ( Figure 2 ) coupled to the rotor 110 ( Figure 21 ), such as a dovetail or fir tree shape.

[0093] The turbine rotor blade root 144 may further include a grid support structure 340 disposed within the hollow interior 330 of the blade mount 146. The grid support structure 340 may take various forms of hollow support structures. In one example, the grid 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 grid support structure 340 may be made by additive manufacturing. The shank 148 and the blade mount 146 may thus include a plurality of integral material layers.

[0094] The root 144 according to this embodiment, i.e., the root having the grid support structure 340, may also include a platform 150, as described herein with respect to Figures 12 to 18 above. As described above, the platform 150 is positioned radially outside the shank 148 and extends laterally outward relative to the shank, terminating at at least one inclined surface 230. The platform 150 may include a cooling circuit 234 that is defined within the platform and is in fluid communication with a source of coolant flow (e.g., the radially extending chamber 134). Cooling channels 240 ( Figures 12 to 18 ) may be defined within the platform 150 and are in fluid communication with the cooling circuit 234. As described above, the cooling channels 240 extend 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 channels 240 extend. The cooling channels 240 may have: a helical shape ( Figure 15 ); at least one first turn in a first direction and at least one second turn in an opposite second direction ( Figure 16 ); a plurality of branches ( Figure 17 ); or a curved shape (e.g., Figure 12 , Figure 14 , Figure 18 ).

[0095] The root 144 according to this embodiment, i.e., the root having the grid support structure 340, may also include a wing 280 that extends laterally from at least one side of the shank 148, as described herein with respect to Figures 19 to 24 above. As described above, the coolant transfer channel 290 may be defined through the wing 280. As Figure 21 shown, the coolant transfer channel 290 fluidly couples a first wheelspace portion 292 defined between the shank 148A and a first adjacent shank 148B of a first adjacent turbine rotor blade root 144B and a second wheelspace portion 294 defined between the shank 148A and a second adjacent shank 148C of a second adjacent turbine rotor blade root 144C. The coolant transfer channel 290 includes a first open end 310 that is in fluid communication with the first wheelspace portion 292 and a second open end 312 that is in fluid communication with the second wheelspace portion 294. As Figure 22 shown, the first open end 310 and the second open end 312 may face circumferentially relative to the shank 148. The coolant transfer channel 290 may be fluidly isolated from the radially extending chamber 134 within the shank 148.

[0096] The root 144 according to this embodiment, i.e., the root having the grid support structure 340, may also include both the platform 150 and the wing 280, as described herein. Additive manufacturing allows for the formation of a root 144 having a shank 148, a hollow blade mount 146, a grid support structure 340, and a platform 150 and / or a wing 280, thereby forming multiple integral material layers for any feature provided.

[0097] The root 144 including the integral grid support structure 340 in the hollow interior 330 of the blade mount 146 provides additional cooling for 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, approximating language may be used to modify any quantitative representation that may permit variation without resulting in a change in the basic function associated therewith. Thus, a value modified by one or more terms, such as "about", "approximately", and "substantially", is not limited to the specified exact value. In at least some instances, the approximating 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 recognized and include all subranges subsumed therein. The "about" applied to a particular value in a range applies to both end values, and may indicate + / −10% of one or more of said values, unless otherwise dependent on the precision of the instrument measuring the value.

[0100] All structural, material, act, and equivalent elements corresponding to the means or step plus function elements in the following claims are intended to include any structure, material, or act for performing the function in combination with other claimed elements for specifically claimed protection. The description of the present disclosure has been given for purposes of illustration and description, but 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 spirit of the present disclosure. The embodiments were chosen and described in order to best explain the principles of the present disclosure and its practical application, and to enable others of ordinary skill in the art to understand the present disclosure with various modifications suitable for the particular use contemplated.

Claims

1. A turbine rotor blade (120), comprising: 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 and a trailing edge (128, 130), the outer walls (124, 126) defining a radially extending chamber (134) for receiving a coolant (236) flow; a platform (150) that extends laterally outward relative to the airfoil body (122) and terminates 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 the coolant (236) flow; at least one cooling channel (172, 240) defined within the platform (150) and in fluid communication with the cooling circuit (234), the at least one cooling channel (172, 240) extending in a non-linear configuration from the cooling circuit (234) through the at least one bevel (230, 230PS, 230SS) of the platform (150) to an outlet (262); and an impingement cooling structure (160) located within the radially extending chamber (134), wherein the source of the coolant (236) leading to the cooling circuit (234) provides the coolant (236) after passing through the impingement cooling structure (160).

2. The turbine rotor blade (120) according to claim 1, wherein the source of the coolant (236) flow leading to the cooling circuit (234) directly provides the coolant (236) from the radially extending chamber (134).

3. A turbine rotor blade (120), comprising: 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 and a trailing edge (128, 130), the outer walls (124, 126) defining a radially extending chamber (134) for receiving a coolant (236) flow; a platform (150) that extends laterally outward relative to the airfoil body (122) and terminates 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 the coolant (236) flow; at least one cooling channel (172, 240) defined within the platform (150) and in fluid communication with the cooling circuit (234), the at least one cooling channel (172, 240) extending in a non-linear configuration from the cooling circuit (234) through the at least one bevel (230, 230PS, 230SS) of the platform (150) to an outlet (262), wherein the at least one bevel (230, 230PS, 230SS) includes an extension member (244), and wherein the at least one cooling channel (172, 240) extends through the extension member (244).

4. A turbine rotor blade (120), comprising: 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 and a trailing edge (128, 130), the outer walls (124, 126) defining a radially extending chamber (134) for receiving a flow of coolant (236); a platform (150) that extends laterally outward relative to the airfoil body (122) and terminates 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 the coolant (236) flow; at least one cooling channel (172, 240) defined within the platform (150) and in fluid communication with the cooling circuit (234), the at least one cooling channel (172, 240) extending in a non-linear configuration from the cooling circuit (234) through the at least one bevel (230, 230PS, 230SS) of the platform (150) to an outlet (262), wherein the at least one cooling channel (172, 240) has a helical shape.

5. A turbine rotor blade (120), comprising: 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 and a trailing edge (128, 130), the outer walls (124, 126) defining a radially extending chamber (134) for receiving a flow of coolant (236); a platform (150) that extends laterally outward relative to the airfoil body (122) and terminates 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 the coolant (236) flow; at least one cooling channel (172, 240) defined within the platform (150) and in fluid communication with the cooling circuit (234), the at least one cooling channel (172, 240) extending in a non-linear configuration from the cooling circuit (234) through the at least one bevel (230, 230PS, 230SS) of the platform (150) to an outlet (262), wherein the at least one cooling channel (172, 240) has at least one first turn (252) in a first direction and at least one second turn (254) in an opposite second direction.

6. The turbine rotor blade (120) according to claim 5, wherein the at least one cooling channel (172, 240) has a sinusoidal shape, and wherein the at least one first turn and second turn (252, 254) have equal amplitudes.

7. The turbine rotor blade (120) according to claim 1, wherein the at least one cooling channel (172, 240) has a plurality of branches (264).

8. A turbine rotor blade (120) comprising: 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 and a trailing edge (128, 130), the outer walls (124, 126) defining a radially extending chamber (134) for receiving a flow of coolant (236); a platform (150) that extends laterally outward relative to the airfoil body (122) and terminates 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 the coolant (236) flow; at least one cooling channel (172, 240) defined in the platform (150) and in fluid communication with the cooling circuit (234), the at least one cooling channel (172, 240) extending in a non-linear configuration from the cooling circuit (234) through the at least one bevel (230, 230PS, 230SS) of the platform (150) to an outlet (262), wherein the at least one cooling channel (172, 240) has a curved shape, and wherein the outlet (262) of the at least one cooling channel (172, 240) meets the bevel (230, 230PS, 230SS) of the platform (150) at an angle less than 15°.

9. The turbine rotor blade (120) according to any one of claims 1 to 8, wherein the airfoil body (122) and the platform (150) comprise a plurality of monolithic material layers.

10. The turbine rotor blade according to any one of claims 1 to 8, wherein the cooling circuit (234) includes a first portion (234SS) located on the suction side (284) of the platform (150) and a second portion (234PS) located on the pressure side (232) of the platform (150), wherein the at least one bevel (230, 230PS, 230SS) includes a suction side (284) bevel (230, 230PS, 230SS) and a pressure side (284)(232) bevel (230, 230PS, 230SS), and wherein the at least one cooling channel (172, 240) defined in the platform (150) comprises: At least one first cooling channel (172, 240), the at least one first cooling channel being in fluid communication with the first portion (234SS) of the cooling circuit (234) and exiting the suction side (231) inclined surface (230, 230PS, 230SS); and at least one second cooling channel (172, 240), the at least one second cooling channel being in fluid communication with the second portion (234PS) of the cooling circuit (234) and exiting the pressure side (232) inclined surface (230, 230PS, 230SS).

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