Turbine rotor blade with wing seal by additive manufacturing and coolant transport channels between adjacent wheel space portions

Abstract

A turbine rotor blade is additively manufactured and includes 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 is at a radially inner end of the airfoil body, and at least one wing extends laterally from at least one side of the shank. A coolant delivery passage is defined through the at least one wing. The coolant delivery passage fluidly couples a first wheel space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade and a second wheel space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade. The coolant delivery passage allows coolant to pass between the wheel space portions of the adjacent turbine rotor blades.

Description

Background Technology

[0001] This disclosure relates generally to turbines, and more specifically to turbine rotor blades having multiple additively manufactured integral features, such as integral impact sleeves, nonlinear cooling channels in a platform, airfoils with coolant delivery channels, and hollow dovetails with integral grid support structures.

[0002] A turbine comprises multiple turbine blades coupled to a rotor. A working fluid, such as steam or combusted fuel, is forced against the blades to rotate the rotor. The turbine blades operate under extremely hot conditions and require cooling. Cooling features can be provided in various ways.

[0003] One mechanism for providing cooling is the impact insert. An impact insert, or sleeve, comprises a hollow body with cooling channels within its walls, allowing coolant to be delivered through the channels to impact or impinge on the surface to be cooled. Impact 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 challenge with impact inserts is positioning them close enough within the tapered or curved cavity of the HGP component to achieve high cooling performance, but not so close that cooling is ineffective. One indicator of the cooling performance of an impact insert is the Z / D parameter, which is the ratio of the distance Z between the insert and the inner surface of the HGP component to the diameter D of the cooling channel (orifice) in the impact insert. The Z / D parameter value of the insert is typically designed to be within the desired range that results in better cooling performance.

[0004] Where the necessary spacing cannot be achieved, impact cooling is typically not provided by the inserts. For example, if the cavity curvature in an HGP component is so significant that the impact inserts cannot be manufactured thin enough or bent to the appropriate necessary spacing, impact cooling cannot be provided. One approach to this problem is to provide impact inserts in multiple flexible longitudinal sections to facilitate their insertion into the HGP component. However, the need to sequentially position and connect multiple insert sections together or attach them to the HGP component increases manufacturing complexity, time, and cost. The flexible impact insert sections also do not provide adjacent elements around their perimeter (i.e., transverse (section)), which can impair cooling performance at their discontinuities.

[0005] Shock cooling is applied in a limited way to rotating turbine rotor blades in turbines, for example, to their leading edge. However, shock cooling has not been applied more widely to the entire inner surface of turbine rotor blades because the centrifugal force experienced by the rotating blades during their rotation forces the coolant to the radially outer ends of the blades, making shock cooling less effective.

[0006] Another cooling feature includes cooling channels that pass through a portion of the turbine rotor blade to be cooled. For example, a turbine rotor blade includes a platform that extends laterally to cooperate with the platforms of adjacent turbine rotor blades to form part of the working fluid path through the turbine. Due to the high temperature of the working fluid, the platform typically includes a cooling circuit that feeds into multiple cooling channels exiting through an inclined plane of the platform. Some platforms include damping pin seats in the inclined plane that receive axially extending pins that engage with adjacent damping pin seats in adjacent platforms to seal the working fluid path. Cooling channels are typically drilled into the inclined plane to fluidly connect the channels to the cooling circuit. Therefore, the cooling channels have a linear configuration that may not adequately cool all platforms. For example, a cooling channel may pass through the extension forming the damping pin seat but not adequately cool other portions of the inclined plane.

[0007] Cooling features can also be used in conjunction with airfoils. In this regard, another cooling feature includes cooling channels that deliver coolant into or radially surround the airfoil. Mounting brackets for turbine rotor blades may also include cooling features. Summary of the Invention

[0008] A first aspect of this disclosure provides 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, the outer wall having an airfoil inner surface defining a radially extending chamber for receiving a coolant flow; an end 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 extending at least partially into the shank to define an inner surface of the shank; and an impingement cooling structure. The structure is located in a radially extending chamber. The impact cooling structure includes: 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 from the inner surface of the hollow body to impact at least the inner surface of the airfoil, wherein the first end of the hollow body is integrally formed with the inner surface of the shank, and wherein the outer surface of the hollow body is uniformly spaced from the inner surface of the airfoil between the first end and the second end of the hollow body.

[0009] A second aspect of this disclosure provides 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, the outer wall having an airfoil inner surface defining a radially extending chamber for receiving a flow of coolant; and an integral impact cooling structure located within the radially extending chamber, the integral impact 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 a flow of coolant from the inner surface of the hollow body to impact at least the inner surface of the airfoil, wherein the outer surface of the hollow body is uniformly spaced from the inner surface of the airfoil between the first end and the second end of the hollow body.

[0010] A third aspect of this 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 wall having an airfoil inner surface defining a radially extending chamber for receiving a flow of coolant; and an impact cooling structure located within the radially extending chamber, the integral impact 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 a flow of coolant from the inner surface of the hollow body to impact at least the inner surface of the airfoil, wherein the outer surface of the hollow body is uniformly spaced from the inner surface of the airfoil between the first end and the second end of the hollow body.

[0011] A fourth aspect of this disclosure provides 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, the outer wall 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 ramp; a cooling circuit defined within the platform and in fluid communication with a source of the coolant flow; and at least one cooling passage defined within the platform and in fluid communication with the cooling circuit, the at least one cooling passage extending in a non-linear configuration from the cooling circuit through at least one ramp of the platform to an outlet.

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

[0013] The 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 wing extending laterally from at least one side of the shank; and a coolant delivery passage defined by the at least one wing, the coolant delivery passage fluidly connecting a first wheel space portion defined between the shank and a first adjacent shank portion of a first adjacent turbine rotor blade and a second wheel space portion defined between the shank and a second adjacent shank portion of a second adjacent turbine rotor blade.

[0014] The seventh aspect of this 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 located at a radially inner end of the airfoil body; at least one wing extending laterally from at least one side of the shank; and a coolant delivery passage defined by the at least one wing, the coolant delivery passage fluidly connecting a first wheel space portion defined between the shank and a first adjacent shank portion of a first adjacent turbine rotor blade and a second wheel space portion defined between the shank and a second adjacent shank portion of a second adjacent turbine rotor blade.

[0015] The 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 and third turbine rotor blades, each 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 wing extending laterally from at least one side of the shank, wherein the shanks of the first and second turbine rotor blades define a first wheel space portion therebetween, and the shanks of the first and third turbine rotor blades define a second wheel space portion therebetween; and a coolant delivery passage defined by at least one wing of the first turbine rotor blade, the coolant delivery passage fluidly connecting the first and second wheel space portions.

[0016] The ninth aspect relates to a turbine rotor blade root, the turbine rotor blade root comprising: a shank having a radially extending chamber defined therein; a blade mount located at a radially inner end of the shank having a hollow interior defined therein 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 comprising: a shank having a radially extending chamber defined therein; a blade mount located at a radially inner end of the shank having a hollow interior defined therein in fluid communication with the radially extending chamber; a grid support structure disposed within the hollow interior of the blade mount; at least one wing extending laterally from at least one side of the shank; and a coolant delivery passage defined by the at least one wing, the coolant delivery passage fluidly connecting a first wheel space portion defined between the shank and a first adjacent shank portion of a first adjacent turbine rotor blade root, and a second wheel space portion defined between the shank and a second adjacent shank portion of a second adjacent turbine rotor blade root.

[0018] The eleventh aspect relates to a turbine rotor blade root, the turbine rotor blade root comprising: a shank having a radially extending chamber defined therein; a blade 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 wing extending laterally from at least one side of the shank; and a coolant delivery passage defined by the at least one wing, the coolant delivery passage fluidly connecting a first wheel space portion defined between the shank and a first adjacent shank portion of a first adjacent turbine rotor blade root and a second wheel space portion defined between the shank and a second adjacent shank portion of a second adjacent turbine rotor blade root.

[0019] The exemplary aspects of this disclosure are designed to address the problems described herein and / or other problems not discussed herein. Attached Figure Description

[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 accompanying drawings depicting various embodiments thereof, wherein:

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

[0022] Figure 2 It shows that it can be used with Figure 1 A cross-sectional view of an exemplary gas turbine assembly used in the GT system.

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

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

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

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

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

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

[0029] Figure 9 A plan sectional view of an impact 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 impact sleeve according to an embodiment of the present disclosure is shown.

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

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

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

[0034] Figure 14 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 is shown.

[0035] Figure 15 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 is shown.

[0036] Figure 16 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 is shown.

[0037] Figure 17 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 is shown.

[0038] Figure 18 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 is shown.

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

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

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

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

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

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

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

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

[0047] It should be noted that the accompanying drawings of this disclosure are not necessarily drawn to scale. The drawings are intended to depict only typical aspects of this disclosure and should therefore not be considered as limiting the scope of this disclosure. In the drawings, similar numbers denote similar elements between figures. Detailed Implementation

[0048] As an initial matter, in order to clearly describe the present disclosure, it will be necessary to choose certain terms when referring to and describing relevant machine parts, such as those within a turbine. In doing so, common industry terms will be used and adopted in a manner consistent with their accepted meanings, where possible. Unless otherwise stated, such terms should be given a broad interpretation consistent with the context of this application and the scope of the appended claims. Those skilled in the art will understand that several different or overlapping terms may generally be used to refer to a particular part. An object that can be described herein as a single part may include multiple parts and is referred to in another context as being composed of multiple parts. Alternatively, an object that can be described herein as comprising multiple parts may elsewhere be referred to as a single part.

[0049] In addition, several descriptive terms may be used periodically throughout this document, and it should prove helpful to define these terms at the beginning of this section. Unless otherwise stated, these terms and their definitions are as follows. As used herein, “downstream” and “upstream” are terms indicating the direction of fluid flow, such as the working fluid through a turbine, or, for example, the airflow through a combustor or the coolant through a component 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. Without any other 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 accompanying drawings, front and rear are generally indicated by the X direction. It is often necessary to describe parts that are in different radial positions relative to a central axis. The term “radial” refers to movement or position perpendicular to an 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, this document will state that the first component is “radially inward” or “inner” of the second component. On the other hand, if the first component resides further away from the axis than the second component, this document may state that the first component is "radially outward" or "outer" of the second component. In the accompanying drawings, the radial direction is typically represented by the Z-direction. The term "axial" refers to movement or position parallel to an axis, such as the axis of a turbine rotor. Finally, the term "circumferential" refers to movement or position about an axis. Although not shown as curved in the legends of the accompanying drawings, the circumferential direction is typically represented by the Y-direction in the drawings. It should be understood that such terms can be applied relative to the rotor axis of a turbine.

[0050] In addition, several descriptive terms may be used regularly in this document, 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 location or importance of a single component.

[0051] The terminology used herein is for the purpose of describing particular embodiments only and is not intended to limit this 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 “comprising” and / or “including” specify the presence of the stated features, integers, steps, operations, elements, and / or components, but do not exclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and / or groups thereof. “Optional” or “optionally” means that an event or condition subsequently described 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 “on,” “joined to,” “connected to,” or “linked to” another element or layer, it may be directly on, joined to, connected to, or linked to the other element or layer, or an intermediary element or layer may be present. In contrast, when an element is referred to as “directly on,” “directly joined to,” “directly connected to,” or “directly linked to” another element or layer, an intermediary element or layer may not be present. Other terms used to describe relationships between elements should be interpreted in a similar manner (e.g., “between” vs. “directly between,” “adjacent” vs. “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, this disclosure provides a turbine rotor blade or turbine rotor blade root comprising a plurality of integral features made possible by additive manufacturing of the blade and / or root. Additive manufacturing allows for the formation of structures that provide cooling where it was previously not permitted, improve cooling compared to conventional systems, provide additional structural strength, and / or reduce blade weight.

[0054] A. Introduction

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

[0056] During 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 communication with fuel source ( Figure 1 (Not shown) is in fluid communication with a burner 104 and directs fuel and air to the combustion zone 105. The burner 104 is ignited and the fuel is burned. The burner 104 is in fluid communication with a turbine 108 so that the thermal energy of the gas flow is converted into mechanical rotational energy. The turbine 108 is rotatably coupled to and drives a rotor 110. A compressor 102 is also rotatably coupled to the rotor 110. In an exemplary embodiment, multiple burners 104 and fuel nozzle assemblies 106 are present.

[0057] Figure 2 It shows that it can be used with Figure 1 A cross-sectional view of an exemplary turbine 108 with three stages, used in the GT system 100. Each stage includes a fixed blade or nozzle assembly 112 and turbine rotor blades 120. The fixed nozzles 112 are held in the turbine 108 by radially outer platforms 114 and radially inner platforms 116. The fixed 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 fixed nozzles 112. Combustion gas is guided by the fixed nozzles 112 against the turbine rotor blades 120 to cause the rotor 110 ( Figure 1 Rotate.

[0058] Figure 3 A perspective view of an exemplary turbine rotor blade 120 of a GT system 100 is shown, in which the overall feature portion may be adopted according to various embodiments of the present disclosure. Figure 4 An axial cross-sectional view of a turbine rotor blade 120 according to various embodiments of the present disclosure is shown, the turbine rotor blade including an integral feature in the form of an integral impact cooling structure 160. The turbine rotor blade 120 includes an airfoil body 122, the airfoil body 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. See again Figure 3 The turbine rotor blade 120 may also include a tip 140 located at the radially outer end 142 of the airfoil body 122. The turbine rotor blade 120 may also include a turbine rotor blade root 144 (hereinafter referred to as "root 144"), through which the turbine rotor blade 120 passes, for example via the rotor wheel 147 (…). Figure 21 ) Attached to rotor 110 ( Figure 1 For the purposes of this disclosure, root 144 may include any portion of turbine rotor blade 120, including platform 150 and radially inward of platform. Root 144 may include blade mount 146 configured for mounting on rotor wheel 147. Figure 21 The blade mounting bracket 146 may have corresponding slots around the rotor disk 147. Figure 21 Any now-known or later-developed external configuration, such as, but not limited to, dovetail or fir arrangement. The turbine rotor blade 120, i.e., its root 144, may also include a shank 148 extending between the blade mount 146 and the platform 150. The platform 150 is located at the junction of the airfoil body 122 and the shank 148, and defines the space through the turbine 108 ( Figures 1 to 2 The shank 148 is a portion of the inner boundary of the flow path of the airfoil body 122. The shank 148 is therefore located at the radially inner end 152 of the airfoil body 122, and the blade mount 146 is located radially inner to the shank 148. The platform 150 extends laterally outward relative to the shank 148. As will be further described herein, a radially extending chamber 134 may extend at least partially into the shank 148 to define the 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 end 140 in the radial (Z) direction. It should be understood that the airfoil body 122 is a moving part of the turbine rotor blade 120, which intercepts the flow of working fluid and causes the rotor to rotate.

[0059] B. Integral impact cooling structure

[0060] In some embodiments, the turbine rotor blade 120 may include, in particular, an airfoil body 122 and an integral feature in the form of an integral impact cooling structure 160 therein. The impact cooling structure is not an insert but is integrally formed with the rest of the blade by, for example, 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 terminal end 140 located at a radially outer end 142 of the airfoil body 122 and a shank 148 located at a radially inner end 152 of the airfoil body 122. The radially extending chamber 134 may extend at least partially into the shank 148 to define the shank inner surface 154. The integral impact cooling structure 160 is located within a radially extending chamber 134 and may include a hollow body 162 comprising 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 coolant flow from the inner surface 168 of the hollow body 162 to impact at least the inner surface of the airfoil body 122. Compared to conventional impact inserts, the first end 164 of the hollow body 162 is integrally formed with the inner surface 154 of the shank, i.e., formed via additive manufacturing. Therefore, the outer surface 170 of the hollow body 162 may be uniformly spaced from the inner surface 132 of the airfoil between the first end 164 and the second end 166 of the hollow body 162, regardless of the curvature of the inner surface 132. In another embodiment, a non-uniform but customized spacing may be employed to provide different impact cooling, heat absorption, and / or reuse. For example, closer spacing can be used where increased shock cooling is required, and wider spacing can be used where less shock cooling is required. Additionally, the hollow body 162 can have cooling channels around its entire periphery and radial span to provide shock cooling across the entire blade, not just at its leading edge. Therefore, the integral shock cooling structure allows for maximum shock range with the limited sacrifices typically associated with shock inserts, and can have multiple variable cooling features. For example, the turbine rotor blade can have: a variable chord width of the shock cooling structure or its rear pin assembly; a shock cooling structure wall thickness tailored to different cooling loads; and different supports to address the different coefficients of thermal expansion (CTE) between the airfoil body and the shock cooling structure.

[0061] like Figure 4 as well as Figure 5As shown in the radial circumferential sectional view, the turbine rotor blade 120 may include an impact cooling structure 160 located within a radially extending chamber 134. Impact cooling is typically provided by one or more impact inserts inserted into the radially extending chamber 134 and attached to the airfoil body 122, for example, by fasteners or welding. The impact inserts are typically linear but may include some curvature. The radially extending chamber 134 has, for example,... Figure 5 In the case of the curved airfoil inner surface 132 shown, it is impossible to have impact inserts that are uniformly spaced from the inner surface along the entire radial span of the blade. To address this challenge, the impact cooling structure 160 according to an embodiment of the present disclosure is integrally formed with the remainder of the turbine rotor blade 120 via additive manufacturing.

[0062] like Figure 5 As shown, the impact cooling structure 160 includes a hollow body 162, which includes a first end 164, a second end 166, an inner surface 168, and an outer surface 170. The impact cooling structure 160 also includes a plurality of cooling channels 172 that pass through the hollow body 162 and are in fluid communication with a radially extending chamber 134 to allow a coolant flow 136 to pass through the inner surface 168 of the hollow body to impact at least the inner surface 132 of the airfoil, for example, impacting the inner surfaces, particularly the airfoil body 122, the end 140, the shank 148, and / or the platform 150. Compared to conventional turbine rotor blades, and as... 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 impact cooling structure 160 can (simultaneously with, for example, the airfoil inner surface 132) be formed to have the same curvature, bending, twisting, and any other shape or size to match the shape or size of its adjacent inner surface. Notably, the impact cooling structure 160 can be uniformly spaced from the airfoil inner surface 132 along its entire radial span, thereby ensuring the desired Z / D parameter on all turbine rotor blades 120. The Z / D parameter is the ratio of the distance Z between the outer surface 170 and the inner surface (e.g., airfoil inner surface 132, shank inner surface 154, etc.) of the turbine rotor blade 120 to the diameter D of the cooling channel 172 (hole) in the impact cooling structure 160. In one example, Z / D is in the range of about 1 to about 10. In another example, Z / D can range from about 2 to about 6. The spacing Z can be smaller than the conventionally available distance for castings, for example, less than about 1.27 mm (0.05 inches). Cooling channels 172 with a smaller diameter D than conventional castings can also be used, for example, to address debris blockage. Advantageously, cooling channels 172 can extend around the entire periphery of the hollow body 162, allowing coolant flow 136 to exit the hollow body 162 in all directions to provide impingement cooling to all airfoil inner surfaces 132 of the airfoil body 122. Alternatively, cooling channels 172 can be omitted in areas where impingement cooling of the inner surfaces 132, 154 is not desired or required. Cooling channels 172 can extend along any desired radial range of the hollow body 162.

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

[0064] Figure 6 The meeting point 174 of the impact cooling structure 160 and the shank 148 of the turbine rotor blade 120 according to various embodiments of the present disclosure is shown. Figure 5 An enlarged sectional view of (e.g.) Figure 5 and Figure 6 As shown, the first end 164 of the hollow body 162 may extend substantially in the radial direction (arrow Z) relative to the meeting point 174 of the first end 164 of the hollow body 162 and the inner surface 154 of the handle. As used herein, "substantially in the radial direction" indicates that the first end 164 is located away from the rotor 110 by a certain tolerance (e.g., + / - 5°). Figure 1 ) Extends radially. In contrast, at least a portion of the inner surface 154 of the handle extends at an angle α relative to the radial direction Z from the first end 164 of the hollow body 162 to the meeting position 174 of the handle 148. Figure 6 In another embodiment, the inner surface 154 of the handle is substantially aligned in the radial direction, and the first end 164 of the hollow body gradually bends or transitions toward the meeting point 174 to maintain an angle α, for example, <30°. In another embodiment, both the inner surface 154 of the handle and the first end 164 of the hollow body 162 gradually bend or transition toward the meeting point 174. The angle α can be any desired angle and is within the range of additive manufacturing, for example, <45° with respect to the vertical direction. To maintain structural integrity, the angle α is desired to be as small as possible, for example, <10°, <20°, or <30°. Figure 6 As shown, in some embodiments, the support structure 180 may be positioned between the first end 164 of the hollow body 162 and the inner surface 154 of the handle, for example, radially outward of the meeting point 174 and radially inward of the platform 150. In other embodiments, the support structure 180 may be positioned anywhere between the outer surface 170 of the hollow body 162 and the inner surface 132 of the airfoil, the inner surface 154 of the handle, etc. In other embodiments, at least a portion of the support structure 180 includes a hollow support element (e.g., a grille) to allow cooling flow to 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 now-known or later-developed element capable of positioning the first end 164 of the hollow body 162 relative to the inner surface 154 of the shank. The support structure 180 may include, but is not limited to, a grid structure, straight or curved rods, etc. The support structure 180 may also be integrally formed via additive manufacturing.

[0065] The impact cooling structure 160 may also include a variety of optional integral cooling features. In one example, the impact cooling structure 160 may be optionally topped with varying wall thicknesses. Different wall thicknesses can facilitate different CTEs between, for example, the impact cooling structure 160 and the hotter airfoil body 122, the shank 148, and / or the platform 150. Figure 5 As shown, the airfoil body 122, the handle 148, and / or the platform 150 may have a variety of wall thicknesses, and may have varying thicknesses within their range. Figure 7 An enlarged partial cross-sectional view of a portion of the impact cooling structure 160 adjacent to the airfoil body 122, platform 150, or handle 148 is shown. As described above, in some 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 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 impact cooling structure 160. The thickness of portions 182, 184 may be any dimension required to meet the structural requirements and / or thermal requirements of the location.

[0066] In another exemplary alternative configuration, additional supports may be desired and / or required to support the overall impact cooling structure 160 relative to the inner surfaces 132, 154. For example, in the thinner wall portion 184 of the impact cooling structure 160... Figure 7 Additional support components may be expected and / or required at this location. Therefore, as... Figure 7 As 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 relatively thin wall thickness W2. The support 186 may be integrally formed with the hollow body 162 (and the rest of the turbine rotor blade 120) to be positioned at the first end 164 of the hollow body 162. Figure 5 The outer surface 170 of the hollow body 162 is spaced from, for example, the airfoil-shaped inner surface 132 between the first end 164 and the second end 166. Any number of supports 186 may be provided in the thinner wall portion 184. The supports 186 may include a channel 188 therethrough, which is in fluid communication with one of a plurality of cooling channels 172, i.e., to allow a coolant flow 136 to pass through and impinge on the inner surfaces 132, 154. In some embodiments, regardless of 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 be positioned at the first end 164 of the hollow body 162. Figure 5 The outer surface 170 of the hollow body 162 is spaced from, for example, the inner surface 132 of the airfoil between the second end 166 and the second end 166. Supports 186, 189 may take any form that allows for: reducing stress between the hotter outer walls 124, 126 of the airfoil body 122 and the colder impact 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. Supports 186, 189 may have any desired size and / or shape, such as, but not limited to, tubes, rods, etc.

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

[0068] Figure 9 A cross-sectional view of a turbine rotor blade 120 is shown, which includes an integral shock cooling structure 160 and additional optional integral cooling features. In an alternative embodiment, the shock cooling structure 160 may optionally be configured with a varying spacing Z relative to the inner surfaces 132, 154. The spacing Z may be customized to provide desired Z / D parameters and desired cooling at various locations. For example, the turbine rotor blade 120 may have multiple 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 following regions: a region near the leading edge 128 (195A), a region near the pressure-side outer wall 124 of the trailing edge 130 (195B), and a region downstream of the leading edge 128, on the suction side wall 126 (195C). In the high heat load region 195, a first spacing Z1 can be used between the integral impact cooling structure 160 and the inner surfaces 132, 154 of the high heat load region 195, while a larger second spacing Z2 can be used in other locations without such high heat loads. Thus, when necessary (i.e., in the high heat load region 195), the first spacing Z1 can be used to provide more cooling, wherein the spacing between the impact cooling structure 160 and the inner surfaces 132, 154 is increased to a second, larger spacing Z2 for lower heat load regions. Figure 9As shown, a larger second spacing Z2 allows 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 greater heat absorption capacity for downstream areas such as the helical cooling channel 200 and / or pin assembly 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 any value, for example, between 1.01 times and 3.00 times the first spacing Z1. The Z / D parameter can be customized for each area of ​​interest. As mentioned above, in one example, Z / D is in the range of about 1 to about 10. In another example, Z / D can be in the range of about 2 to about 6. The diameter D of the cooling channel 172 can also be configured to customize the Z / D parameter for different areas.

[0069] Figure 9 Also shown is a turbine rotor blade 120 including one or more post-impact target features 196 located on its inner surface 132. The post-impact target features 196 may include any now-known or later-developed structures on the inner surface 132 to facilitate cooling. In the example shown, the post-impact target features 192 include bumps, but it may also include any structure. In one case, the hollow body 162 may include localized bulges 198 to match the contour of the post-impact 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 may be used. In one embodiment, the post-impact target features 196 may also optionally include additional integral cooling features, such as, but not limited to, film cooling holes 199. After impact with the post-impact target features 196 (i.e., its inner surface 132), the film cooling holes 199 guide a coolant flow 136 to form a cooling film 201 over the sidewalls 124, 126. Any number of film cooling holes 199 can be applied within each post-impact cooling feature 196.

[0070] Figure 10 It shows along Figure 4 The first radial section view of line 10-10 in the middle, and Figure 11 It shows along Figure 4 The second radial section view of line 11-11, which is in line with Figure 10 In slightly different planes and in opposite directions. For example... Figure 10 As shown, in some embodiments, the hollow body 162 has a smaller chordal width WC1 near the end 140 than the handle 148. Most conventional impact inserts have opposite chordal width arrangements to allow them to be inserted through the open end of the airfoil body. Additionally, the hollow body 162 may have a radial span ( Figures 10 to 11The upper and lower pages of the hollow body 162 have alternating wider and narrower chord widths WC1. Therefore, the axial rear end 194 of the hollow body 162 can vary in chordal position along the radial span of the hollow body 162. Thus, the impact cooling structure 160 can have a radial span that bends to meet the inner surface 132 of the airfoil and / or the inner surface 154 of the shank. Figure 5 The evenly spaced shapes (chord width WC1) are independent of the shape of the other component. Conventional impact inserts do not provide this feature.

[0071] like Figure 4 , Figure 10 and Figure 11 As shown, the airfoil body 122 also includes at least one chordally extending helical cooling channel 200 extending from the inner surface 132 of the airfoil behind the hollow body 162 toward the trailing edge 130. Figure 10 and Figure 11 As best shown, each chordally extending helical cooling channel 200 may have the same chordal width WC2, for example, shorter than the chordal 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., from the helical cooling channels 200). Each of the plurality of trailing-edge cooling channels 202 has the same chordal width WC3, i.e., along the radial span of the turbine rotor blade 120. Figure 10 and Figure 11 As shown, the space 204 between the trailing edge cooling channel 202 and the helical cooling channel 200 has a varying chordal width WC4, i.e., along the radial span of the turbine rotor blade 120. The turbine rotor blade 120 may also include sections located at the front end of the plurality of trailing edge cooling channels 202. Figure 10 On its right side, Figure 11 (on its left side) and the rear end of the chordally extending spiral cooling channel 200 ( Figure 10 On its left side, Figure 11 The pin group 206 is located between the right side of the pin. Therefore, as... Figures 10 to 11 As shown, pin assembly 206 may have a varying chordal width WC4 along its radial span. Figure 10 and Figure 11 Various cooling features are shown, including but not limited to: film cooling via opening 208, chordal impact cooling via impact cooling structure 160, cooling near trailing edge 130 via helical cooling channel 200, and trailing edge 130 pin group cooling via pin group 206.

[0072] Additive manufacturing (AM) encompasses 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 with respect to turbine rotor blade 120, without the use of any kind of tooling, molds, or jigs, and with little or no material waste. Instead of machining parts from solid material blanks (many of which are cut away and discarded), the only material used in additive manufacturing is the material required to shape the part. Additive manufacturing techniques typically involve taking a three-dimensional computer-aided design (CAD) file of the part to be formed (e.g., turbine rotor blade 120), electronically cutting the part into layers (e.g., 18–102 micrometers thick), and creating a file containing a two-dimensional image (including vectors, images, or coordinates) of each layer. This file can then be loaded into a fabrication software system that interprets the file, allowing the part to be constructed using different types of additive manufacturing systems. In the forms of additive manufacturing, such as 3D printing, rapid prototyping (RP), and direct digital manufacturing (DDM), material layers are selectively applied, sintered, shaped, deposited, etc., to form the part. Although other manufacturing processes such as casting may also be used, the turbine rotor blades 120 can be advantageously manufactured by additive manufacturing.

[0073] In metal powder additive manufacturing technologies 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 other possible forms of additive manufacturing, layers of metal powder are sequentially melted together to form the part. More specifically, fine layers of metal powder are sequentially melted after being uniformly distributed onto a bed of metal powder using an applicator. Each applicator comprises 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 uniformly spreads the metal powder above a build platform. The metal powder bed can move in a vertical axis. The process occurs in a processing chamber with a precisely controlled atmosphere. Once each layer is formed, each two-dimensional slice of the part geometry can be fused by selectively melting the metal powder. This melting can be performed using a high-power melting beam (such as a 100-watt ytterbium laser) to completely weld (melt) the metal powder to form a solid metal. The melting beam is moved in the XY direction using a scanning mirror, and its intensity is sufficient to completely weld (melt) the metal powder to form a solid metal. The metal powder bed can be lowered for each subsequent two-dimensional layer, and this process is repeated until the part is completely formed. To form certain larger blades more quickly, some metal additive manufacturing systems employ a pair of high-power lasers working together to form the blades. Here, the method of fabricating turbine rotor blades 120 may include sequentially forming material layers and applying a heat source to sinter the material layers, thereby forming the structure described herein. Thus, additive manufacturing produces an airfoil body 122, a tip 140, a shank 148, and an impact cooling structure 160 comprising multiple integral material layers.

[0074] The turbine rotor blades 120 may be made of a metal capable of withstanding the environment in which they are employed, and this metal may include pure metals or alloys. In one example, the metal may include virtually any non-reactive metal powder, i.e., a non-explosive or non-conductive powder, such as, but not limited to: cobalt-chromium-molybdenum (CoCrMo) alloys, stainless steel, austenitic nickel-chromium based alloys such as nickel-chromium-molybdenum-niobium alloys (NiCrMoNb) (e.g., Inconel 625 or Inconel 718), nickel-chromium-iron-molybdenum alloys (NiCrFeMo) (e.g., available from Haynes International, Inc.). X) or nickel-chromium-cobalt-molybdenum alloys (NiCrCoMo) (e.g., Haynes 282, available from Haynes International, Inc.). In another example, the metal may include virtually any metal, such as, but not limited to: tool steel (e.g., H13), titanium alloys (e.g., Ti6Al4V), stainless steel (e.g., 316L), cobalt-chromium alloys (e.g., CoCrMo), and aluminum alloys (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. Therefore, the outer surface 170 of the hollow body 162 can be uniformly spaced from the inner surface between the first end 164 and the second end 166 of the hollow body 162, regardless of, for example, the curvature of the airfoil inner surface 132 and / or the shank inner surface 154. Additionally, the hollow body 162 may have cooling channels 172 around its entire periphery and radial span to provide impact cooling over the entire blade, not just at its leading edge. Thus, the integral impact cooling structure 160 allows for maximum impact range without the sacrifices typically associated with impact inserts and can have multiple variable cooling features. For example, the turbine rotor blade 120 may have: a variable chord width of the impact cooling structure 160 (i.e., width WC1) or the pin assembly 206 (i.e., WC4) behind the structure 160; a customized impact cooling structure 160; different wall thicknesses for different cooling and / or structural loads; and different supports 186, 189 to address the different coefficients of thermal expansion (CTE) between the airfoil body 122 and the impact cooling structure 160. Additional cooling features, such as turbulence generators (not shown), may also be provided and customized around each cooling channel 172 to optimize impact cooling. The turbine rotor blade 120 may also include axial ventilation through the trailing edge 130, such as relative to... Figures 10 to 11 As stated above.

[0076] C. Platform with non-linear cooling channels

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

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

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

[0080] The cooling channel 240 can employ any of a variety of nonlinear configurations to provide the desired cooling. The nonlinear configuration (e.g., a curved shape) can extend in any desired direction within the platform 150, such as radial (inward or outward), axial (backward or forward), or circumferential (clockwise or counterclockwise), or a combination of these directions. The cooling channels 240 can all have the same shape to provide the same cooling properties at every location where the cooling channel is set, or the shape of the cooling channel can vary within the platform 150 to provide customized cooling for each location where the cooling channel is set. In addition... 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 (screw) shape, i.e., it may have multiple 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 deflection 252 along the first direction FD. Figure 16 ) and at least one second steering 254 along the opposite second direction SD. Figure 16 This creates a roughly sawtooth-shaped path. Any number of first steering 252 and second steering 254 ( Figure 16 It can be used for each cooling channel 240. (e.g.) Figure 16 As shown on the left, the amplitude A of each steering 252, 254 can be consistent so as to form a sinusoidal shape with at least one first steering and second steering having the same amplitude A. Alternatively, as Figure 16 As shown on the right, the amplitude of each steering step 252, 254 can be inconsistent to create a more random zigzag path with steering steps 252, 254. Similarly... Figure 16 As shown on the right, 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 multiple branches 264, such as a tree. Any branching configuration may be adopted.

[0081] Figure 18 One embodiment is shown in which the cooling channel 240 has a curved shape, for example, within the platform 150 compared to... Figure 12 Flatter. 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 to provide film cooling to the inclined surface 230. Although shown separately, any of the cooling channel examples or aspects thereof can be combined with other examples.

[0082] Cooling circuit 234 and cooling channel 240 can be separately installed in the pressure side 232 of the platform (only) Figure 13 ), separately installed in the suction side 231 of platform 150 (only Figure 12 ), or set in the two sides 231, 232 of platform 150 ( Figures 12 to 13 If it is set on only one side of platform 150, any other conventional structure can be set on the other side of the platform. In the latter case, such as Figure 12 and Figure 13 As shown, the cooling circuit 234 may include a first portion 234SS located on the suction side 231 of the platform 150 and a second portion 234PS located on the pressure side 232 of the platform 150. The portions 234SS and 234PS may be separate or fluidly connected. In this regard, the ramp 230 includes a suction-side ramp 230SS and a pressure-side ramp 230PS. Here, the cooling passage 240 in the platform 150 may include at least one first cooling passage 240 fluidly communicating with the first portion 234SS of the cooling circuit and exiting the suction-side ramp 230SS, and at least one second cooling passage 240 fluidly communicating with the second portion 234PS of the cooling circuit and exiting the pressure-side ramp 230PS.

[0083] Compared to the linear coolant channels of conventional drilling, the nonlinear cooling channels 240 allow coolant to be directed where needed within the platform 150. The additive manufacturing of the coolant channels 140 enables a wide variety of nonlinear configurations that provide direct cooling when necessary and enhanced cooling through their shape.

[0084] D. Wing with coolant delivery channel

[0085] See Figure 2 and Figures 19 to 22 Another overall feature of the embodiments of this disclosure includes a wing 280 having a coolant delivery channel therein. Figure 19 A radial cross-section is shown through the turbine rotor blade 120, including the airfoil 280. Figure 20 A transparent perspective view of the turbine rotor blades 120, including the airfoil 280, is shown. Figure 21 An axial view of a set of turbine rotor blades 120A-C including airfoil 280 is shown, and Figure 22 A top view of the turbine rotor blade 120, including the airfoil 280, is shown. Further reference... 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 (which may be turbine rotor blades 120 together or individually). 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, the airfoil body comprising along the leading edge 128 and the trailing edge 130. Figure 3 ) connected to the concave pressure side outer wall 124 ( Figure 3 ) and convex suction side outer wall 126 ( Figure 3 The turbine rotor blade 120 may also include a shank 148 located at the radially inner end 152 of the airfoil body 122. In addition, the turbine rotor blade 120 includes at least one airfoil 280 extending laterally from at least one side 282, 284 of the shank 148.

[0086] like Figure 2 As shown in the nozzle-blade junction, an opening 286 exists at the junction between adjacent nozzles 112 and turbine rotor blades 120. This opening allows hot working fluid to exit the hot gas path and enter the wheel space 300 of the turbine 108. To limit this leakage of hot gas, the turbine rotor blades 120 typically include axially projecting airfoil seals 280, also simply referred to as "airfoil". Airfoil 280 mates with a protruding section or "hindrance" 288 extending from nozzle 112. Airfoil 280 and hinder 288 overlap (or nearly overlap) but do not contact each other, thereby restricting fluid flow.

[0087] Go to Figures 19 to 22 According to embodiments of this disclosure, the turbine rotor blade 120 may further include a coolant transfer passage 290 defined by at least one airfoil 280. For example, for the first turbine rotor blade 120A ( Figure 21 The coolant transfer channel 290 is fluidly connected to the handle 148A. Figure 21 ) and the first adjacent shank 148B of the first adjacent turbine rotor blade 120B ( Figure 21 The first round space portion 292 between and the handle 148A ( Figure 21 ) and the second adjacent turbine rotor blade 120C ( Figure 21The second wheel space portion 294 between the second adjacent handle portions 148C. (As observed...) Figure 2 , Figure 21 and Figure 22 As shown in the optimal configuration, each wheel space portion 292, 294 is part of the wheel space 300. The wheel space 300 is defined as follows: circumferentially defined in the shanks 148A-C of adjacent turbine rotor blades 120A-C. Figure 21 The blade is axially defined between the shank 148 and the adjacent nozzle 112, and radially defined by the platform 150 and the rotor disk 147. The wheel space portions 292 and 294 are wheel space portions 300 axially adjacent to the shank 148 of a particular blade.

[0088] like Figure 22 As shown, the coolant delivery channel 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. Therefore, the coolant delivery channel 290 allows wheel space coolant 316 to pass between the wheel space portions 292, 294 on circumferentially opposite sides of the handle 148. The first open end 310 and the second open end 312 may face relative to the airfoil body 122 in the axial-circumferential direction, or in any direction that would allow wheel space coolant 316 to pass between the wheel space portions 292, 294. The wheel space coolant 316 can be any coolant now known or later developed, for example from compressor 102 (… Figure 1 The coolant is guided by the airfoil body 122. As previously described, the outer walls 124, 126 of the airfoil body 122 define a radially extending chamber 134 that extends into the shank 148. Figure 22 As shown, the coolant delivery channel 290 is fluidly isolated from the radially extending chamber 134, i.e., the coolant flow 136 from chamber 134 ( Figure 4 It does not mix with wheel space coolant 316.

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

[0090] Coolant transfer channel 290 allows wheel space for coolant 316 ( Figure 22 It can move between wheel space sections 292 and 294, allowing cooling of the wing section 280 and reducing the weight of the turbine rotor blades 120.

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

[0092] See Figure 19 , Figure 20 , Figure 25 and Figure 26 Another overall feature of an embodiment of this disclosure includes a hollow blade mount 146. In this embodiment, a root 144 is provided, which includes a shank 148 having a radially extending chamber 134 defined therein. The blade mount 146 is located at the radially inner end of the shank 148. Compared to many conventional blade mounts, the blade mount 146 has a hollow interior 330, for example defined therein 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, as shown in the figure. Figure 25 As shown, the blades are radially extended. The blade mount 146 may be configured for mounting to the rotor 110. Figure 2 ) rotor wheel 147 ( Figure 21 Any external shape that is now known or later developed, such as a swallowtail or fir tree shape.

[0093] The turbine rotor blade root 144 may also 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 hollow support structure forms. In one example, the grid support structure 340 may include multiple 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 manufactured by additive manufacturing. The shank 148 and the blade mount 146 may therefore include multiple integral material layers.

[0094] According to this embodiment, the root 144, 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 As described above, platform 150 is positioned radially outward of handle 148 and extends laterally outward relative to handle, terminating at at least one ramp 230. Platform 150 may include a cooling circuit 234 defined within the platform and in fluid communication with a source of coolant flow (e.g., radially extending chamber 134). Cooling passage 240 ( Figures 12 to 18 The cooling channel 240 may be confined within the platform 150 and in fluid communication with the cooling circuit 234. As described above, the cooling channel 240 extends in a non-linear configuration from the cooling circuit 234 through the inclined surface 230 of the platform to an outlet. 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 steering along a first direction and at least one second steering along the opposite second direction ( Figure 16 Multiple branches Figure 17 ); or curved shapes (e.g., Figure 12 , Figure 14 , Figure 18 ).

[0095] According to this embodiment, the root 144, i.e., the root having the grid support structure 340, may also include a wing 280 extending laterally from at least one side of the handle 148, as described herein with respect to... Figures 19 to 24 As described above, the coolant delivery channel 290 can be defined by the wing 280. Figure 21 As shown, the coolant delivery channel 290 fluidly connects a first wheel space portion 292 defined between a first adjacent shank 148B and a first adjacent turbine rotor blade root 144B, and a second wheel space portion 294 defined between a second adjacent shank 148A and a second adjacent turbine rotor blade root 144C. The coolant delivery channel 290 includes a first opening end 310 in fluid communication with the first wheel space portion 292 and a second opening end 312 in fluid communication with the second wheel space portion 294. Figure 22 As shown, the first opening end 310 and the second opening end 312 may face in the circumferential direction relative to the handle 148. The coolant delivery channel 290 may be fluidly isolated from the radially extending chamber 134 in the handle 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 the root 144 having the shank 148, the hollow blade mount 146, the grid support structure 340, and the platform 150 and / or the wing 280, thereby forming multiple integral material layers for any feature provided.

[0097] The root 144 of the integral grid support structure 340, which is located in the hollow interior 330 of the blade mount 146, provides additional cooling for the lighter turbine rotor blades 120 and the blade mount 146.

[0098] Although various implementation schemes have been described and shown for use in this document, it should be understood that various implementation schemes may be used alone or in combination.

[0099] As used throughout the specification and claims, approximate language can be used to modify any quantitative expression that allows for variation without causing a change in its underlying function. Therefore, values ​​modified by one or more terms (such as “about,” “approximately,” and “substantially”) are not limited to specified exact values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value. Range limitations may be combined and / or interchanged herein and throughout the specification and claims; unless the context or language otherwise indicates, these ranges are identified and include all subranges contained therein. The term “about” applied to a specific value within a range applies to both terminating values ​​and may indicate + / - 10% of said value unless otherwise dependent on the precision of the instrument used to measure the value.

[0100] All means or steps plus functional elements in the following claims are intended to include any structure, material, action, and equivalent for performing a function in conjunction with other claimed elements of a particular claim. This disclosure has been described for purposes of illustration and description, but it is not intended to be exhaustive or to limit the disclosure to the forms disclosed. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of this disclosure. Embodiments have been selected and described to best explain the principles and practical application of this disclosure and to enable others skilled in the art to understand various embodiments of this disclosure with various modifications suitable for the intended particular use.

Claims

1. A turbine rotor blade (120, 120A, 120B, 120C), 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 (128) and a trailing edge (128, 130); a shank (148, 148A) at a radially inner end (152) of the airfoil body (122); at least one angel wing (280) laterally extending from at least one side (282, 284) of the shank (148, 148A); and a coolant delivery passage (290) defined through the at least one angel wing (280), the coolant delivery passage (290) fluidly coupling a first wheel space portion (292) defined between the shank (148, 148A) and a first adjacent shank (148B) of a first adjacent turbine rotor blade (120) and a second wheel space portion (294) defined between the shank (148, 148A) and a second adjacent shank (148C) of a second adjacent turbine rotor blade (120).

2. The turbine rotor blade (120) of claim 1, wherein the coolant delivery passage (290) includes a first open end (310) in fluid communication with the first wheel space portion (292) and a second open end (312) in fluid communication with the second wheel space portion (294), wherein the first open end (310) and the second open end (312) face in a circumferential direction relative to the airfoil body (122).

3. The turbine rotor blade (120) of claim 1, wherein the outer walls (124, 126) define a radially extending chamber (134) that extends into the shank (148, 148A), and wherein the coolant delivery passage (290) is fluidly isolated from the radially extending chamber (134).

4. The turbine rotor blade (120) of claim 1, wherein the at least one angel wing (280) laterally extending from the at least one side (282, 284) of the shank (148, 148A) includes a first pair of radially spaced apart angel wings (280) laterally extending from a first side (282) of the shank (148, 148A), each angel wing (280) of the first pair including a respective coolant delivery passage (290).

5. The turbine rotor blade (120) of claim 4, wherein the at least one angel wing portion (280) extending laterally from the at least one side (282, 284) of the shank portion (148, 148A) includes a second pair of radially spaced apart angel wing portions (280) extending laterally from an opposite second side (284) of the shank portion (148, 148A), each angel wing portion (280) of the second pair including a respective coolant routing passage (290).

6. The turbine rotor blade (120) of claim 1, wherein the at least one angel wing portion (280) extending laterally from the at least one side (282, 284) of the shank portion (148, 148A) includes a first angel wing portion (280) extending laterally from a first side (282) of the shank portion (148, 148A) and a second angel wing portion (280) extending laterally from an opposite second side (284) of the shank portion (148, 148A), each of the first angel wing portion (280) and the second angel wing portion (280) including a respective coolant routing passage (290).

7. An additively manufactured 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 (128) and a trailing edge (130); a shank portion (148, 148A) at a radially inner end (152) of the airfoil body (122); at least one angel wing portion (280) extending laterally from at least one side (282, 284) of the shank portion (148, 148A); and a coolant routing passage (290) defined through the at least one angel wing portion (280), the coolant routing passage (290) fluidly coupling a first wheel space portion (292) defined between the shank portion (148, 148A) and a first adjacent shank portion (148B) of a first adjacent turbine rotor blade (120) and a second wheel space portion (294) defined between the shank portion (148, 148A) and a second adjacent shank portion (148C) of a second adjacent turbine rotor blade (120).

8. The additively manufactured turbine rotor blade (120) of claim 7, wherein the coolant routing 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), wherein the first open end (310) and the second open end (312) face in a circumferential direction relative to the airfoil body (122). ​ 9. The additively manufactured turbine rotor blade (120) of claim 7, wherein the outer wall (124, 126) defines a radially extending chamber (134) that extends into the shank (148, 148A), and wherein the coolant delivery passage (290) is fluidly isolated from the radially extending chamber (134).

10. The additively manufactured turbine rotor blade (120) of claim 7, wherein the at least one angel wing (280) extending laterally from the at least one side (282, 284) of the shank (148, 148A) includes a first pair of radially spaced apart angel wings (280) extending laterally from a first side (282) of the shank (148, 148A), each angel wing (280) of the first pair including a respective coolant delivery passage (290).

11. The additively manufactured turbine rotor blade (120) of claim 10, wherein the at least one angel wing (280) extending laterally from the at least one side (282, 284) of the shank (148, 148A) includes a second pair of radially spaced apart angel wings (280) extending laterally from an opposite second side (284) of the shank (148, 148A), each angel wing (280) of the second pair including a respective coolant delivery passage (290).

12. The additively manufactured turbine rotor blade (120) of claim 7, wherein the at least one angel wing (280) extending laterally from the at least one side (282, 284) of the shank (148, 148A) includes a first angel wing (280) extending laterally from a first side (282) of the shank (148, 148A) and a second angel wing (280) extending laterally from an opposite second side (284) of the shank (148, 148A), each of the first angel wing (280) and the second angel wing (280) including a respective coolant delivery passage (290).

13. A set of turbine rotor blades (120) comprising: a first turbine rotor blade (120A), a second turbine rotor blade (120B), and a third turbine rotor blade (120C), the first turbine rotor blade (120A) positioned between the second turbine rotor blade (120B) and the third turbine rotor blade (120C), each 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 (128) and a trailing edge (130); a shank (148, 148A) at a radially inner end (152) of the airfoil body (122); and at least one angel wing (280) extending laterally from at least one side (282, 284) of the shank (148, 148A), wherein the shanks (148A-C) of the first turbine rotor blade (120A) and the second turbine rotor blade (120B) define a first wheel space portion (292) between the shanks, and the shanks (148A-C) of the first turbine rotor blade (120A) and the third turbine rotor blade (120C) define a second wheel space portion (294) between the shanks; and a coolant transfer passage (290) defined through the at least one angel wing portion (280) in the first turbine rotor blade (120A), the coolant transfer passage (290) fluidly coupling the first wheel space portion (292) and the second wheel space portion (294).

14. The set of turbine rotor blades (120) of claim 13, wherein 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), and wherein the first open end (310) and the second open end (312) face in a circumferential direction relative to the airfoil body (122) of the first turbine rotor blade (120A).

15. The set of turbine rotor blades (120) of claim 13, wherein the outer wall (124, 126) defines a radially extending chamber (134) that extends into the shank (148, 148A) of the first turbine rotor blade (120A), and wherein the coolant transfer passage (290) is fluidly isolated from the radially extending chamber (134).

16. The set of turbine rotor blades (120) of claim 13, wherein the at least one angel wing portion (280) laterally extending from the at least one side (282, 284) of the shank (148, 148A) of the first turbine rotor blade (120A) includes a first pair of radially spaced apart angel wing portions (280) laterally extending from a first side (282) of the shank (148, 148A), each angel wing portion (280) of the first pair including a respective coolant transfer passage (290).

17. The set of turbine rotor blades (120) of claim 16, wherein the at least one angel wing portion (280) laterally extending from the at least one side (282, 284) of the shank (148, 148A) of the first turbine rotor blade (120A) includes a second pair of radially spaced apart angel wing portions (280) laterally extending from an opposing second side (284) of the shank (148, 148A), each angel wing portion (280) of the second pair including a respective coolant transfer passage (290).

18. The set of turbine rotor blades (120) of claim 13, wherein the at least one angel wing portion (280) laterally extending from the at least one side (282, 284) of the shank portion (148, 148A) of the first turbine rotor blade (120A) includes a first angel wing portion (280) laterally extending from a first side (282) of the shank portion (148, 148A) and a second angel wing portion (280) laterally extending from an opposite second side (284) of the shank portion (148, 148A), each of the first angel wing portion (280) and the second angel wing portion (280) including a respective coolant routing passage (290).

19. The set of turbine rotor blades (120) of claim 13, wherein the at least one angel wing portion (280) of the second turbine rotor blade (120B) and the third turbine rotor blade (120C) each includes a coolant routing passage (290) defined through the at least one angel wing portion, the coolant routing passage (290) fluidly coupling wheel space portions (292, 294) on opposite sides (282, 284) of the shank portion (148, 148A) of each respective turbine rotor blade (120).

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