Additively manufactured turbine rotor blade with integrated impingement sleeve

Abstract

The present invention is entitled "Turbine rotor blade with integral impingement sleeve manufactured by additive manufacturing." A turbine rotor blade (120) is disclosed, the turbine rotor blade being manufactured by additive manufacturing and comprising an airfoil body (122) having a radially extending chamber (134) for receiving a coolant flow (136), a tip (140) at a radially outer end (142) of the airfoil body (122), and a shank (148, 148A) at a radially inner end (152) of the airfoil body (122). The radially extending chamber (134) extends at least partially into the shank (148, 148A) to define a shank inner surface (154). An integral impingement cooling structure (160) is located within the radially extending chamber (134). The integral impingement cooling structure (160) allows an outer surface (170) of the hollow body (162) thereof to be uniformly spaced from the airfoil inner surface (132) despite the curvature of the chamber (134). A turbine rotor blade (120) has impingement cooling (160) across the blade (120).

Description

Background Art

[0001] The present disclosure relates integrally to turbines, and more particularly to turbine rotor blades having multiple integral, additively manufactured features, such as an integral impingement sleeve, nonlinear cooling channels in the platform, angel wings with coolant delivery channels, and a hollow dovetail with an integral grid support structure.

[0002] A turbine includes multiple turbine rotor blades coupled to a rotor. A working fluid, such as steam or burning fuel, is pushed against the blades, forcing them to turn the rotor. 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. An impingement insert or sleeve comprises a hollow body having cooling channels in its walls that allow coolant to be delivered through the cooling channels to impinge or impinge on the surface to be cooled. Impingement inserts are used in various hot gas (HGP) components in turbomachinery, such as turbine rotor blades, for example, to increase the cooling performance of the cooling circuit therein. One challenge with impingement inserts is positioning the impingement insert within a tapered or curved cavity in the HGP component close enough to achieve high cooling performance, but not so close that cooling is ineffective. One indicator 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 interior surface of the HGP component to the diameter D of the cooling channel (hole) 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] Impingement cooling is typically not provided by the insert if the necessary spacing distance cannot be achieved. For example, if the cavity curvature in the HGP assembly is so significant that the impingement insert cannot be made thin enough or curved enough to adhere to the necessary spacing distance, impingement cooling cannot be provided. One approach to addressing this challenge involves arranging the impingement insert in multiple flexible longitudinal sections to facilitate their insertion into the HGP component. However, the necessity of sequentially positioning and coupling multiple insert sections together or to the HGP component increases manufacturing complexity, time, and cost. Flexible impingement insert sections also lack adjacent elements around their perimeter (i.e., transversely (cross-sectionally)), which can detract from cooling performance at discontinuities.

[0005] Impingement cooling has been applied in a limited manner to rotating turbine rotor blades in turbines, such as at their leading edges. However, because the centrifugal forces experienced by the rotating blades cause coolant to reach the radially outer tips of the blades as they rotate, impingement cooling has not been applied more broadly across the entire inner surface of the turbine rotor blades, making impingement cooling less effective.

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

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

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

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

[0010] A third aspect of the present disclosure provides a method comprising: sequentially generating material layers and applying a heat source to sinter the material layers to form: a wing-shaped body comprising a concave pressure side outer wall and a convex suction side outer wall connected along a leading edge and a trailing edge, the outer wall having a wing-shaped inner surface, the wing-shaped inner surface defining a radially extending chamber for receiving a coolant flow; and an impingement cooling structure within the radially extending chamber, the integrated impingement cooling structure comprising: a hollow body comprising 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 fluidly connected to the radially extending chamber to allow a coolant flow to pass from the inner surface of the hollow body to at least impinge on the wing-shaped inner surface, wherein the outer surface of the hollow body is evenly spaced from the wing-shaped inner surface between the first end and the second end of the hollow body.

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

[0012] A fifth aspect of the present disclosure provides a turbine rotor blade manufactured by additive manufacturing, comprising: an airfoil body comprising a concave pressure side outer wall and a convex suction side outer wall connected along a leading edge and a trailing edge, 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 in at least one inclined surface; a cooling circuit defined within the platform and fluidly connected to a source of the coolant flow; and at least one cooling channel defined in the platform and fluidly connected to the cooling circuit, the at least one cooling channel extending from the cooling circuit in a nonlinear configuration to exit through at least one inclined surface of the platform.

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

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

[0016] A ninth aspect relates to a turbine rotor blade root comprising a shank having a radially extending chamber defined therein; a blade mount at a radially inner end of the shank, the blade mount having a hollow interior defined therein, the hollow interior being fluidically connected to the radially extending chamber; and a grating support structure disposed in the hollow interior of the blade mount.

[0017] A tenth aspect includes a turbine rotor blade root, the turbine rotor blade root comprising: a shank having a radially extending chamber defined therein; a blade mounting frame located at a radially inner end of the shank, the blade mounting frame having a hollow interior defined therein, the hollow interior being in fluid communication with the radially extending chamber; a grating support structure disposed in the hollow interior of the blade mounting frame; at least one angel wing extending laterally from at least one side of the shank; and a coolant transfer passage defined by the at least one angel wing, the coolant transfer passage being fluidly coupled to a first impeller space portion and a second impeller space portion, the first impeller space portion being defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade root, and the second impeller space portion being defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade root.

[0018] An eleventh aspect relates to a turbine rotor blade root, comprising: 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 grating support structure disposed within the hollow interior of the blade mount; at least one angel wing extending laterally from at least one side of the shank; and a coolant transfer passage defined by the at least one angel wing, the coolant transfer passage being fluidly coupled to a first impeller space portion and a second impeller space portion, the first impeller space portion being defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade root, and the second impeller space portion being defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade root.

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

[0047] It should be noted that the drawings of the present disclosure are not necessarily drawn to scale. The drawings are intended to depict only typical aspects of the present disclosure and therefore should not be considered to limit the scope of the present disclosure. In the drawings, similar numbers represent similar 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, common industry terms will be used and adopted, where possible, in a manner consistent with their accepted meanings. Unless otherwise indicated, such terms should be given a broad interpretation consistent with the context of this application and the scope of the appended claims. One of ordinary skill in the art will understand that several different or overlapping terms may often be used to refer to a particular component. An object that may be described herein as a single part may include multiple components and be referenced in another context as consisting of multiple components. Alternatively, an object that may be described herein as comprising multiple components may be referred to elsewhere as a single part.

[0049] In addition, several descriptive terms may be used regularly throughout this document, and it may be 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 that indicate the direction relative to the flow of a fluid, such as the working fluid through a turbine, or, for example, the flow of air through a combustor or the coolant through one of the turbine's components. The term "downstream" corresponds to the direction of fluid flow, and the term "upstream" refers to the direction opposite to the flow. Without any further specificity, the terms "front" and "rear" refer to directions, where "front" refers to the front end (compressor end) of the turbine or the front section of a component thereof, and "rear" refers to the rear end (turbine end) of the turbine or the rear end of a component thereof. In the accompanying drawings, forward and rearward are generally represented by the X direction. It is often necessary to describe parts that are in different radial positions relative to the central axis. The term "radial" refers to movement or position perpendicular to, for example, the axis of the turbine rotor. In cases such as these, if a first component resides closer to the axis than a second component, this document will describe the first component as being "radially inward" or "inboard" of the second component. On the other hand, if a first component resides farther from the axis than a second component, the present invention may describe the first component as being "radially outward" or "outboard" of the second component. In the accompanying drawings, the radial direction is generally represented by the Z direction. The term "axial" refers to movement or position parallel to the axis (i.e., the turbine rotor axis). Finally, the term "circumferential" refers to movement or position around the axis. Although the circumferential direction is not shown as curved in the legends of the accompanying drawings, it is generally represented by the Y direction in the accompanying drawings. It should be understood that such terms can be applied relative to the rotor axis of the turbine.

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

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

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

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

[0054] A. Introduction

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

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

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

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

[0059] B.Integrated impingement cooling structure

[0060] In certain embodiments, a turbine rotor blade 120 may include, among other things, an airfoil body 122 and an integral feature in the form of an integral impingement cooling structure 160 therein. The impingement cooling structure is not an insert, but is manufactured integrally 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 also 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 extend at least partially into the shank 148 to define a shank inner surface 154. An integral impingement cooling structure 160 is located within the radially extending chamber 134 and may include a hollow body 162 having a first end 164, a second end 166, an interior surface 168, and an exterior 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 coolant flow from the interior surface 168 of the hollow body 162 to impinge on at least the inner surface of the airfoil body 122. In contrast to conventional impingement inserts, the first end 164 of the hollow body 162 is integrally formed to the shank inner surface 154, i.e., additively manufactured. Consequently, regardless of the curvature of the airfoil inner surface 132, the exterior surface 170 of the hollow body 162 may be uniformly spaced from the airfoil inner surface 132 between the first and second ends 164, 166 of the hollow body 162. In another embodiment, a non-uniform but customized spacing distance may be employed to provide varying levels of impingement cooling, heat absorption, and / or reuse. For example, a closer spacing distance may be employed where more impingement cooling is desired, while a wider spacing distance may be employed where less impingement cooling is desired. Additionally, the hollow body 162 may have cooling channels around its entire circumference and radial span to provide impingement cooling across the entire blade, not just at its leading edge. Thus, an integrated impingement cooling structure allows for maximum impingement coverage with limited sacrifices typically associated with impingement inserts, and may have multiple variable cooling features. For example, a turbine rotor blade may have: a variable chord-wise width behind the impingement cooling structure or its pin bank; impingement cooling structure wall thickness tailored for different cooling loads; and different supports to account for different coefficients of thermal expansion (CTE) between the airfoil body and the impingement cooling structure.

[0061] like Figure 4 As shown, and as Figure 5As shown in the radial circumferential cross-sectional view of FIG, the turbine rotor blade 120 may include an impingement cooling structure 160 within the 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. Figure 5 With the curved airfoil inner surface 132 shown, it is not possible to have impingement inserts evenly spaced from the inner surface along the entire radial span of the blade. To address this challenge, an impingement cooling structure 160 according to an embodiment of the present disclosure is integrally formed with the remainder of the turbine rotor blade 120 by additive manufacturing.

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

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

[0064] Figure 6 FIG. 1 shows the junction 174 of the impingement cooling structure 160 and the shank 148 of the turbine rotor blade 120 according to various embodiments of the present disclosure. Figure 5 ) is an enlarged cross-sectional view. Figure 5 and Figure 6 As shown, the first end 164 of the hollow body 162 may extend substantially in a radial direction (arrow Z) relative to a location 174 where the first end 164 of the hollow body 162 meets the shank inner surface 154. As used herein, "substantially in a radial direction" means that the first end 164 is spaced apart from the rotor 110 ( Figure 1 ) extends radially. In contrast, at least a portion of the shank inner surface 154 extends at an angle α relative to the radial direction Z from the junction 174 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 aligned substantially in a radial direction, and the first end 164 of the hollow body is gradually curved or transitioned toward the confluence location 174 to maintain the angle α (for example) less than 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 confluence location 174. The angle α can be any angle desired and within the scope of additive manufacturing, for example, an angle less than 45° from the vertical. To maintain structural integrity, it is desirable that the angle α be as small as possible, for example, less than 10°, less than 20°, or less than 30°. Only as Figure 6 As shown, in certain embodiments, the support structure 180 may be positioned between the first end 164 of the hollow body 162 and the shank inner surface 154, for example, radially outward of the junction location 174 and radially inward of the platform 150. In other embodiments, the support structure 180 may be positioned anywhere between the exterior surface 170 of the hollow body 162 and the airfoil inner surface 132, the shank inner surface 154, etc. In other embodiments, at least a portion of the support structure 180 includes hollow support elements (e.g., a grid) to enable cooling flow to be directed from the chamber 134 to the outer walls 124 and / or 126 ( Figure 4 For example, it may be desirable to provide film cooling directly from the radially extending chamber 134 to certain areas of the airfoil body 122, such as the leading edge 128 and / or the trailing edge 130. The support structure 180 may include any now known or later developed element capable of positioning the first end 164 of the hollow body 162 relative to the shank inner surface 154. The support structure 180 may include, but is not limited to, a lattice structure, straight or curved rods, and the like. 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 integrated cooling features. In one example, the impingement cooling structure 160 may be optionally formed with varying wall thicknesses. The varying wall thicknesses may be advantageous, for example, to accommodate different CTEs between the impingement cooling structure 160 and the hotter wing-shaped body 122, shank 148, and / or platform 150. Figure 5 As shown, the wing-shaped body 122, the stem 148, and / or the 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, such as Figure 7 As shown, the hollow body 162 may include at least one first portion 182 having a first wall thickness W1 between its interior surface (168) and its exterior surface (170), and at least one second portion 184 having a second wall thickness W2 between its interior surface (168) and its exterior surface (170). In the example shown, the first wall thickness W1 is greater than the second wall thickness W2. Any number of thicker portions 182 and / or thinner portions 184 may be provided in the impingement cooling structure 160. The thickness of the portions 182, 184 may be any size required to meet the structural and / or thermal requirements of the location.

[0066] In another exemplary optional configuration, additional supports may be desired and / or required to support the integral impingement cooling structure 160 relative to the inner surfaces 132, 154. For example, at the thinner wall portion 184 ( Figure 7 ) may be desired and / or required to have additional supports. Figure 7 As shown, the turbine rotor blade 120 may further include a support member 186 located on the outer surface 170 of the hollow body 162 in at least one portion 184 having a thinner wall thickness W2. The support member 186 may be integrally formed with the hollow body 162 (and the remainder of the turbine rotor blade 120) to extend at the first end 164 ( Figure 5) and the second end 166 to space the exterior surface 170 of the hollow body 162 from, for example, the airfoil-shaped inner surface 132. 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 upon the interior surface 132, 154. In certain embodiments, regardless of the wall thickness, the turbine rotor blade 120 may include a support 189 on the exterior surface 170 of the hollow body 162. The support 189 may be integrally formed with the hollow body 162 (and the remainder of the turbine rotor blade 120) to form a support between the first end 164 ( Figure 5 ) and the second end 166 to space the exterior surface 170 of the hollow body 162 from, for example, the wing-shaped inner surface 132. The supports 186, 189 may take any form that allows for reducing stress between the hotter exterior walls 124, 126 of the wing-shaped body 122 and the cooler impingement cooling structure 160, provides for any necessary thermal expansion, provides structural support, and / or spaces the hollow body 162 from the inner surfaces 132, 154 in a desired manner. The supports 186, 189 may have any desired size and / or shape, such as, but not limited to, tubes, rods, and the like.

[0067] Figure 8 1 shows an enlarged cross-sectional view of another alternative embodiment that includes a reinforcement member 190 surrounding at least one of the cooling channels 172. The reinforcement member 190 may include any structural reinforcement member, such as a thicker wall, etc. Figure 4 As shown, certain embodiments may also include reinforcing ribs 192 integrally formed to the interior surface 168 of the hollow body 162. Any number of reinforcing ribs 192 may be provided, and each reinforcing rib may extend any desired radial length of the hollow body 162. The supports 186, 189, the reinforcing member 190, and / or the reinforcing ribs 192 may be integrally formed with the remainder of the turbine rotor blade 120 via additive manufacturing.

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

[0069] Figure 9Also shown is a turbine rotor blade 120 including one or more aft impingement target features 196 located on the inner surface 132. The aft impingement target features 196 may include any now known or later developed structure on the inner surface 132 to facilitate cooling. In the illustrated example, the impingement target features 192 include bumps, but they may include any structure. In one embodiment, the hollow body 162 may include localized protrusions 198 to match the profile of the impingement target features 196, thereby maintaining the spacing Z (i.e., Z1 as shown). While two pairs of target features 196 and protrusions 198 are shown, any number may be used. In one embodiment, the aft impingement target features 196 may also optionally include additional integrated cooling features, such as, but not limited to, film-cooling holes 199. The film-cooling holes 199 direct the coolant flow 136 to impinge upon the aft impingement target features 196 (i.e., their inner surface 132) to form a cooling film 201 over the sidewalls 124, 126. Any number of film-cooling holes 199 may be implemented within each aft impingement cooling feature 196 .

[0070] Figure 10 Shown along Figure 4 The first radial cross-sectional view of the middle sight line 10-10, and Figure 11 Shown along Figure 4 A second radial cross-sectional view of the middle sight line 11-11, the latter being located at the same Figure 10 In slightly different planes and in opposite directions. Figure 10 As shown, in certain embodiments, the hollow body 162 has a chord-wise width WC1 that is smaller near the tip 140 than at the shank 148. Most conventional impact inserts have an opposite chord-width arrangement to allow them to be inserted through the open end of the wing-shaped body. Additionally, the hollow body 162 may have a chord-wise width WC1 that is smaller near the tip 140 than at the shank 148. Figure 10 and Figure 11 The impingement cooling structure 160 may have an alternating wider and narrower chord-wise width WC1 in the vertical direction of the page (in the vertical direction of the page). Therefore, the axial rearward end 194 of the hollow body 162 may vary in chord-wise position along the radial span of the hollow body 162. In this way, the impingement cooling structure 160 may have a curvature over its radial span to align with the wing-shaped inner surface 132 and / or the shank inner surface 154 ( Figure 5 ) are evenly spaced apart in shape (chord-wise width WC1), regardless of the shape of the latter. Conventional impact inserts do not provide such a feature.

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

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

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

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

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

[0076] C. Platform with nonlinear cooling channels

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

[0078] The turbine rotor blade 120 also includes a cooling passage 240 extending from the cooling circuit 234 through the surface 242 of the ramp 230, i.e., the cooling ramp 230 and other structures. The cooling passage 240 is in the platform 150 and is in fluid communication with the cooling circuit 234. Compared to conventional linear cooling passages, the cooling passage 240 extends from the cooling circuit 234 in a nonlinear configuration, exiting through at least one ramp 230 of the platform 150, thereby providing improved cooling compared to linear cooling passages. For example, in Figure 12 , the cooling channels 240 have a (gentle) curved shape. Any number of cooling channels 240 may be employed to provide the desired cooling. Furthermore, they may have any desired uniform or non-uniform cross-sectional shape and may be spaced uniformly or non-uniformly to provide the desired cooling. Non-linear configurations are made possible, for example, by additive manufacturing. As described above, the wing-shaped body 122 and the platform 150 (including the components defining the cooling channels 240) may comprise multiple, integral layers of material.

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

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

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

[0082] The cooling circuit 234 and the cooling passage 240 may be provided separately in the pressure side 232 of the platform (only Figure 13 ), which is separately provided in the suction side 231 of the platform 150 (only Figure 12 ), or provided on both sides 231, 232 of the platform 150 ( Figures 12 to 13 If the cooling circuits and cooling channels are provided on only one side of the platform 150, any other conventional structures may be provided on the other side of the platform. Figure 12 and Figure 13 As shown collectively, the cooling circuit 234 may include a first portion 234SS in the suction side 231 of the platform 150 and a second portion 234PS on the pressure side 232 of the platform 150. The portion 234SS and the portion 234PS may be separate or fluidly coupled. In this regard, the slant surface 230 includes a suction side slant surface 230SS and a pressure side slant surface 230PS. Here, the cooling passages 240 in the platform 150 may include at least one first cooling passage 240 in fluid communication with the first portion 234SS of the cooling circuit and exiting the suction side slant surface 230SS, and at least one second cooling passage 240 in fluid communication with the second portion 234PS of the cooling circuit 234 and exiting the pressure side slant surface 230PS.

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

[0084] D. Angel wings with coolant delivery channels

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

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

[0087] Go to Figures 19 to 22 According to an embodiment of the present disclosure, the turbine rotor blade 120 may further include a coolant delivery channel 290 defined by at least one angel wing 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 and a second wheel space portion 294, the first wheel space portion being defined at the shank 148A ( Figure 21 ) and the first adjacent shank 148B of the first adjacent turbine rotor blade 120B ( Figure 21), the second wheel space portion is defined between the shank 148A ( Figure 21 ) and the second adjacent shank 148C of the second adjacent turbine rotor blade 120C ( Figure 21 ). As observed Figure 2 、 Figure 21 and Figure 22 As best shown, each wheel space portion 292, 294 is part of a wheel space 300. The wheel space 300 is defined as a wheel space 300 circumferentially bounded by the shanks 148A-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 wheelspace portions 292, 294 are portions of the wheelspace 300 located axially beside the shank 148 of a particular blade.

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

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

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

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

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

[0093] Turbine rotor blade root 144 may also include a lattice support structure 340 disposed within hollow interior 330 of blade mount 146. Grid support structure 340 may take a variety of hollow support structure forms. In one example, lattice support structure 340 may include a plurality of radially extending V-shaped segments 342. V-shaped segments 342 may be integrally formed with inner wall surfaces 344 of blade mount 146. Root 144 (including shank 148 and blade mount 146, including lattice support structure 340) may be fabricated using additive manufacturing. Thus, shank 148 and blade mount 146 may include multiple, integral layers of material.

[0094] As this article is relative to Figures 12 to 18As described, the root portion 144 according to this embodiment (i.e., having the lattice support structure 340) may also include a platform 150. As described above, the platform 150 is positioned radially outward from the shank 148 and extends laterally outward relative to the shank, terminating in at least one inclined surface 230. The platform 150 may include a cooling circuit 234 defined within the platform and in fluid communication with a source of coolant flow (e.g., the radially extending chamber 134). The cooling channel 240 ( Figures 12 to 18 ) may be defined in the platform 150 and in fluid communication with the cooling circuit 234. As described above, the cooling channel 240 extends from the cooling circuit 234 in a non-linear configuration, exiting through the inclined surface 230 of the platform. The inclined surface 230 may include an extension member 244 through which the cooling channel 240 extends. The cooling channel 240 may have: a spiral 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 ); multiple branches ( Figure 17 ); or curved shapes (e.g. Figure 12 、 Figure 14 、 Figure 18 ).

[0095] As this article is relative to Figures 19 to 24 As described above, the root portion 144 according to this embodiment (i.e., the root portion having the grid support structure 340) may further include an angel wing 280 extending laterally from at least one side of the handle 148. As described above, the coolant delivery channel 290 may be defined by the angel wing 280. Figure 21 As shown, the coolant transfer passage 290 is fluidly coupled to a first wheel space 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 wheel space 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 passage 290 includes a first open end 310 in fluid communication with the first wheel space portion 292 and a second open end 312 in fluid communication with the second wheel space portion 294. Figure 22 As shown, the first open end 310 and the second open end 312 may face in a circumferential direction relative to the shank 148. The coolant delivery passage 290 may be fluidly isolated from the radially extending chamber 134 in the shank 148.

[0096] As described herein, the root portion 144 according to this embodiment (i.e., the root portion having the lattice support structure 340) may also include both the platform 150 and the angel wings 280. Additive manufacturing allows the root portion 144 to be formed with the shank 148, hollow blade mount 146, lattice support structure 340, and the platform 150 and / or angel wings 280, thereby creating multiple, integral layers of material for any features provided.

[0097] The root portion 144 including the integral grid support structure 340 located within the hollow interior 330 of the blade mount 146 provides for a lighter turbine rotor blade 120 and additional cooling of 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 can be used alone or in combination.

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

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

Claims

1. A turbine rotor blade (120), comprising: an airfoil body (122) comprising a concave pressure side outer wall (124) and a convex suction side outer wall (126) connected along leading and trailing edges (128, 130), the outer walls (124, 126) having an airfoil inner surface (132) defining a radially extending chamber (134) for receiving a flow of coolant (236); a tip (140) at a radially outer end (142) of the wing-shaped body (122); a shank (148, 148A) at a radially inner end (152) of the wing-shaped body (122), the radially extending chamber (134) extending at least partially into the shank (148, 148A) to define a shank inner surface (154); and An impingement cooling structure (160) is provided within the radially extending chamber (134), the impingement cooling structure (160) comprising: a hollow body (162) comprising a first end (164), a second end (166), an interior surface (168), and an exterior surface (170), a plurality of cooling passages (172) extending through the hollow body (162) and in fluid communication with the radially extending chamber (134) to allow the coolant (236) to flow from the interior surface (168) of the hollow body (162) to impinge on at least the airfoil inner surface (132), wherein the first end portion (164) of the hollow body (162) is integrally formed with the handle inner surface (154), wherein a first chord-wise width of the hollow body at the top end is smaller than a second chord-wise width of the hollow body at the handle, wherein the exterior surface (170) of the hollow body (162) is evenly spaced from the wing-shaped interior surface (132) between the first end (164) and the second end (166) of the hollow body (162), wherein the hollow body (162) includes at least one first portion (234SS) and at least one second portion (234PS), the at least one first portion having a first wall thickness between the interior surface (168) and the exterior surface (170) of the hollow body, and the at least one second portion having a second wall thickness between the interior surface (168) and the exterior surface (170) of the hollow body, the first wall thickness being greater than the second wall thickness, wherein the turbine rotor blade (120) further comprises a support member (186, 189) on the outer surface (170) of the hollow body (162) in the at least one second portion (234PS) having the second wall thickness, the support member (186, 189) being integrally formed with the hollow body (162) to space (204) the outer surface (170) of the hollow body (162) from the airfoil-shaped inner surface (132) between the first end (164) and the second end (166) of the hollow body (162), and wherein the support member (186, 189) includes a passage therethrough, the passage being in fluid communication with one of the plurality of cooling passages (172, 240).

2. The turbine rotor blade (120) of claim 1, wherein the second end (166) of the hollow body (162) is integrally formed to an inner surface (176) of the tip (140).

3. The turbine rotor blade (120) of claim 1, wherein the first end (164) of the hollow body (162) extends substantially in a radial direction from a location (174) where the first end (164) of the hollow body (162) meets the shank inner surface (154).

4. The turbine rotor blade (120) of claim 3, wherein at least a portion of the shank inner surface (154) extends from the junction (174) of the first end (164) of the hollow body (162) and the shank (148, 148A) at an angle relative to the radial direction.

5. The turbine rotor blade (120) of claim 3, further comprising a platform (150) located between the shank (148, 148A) and the radially inner end (152) of the airfoil-shaped body (122), the platform (150) extending laterally outward relative to the shank (148, 148A), and further comprising a support structure (180) between the first end (164) of the hollow body (162) and the shank inner surface (154), the support structure (180) being located radially outward of the confluence location (174) and radially inward of the platform (150).

6. The turbine rotor blade (120) of claim 1, further comprising a platform (150) located between the shank (148, 148A) and the radially inner end (152) of the airfoil-shaped body (122), the platform (150) extending laterally outward relative to the shank (148, 148A), and wherein the first end (164) of the hollow body (162) is integrally formed to the shank inner surface (154) radially inwardly along the platform (150).

7. The turbine rotor blade (120) of claim 1, further comprising a support member (186, 189) on the outer surface (170) of the hollow body (162), the support member (186, 189) being integrally formed with the hollow body (162) to space (204) the outer surface (170) of the hollow body (162) from the airfoil-shaped inner surface (132) between the first end (164) and the second end (166) of the hollow body (162).

8. The turbine rotor blade (120) of claim 1, further comprising a reinforcement member (190) surrounding at least one cooling channel of the plurality of cooling channels (172, 240).

9. The turbine rotor blade (120) of claim 1, wherein the airfoil body (122), the tip (140), the shank (148, 148A), and the impingement cooling structure (160) comprise a plurality of integral layers of material.

10. The turbine rotor blade (120) of claim 1, further comprising strengthening ribs (192) integrally formed to the interior surface (168) of the hollow body (162).

11. The turbine rotor blade (120) of claim 1, wherein the plurality of cooling passages (172, 240) passing through the hollow body (162) and in fluid communication with the radially extending chamber (134) extend around the entire circumference of the hollow body (162).

12. The turbine rotor blade (120) of claim 1, wherein the axially aft end (194) of the hollow body (162) varies in chordwise position along the radial span of the hollow body (162), and The airfoil body (122) further includes at least one chord-wise extending serpentine cooling channel (172, 240) extending from the airfoil inner surface (132) rearward of the hollow body (162) toward the trailing edge (130), wherein each chord-wise extending serpentine cooling channel (172, 240) has the same chord-wise width.

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