Cast turbine nozzle with heat transfer protrusions on the inner surface of the leading edge

Abstract

The present invention is entitled "Cast Turbine Nozzle with Heat Transfer Protrusions on the Inner Surface of the Leading Edge". The present invention discloses a cast turbine (108) nozzle (112) comprising an airfoil (130) having a body (128) and a cooling cavity (150) defined by an inner surface (152) of the body (128). The body includes a suction side (132), a pressure side (134) opposite to the suction side (132), a leading edge (136) bridging the pressure side (134) and the suction side (132), and a trailing edge (138) opposite to the leading edge (136) and bridging the pressure side (134) and the suction side (132). The nozzle (112) further includes a plurality of heat transfer protrusions (160) extending inwardly within the body (128) from the leading edge (136) in a radially staggered columnar pattern.

Description

Technical Field

[0001] This disclosure relates generally to turbines, and more specifically to cast turbine nozzles having heat transfer protrusions on the inner surface of the leading edge of a cooling cavity in an airfoil. Background Technology

[0002] The turbine nozzle includes a cooling chamber within the airfoil body to guide coolant to cool the airfoil. The cooling chamber provides space for an impact cooling sleeve that guides coolant against the inner surface of the airfoil body defining the cooling chamber. In some nozzle stages, it is advantageous to have a smaller radius at the leading edge of the turbine nozzle, which narrows the airfoil. A narrower airfoil makes it more difficult to maintain cooling using conventional impact cooling. Summary of the Invention

[0003] A first aspect of this disclosure provides a cast turbine nozzle comprising: an airfoil having a body and a cooling cavity defined by an inner surface of the body, the body including a suction side, a pressure side opposite to the suction side, a leading edge bridging the pressure side and the suction side, and a trailing edge opposite to the leading edge and bridging the pressure side and the suction side; at least one end wall connected to the airfoil along the suction side, the pressure side, the trailing edge, and the leading edge; and a plurality of heat transfer protrusions extending inwardly from the inner surface of the body within the cooling cavity, the plurality of heat transfer protrusions extending from the leading edge along the suction side and along the pressure side in a radially staggered columnar pattern, wherein the inner surface includes a flat surface extending between adjacent heat transfer protrusions.

[0004] A second aspect of this disclosure provides a nozzle section for a turbine having a set of nozzles including at least one cast nozzle having: an airfoil having a body and a cooling cavity defined by an inner surface of the body, the body including a suction side, a pressure side opposite to the suction side, a leading edge bridging the pressure side and the suction side, and a trailing edge opposite to the leading edge and bridging the pressure side and the suction side; at least one end wall connected to the airfoil along the suction side, the pressure side, the trailing edge, and the leading edge; and a plurality of heat transfer protrusions extending inwardly from the inner surface of the body within the cooling cavity, the plurality of heat transfer protrusions extending from the leading edge along the suction side and along the pressure side in a radially staggered columnar pattern, wherein the inner surface includes a flat surface extending between adjacent heat transfer protrusions.

[0005] A third aspect of this disclosure provides a turbine including a plurality of cast turbine nozzles, each cast turbine nozzle comprising: an airfoil having a body and a cooling cavity defined by an inner surface of the body, the body including a suction side, a pressure side opposite to the suction side, a leading edge bridging the pressure side and the suction side, and a trailing edge opposite to the leading edge and bridging the pressure side and the suction side; at least one end wall connected to the airfoil along the suction side, the pressure side, the trailing edge and the leading edge; and a plurality of heat transfer protrusions extending inwardly from the inner surface of the body within the cooling cavity, the plurality of heat transfer protrusions extending from the leading edge along the suction side and along the pressure side in a radially staggered columnar pattern, wherein the inner surface includes a flat surface extending between adjacent heat transfer protrusions.

[0006] Exemplary aspects of this disclosure are designed to address the problems described herein and / or other problems not discussed herein. Attached Figure Description

[0007] 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:

[0008] Figure 1 This is a schematic diagram of an exemplary turbine in the form of a combustion turbine or gas turbine (GT) system according to various embodiments of this disclosure;

[0009] Figure 2 Is it possible to... Figure 1 A cross-sectional illustration of an example gas turbine assembly with four-stage turbines used in conjunction with turbines in the image;

[0010] Figure 3 A schematic perspective view of a pair of exemplary turbine nozzles including an airfoil with heat transfer protrusions, according to various embodiments of the present disclosure, is shown.

[0011] Figure 4 A perspective view of an exemplary impact sleeve for use with a turbine nozzle according to an embodiment of the present disclosure is shown;

[0012] Figure 5 A top perspective view of a pair of cast turbine nozzles in a turbine nozzle section according to an embodiment of the present disclosure is shown;

[0013] Figure 6 A slightly enlarged top perspective view of a cast turbine nozzle according to an embodiment of the present disclosure is shown;

[0014] Figure 7 A perspective view of a plurality of heat transfer protrusions according to an embodiment of the present disclosure is shown;

[0015] Figure 8A plan view of the inner surface of the cooling cavity at the top of the heat transfer protrusion according to an embodiment of the present disclosure is shown; and

[0016] Figure 9 The following is illustrated in the embodiments according to this disclosure. Figure 8 A cross-sectional side view of the heat transfer protrusion taken along centerline 9-9.

[0017] 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

[0018] First, in order to clearly describe the subject matter disclosed herein, it will be necessary to select certain terms when referring to and describing the relevant machine parts within a turbine. To the extent possible, common industry terms will be used and adopted in a manner consistent with their accepted meaning. 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.

[0019] 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 relative to fluid flow (such as coolant in the space behind the airfoil, or airflow through the combustor, for example). The term “downstream” corresponds to the direction of fluid flow, and the term “upstream” refers to the direction opposite to the flow. Without further detail, the terms “front” and “rear” refer to directions, where “front” refers to the front end of the engine or compressor end, and “rear” refers to the rear section of the turbine.

[0020] It is often necessary to describe parts positioned at different radial locations relative to the central axis. The term "radial" refers to movement or position perpendicular to the axis. For example, if a first part is closer to the axis than a second part, this document will describe the first part as "radially inward" or "inner" of the second part. On the other hand, if the first part resides further away from the axis than the second part, this document may describe the first part as "radially outward" or "outer" of the second part. The term "axial" refers to movement or position parallel to the axis. Finally, the term "circumferential" refers to movement or position about the axis. It should be understood that such terms can be applied relative to the central axis of the turbine.

[0021] 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.

[0022] 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, or an element or feature subsequently described may or may not be present, and the description includes instances where the event occurs (or the feature is present) and instances where it does not occur (or is not present).

[0023] When an element or layer is referred to as “on another element or layer,” “attached to another element or layer,” “connected to another element or layer,” or “linked to another element or layer,” it may be directly on, attached to, connected to, or linked to another element or layer, or an intervening element or layer may be present. In contrast, when an element is referred to as “directly on another element or layer,” “directly attached to another element or layer,” “directly connected to another element or layer,” or “directly linked to another element or layer,” an intervening element or layer may not be present. Other terms used to describe the relationship 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.

[0024] Embodiments of this disclosure provide a cast turbine nozzle, a turbine nozzle section, and a turbine. The turbine nozzle includes a plurality of heat transfer protrusions on the inner surface of the cooling chamber in its airfoil. The heat transfer protrusions provide an improved cooling effect to maintain part life, turbine efficiency, and power output. More specifically, compared to a flat, unreinforced surface, the heat transfer protrusions (or “bulges”) increase the surface area inside the airfoil and provide additional heat transfer effects by turbulent airflow and “snapping” boundary laminar flow, thereby increasing energy exchange (heat transfer). The heat transfer protrusions are applied only to a portion of the airfoil body, i.e., the region including and surrounding the leading edge, to prevent overheating downstream of the leading edge of the narrower airfoil.

[0025] Refer to the attached diagram. Figure 1 This is a schematic view of an exemplary turbine 90 in the form of a combustion turbine or gas turbine (GT) system 100 (hereinafter “GT system 100”). The GT system 100 includes a compressor 102 and a combustor 104. The combustor 104 includes a combustion zone 105 and a head assembly 106, which includes one or more fuel nozzles. The GT system 100 also includes a turbine 108 and a conventional compressor / turbine shaft 110 (hereinafter referred to as “rotor 110”). In one embodiment, the GT system 100 is a 6F.03FL18 engine, available from General Electric Company, Greenville, SC, South Carolina. This disclosure is not limited to any particular GT system and can be incorporated 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. Furthermore, the teachings of this disclosure are not necessarily limited to GT systems and can be applied to blades and / or nozzles of other types of turbines, such as steam turbines, jet engines, compressors, etc.

[0026] During operation, air flows through compressor 102, and compressed air is supplied to burner 104. Specifically, compressed air is supplied to fuel nozzles integrated with burner 104 in head assembly 106. Head assembly 106 is in fluid communication with combustion zone 105. The fuel nozzles in head assembly 106 are also connected to a fuel source ( Figure 1 (Not shown in the image) is in fluid communication, and fuel nozzles deliver fuel and air to combustion zone 105. Combustor 104 ignites and burns the fuel to generate combustion products. In an exemplary embodiment, there are multiple combustors 104 having head-end assemblies 106, each head-end assembly 106 having one or more fuel nozzles. Combustor 104 is in fluid communication with turbine 108, within which the thermal energy of the combustion product gas vapor is converted into mechanical rotational energy.

[0027] Turbine 108 is rotatably coupled to and drives rotor 110. Compressor 102 is also rotatably coupled to rotor 110. At least one end of rotor 110 may extend axially away from turbine 108 and may be attached to a load or machinery (not shown), such as, but not limited to, a generator, a load compressor, and / or another turbine.

[0028] Figure 2 It shows that it can be used with Figure 1 The image shows an exemplary cross-sectional view of a turbine 108 with four stages L0-L3, used in the GT system 100. These four stages are referred to as L0, L1, L2, and L3. Stage L0 is the first stage and the smallest of the four stages (in the radial direction). Stage L1 is the second stage following the first stage in the axial direction. Stage L2 is the third stage and the next stage following the second stage in the axial direction. Stage L3 is the fourth and final stage in the axial direction, and its blades are the largest (in the radial direction). It should be understood that the four stages are shown only as an example, and each turbine may have more or fewer than four stages.

[0029] A set of stationary blades or nozzles 112 cooperates with a set of rotating blades 114 to form each stage L0-L3 of the turbine 108 and define a portion of the flow path through the turbine 108. The rotating blades 114 in each set are coupled to a corresponding rotor wheel 116, which circumferentially connects them to the rotor 110. Figure 1 That is, multiple rotating blades 114 are mechanically coupled to each rotor wheel 116 in a circumferentially spaced manner. The stationary nozzle section 115 includes multiple stationary nozzles 112 circumferentially spaced around the rotor 110. Each nozzle 112 may include at least one end wall (or platform) 120, 122 connected to the airfoil 130. In the example shown, the nozzle 112 includes a radially outer end wall 120 and a radially inner end wall 122. The radially outer end wall 120 connects the nozzle 112 to the housing 124 of the turbine 108. In some embodiments, the stationary nozzle section 115 is a second-stage nozzle section, i.e. Figure 2 Level L1 in the middle.

[0030] Go to Figure 3 A schematic perspective view of a cast turbine nozzle (or simply nozzle) 112 according to various embodiments is shown to better illustrate the nozzle components. Figure 3 In the diagram, two nozzles 112 are shown as parts of a stationary nozzle section 115. Thus, each nozzle 112 is a stationary nozzle, forming the stationary nozzle section 115. Figure 2The nozzle 112 is a component that forms an annular band of stationary nozzles in the stage of a turbine (e.g., turbine 108), as previously described. During operation of the turbine (e.g., turbine 108), the nozzle 112 may remain stationary to direct the working fluid (e.g., combustion gas, or steam) to one or more movable blades (e.g., blade 114), thereby causing those movable blades to initiate rotation of the rotor 110. It should be understood that the nozzle 112 is configured to be coupled (mechanically coupled via fasteners, welding, slots / grooves, etc.) to a plurality of similar or different nozzles (e.g., nozzle 112 or other nozzles) to form an annular band of nozzles in the stage L0-L3 of turbine 108.

[0031] Each turbine nozzle 112 may include a body 128 having an airfoil 130 having a convex suction side 132 and a concave pressure side 134 opposite to the suction side 132. Figure 3 (Blocked in the middle). The nozzle 112 may also include a leading edge 136 bridging the pressure side 134 and the suction side 132, and a trailing edge 138 opposite to the leading edge 136 and bridging the pressure side 134 and the suction side 132. As shown, and as previously described, the nozzle 112 may also include at least one end wall 120, 122 (two shown) connected to the airfoil 130 along the suction side 132, the pressure side 134, the trailing edge 138 and the leading edge 136. In the example shown, the nozzle 112 includes a radially outer end wall 120 and a radially inner end wall 122. The radially outer end wall 120 is configured to be in the stationary nozzle section 115 ( Figure 2 Aligned on the radially outer side of the nozzle 112 and configured to connect the corresponding nozzle 112 to the turbine 108. Figure 2 ) housing 124 ( Figure 2 The radially inner end wall 122 is configured to align with the radially inner side of the stationary nozzle section 115. Figure 2 ).

[0032] In various embodiments, each nozzle 112 includes fillets 140, 142 connecting the airfoil 130 and each corresponding endwall 120, 122. The fillet 140 may include a welded or brazed fillet, which may be formed via conventional metal-inert gas (MIG) welding, tungsten-inert gas (TIG) welding, brazing, etc. The fillets 140, 142 may overlap with a portion of the airfoil 130. The degree of overlap may vary depending on the nozzle, the stage, and / or the turbine.

[0033] Each nozzle 112 according to an embodiment of this disclosure is cast, for example formed from molten material poured into a casting and hardened. The nozzle 112 may comprise any metal or metal alloy now known or later developed, such as superalloys, capable of withstanding the environment within the turbine 108.

[0034] Each nozzle 112 may also include a cooling chamber 150 having an inner surface 152 defined within the body 128. Figure 4 A perspective view is shown of an exemplary impact insert or sleeve 154 inserted into each cooling chamber 150. That is, in operation, the impact sleeve 154 is positioned within the cooling chamber 150. As shown, the impact sleeve 154 includes a plurality of holes 156 configured to guide coolant against the inner surface 152 and around a plurality of heat transfer protrusions 160 (e.g., ...). Figures 5 to 7 As understood in the art, the cooling chamber 150 is fluidly coupled to a coolant source, such as pressurized air from the compressor 102. Coolant passes through a hole 156 in the impact insert 150 to impact the inner surface 152, thereby cooling the nozzle 112. The positioner 158 can space the impact sleeve 154 from the inner surface 152 to form an impact cooling zone therebetween.

[0035] In some commercial implementations of turbine 108, it has been found advantageous to scale the nozzle 112 for use on turbine 108 of different (e.g., smaller) gas turbines 100. This results in a smaller and / or narrower nozzle 112 (and specifically, airfoil 130), leading to a progressively smaller radius of the leading edge 136. The narrower airfoil 130 makes it more difficult to cool the leading edge 136 using conventional shock cooling. For example, turbine nozzle 112 may comprise a second-stage nozzle for a 6-series gas turbine.

[0036] The embodiments disclosed herein provide a plurality of heat transfer protrusions 160 extending inwardly from the inner surface 152 within the body 128 in a radially staggered columnar pattern. The protrusions 160 are integral with the airfoil 130. Figure 5 A perspective view of a cast turbine nozzle 112 including a heat transfer protrusion 160 is shown, and Figure 6 A slightly enlarged perspective view of the cast turbine nozzle is shown, and Figure 7 An enlarged perspective view of multiple heat transfer protrusions 160 is shown. The heat transfer protrusions 160 extend in a radially staggered columnar pattern from the inner surface 152 at the leading edge 136 along the suction side 132 and along the pressure side 134. The heat transfer protrusions 160 do not extend along the entire chord length of each side 132, 134 as is conventionally done, because it has been found that doing so with a narrower airfoil 130 would cause overheating closer to the trailing edge 138 in the downstream region. Instead, the multiple heat transfer protrusions extend along the suction side 132 in the range of 28% to 32% of the arch length, and along the pressure side 134 in the range of 9% to 13% of the arch length. "Arch length" refers to the distance from the leading edge 136 to the trailing edge 138 through the center of the airfoil 130, equidistant between the suction side 132 and the pressure side 134. A rough approximation of the arch length CL is... Figure 5As shown in the diagram. The extent to which the heat transfer protrusion 160 is defined based on the percentage of the arched length will be on each side 132, 134 at a location perpendicular to the arched length. In any case, only a portion of the inner surface along each side 132, 134 is covered by the heat transfer protrusion 160, and the inner surface 152 downstream of the heat transfer protrusion 160 has no protrusion or other structure that causes turbulence of the coolant flow in the rearward direction toward the trailing edge 138. The heat transfer protrusion 160 may extend radially on each side 132, 134 to achieve the desired heat transfer. For example, they may span the entire radial length between the end walls 120, 122. In contrast, in some embodiments, the heat transfer protrusion 160 may extend radially but terminate within a range of 8 to 13 mm from one or more end walls 120, 122.

[0037] Figure 8 A plan view of the inner surface 152 of the top of the heat transfer protrusion 160 is shown, and Figure 9 It shows along Figure 8 A cross-sectional side view of the heat transfer protrusion 160, taken along centerline 9-9. (See image.) Figure 8 and Figure 9 As shown, the inner surface 152 includes a flat surface 164 extending between adjacent heat transfer protrusions 160. That is, the flat surface 164 separates adjacent heat transfer protrusions 160, wherein the inner surface 152 has no inward or outward curvature other than inward or outward curvature to form the airfoil 130. Furthermore, as... Figure 9 As shown, each heat transfer protrusion 160 may have a truncated conical cross-section over its entire height. Each heat transfer protrusion 160 has an innermost surface 170, with a cooling cavity 150 between the innermost surface and the adjacent heat transfer protrusion 160. Figures 5 to 6 The inner surface 152 of the airfoil 130 is parallel. As used herein, "innermost" refers to the part of the structure closest to the center of the airfoil 130, and "outermost" refers to the part of the structure furthest from the center of the airfoil 130. The height H of each heat transfer protrusion 160 from the inner surface 152 of the cooling cavity 150 to the innermost surface 170 of the heat transfer protrusion 160 may be in the range of 0.5 mm to 1.0 mm.

[0038] The innermost width W1 of the heat transfer protrusion 160 can range from 0.2 mm to 0.8 mm. The outermost width W2 of the heat transfer protrusion 160 can range from 0.6 mm to 1.2 mm. The outermost width W2 is wider than the innermost width W1. The ratio of the innermost width W1 of each heat transfer protrusion 160 to the outermost width W2 of each heat transfer protrusion 160 relative to the inner surface 152 is in the range of 0.2 to 0.9. Figure 9As shown, each heat transfer protrusion 160 may have a circular cross-section over its entire width. However, other non-elongated shapes are also possible. The heat transfer protrusion 160 extends from the inner surface 152 at a substantially perpendicular angle α (i.e., substantially 90°).

[0039] like Figure 6 and Figure 8 As shown, the heat transfer protrusions 160 are arranged in a radially staggered columnar pattern. Figure 8 As best shown, the radially interlaced columnar pattern of the plurality of heat transfer protrusions 160 includes a plurality of radially extending rows 176 that are radially interlaced (perpendicular on the page) with respect to each other. Figure 8 (Three are shown in the figure). Any number of rows necessary to cover the desired percentage of chord length on each side 132, 134 can be used. The first radial distance R1 between the centers of the heat transfer protrusions 160 in the same radially extending row 176 can be in the range of 0.9 mm to 1.4 mm. The second radial distance R2 between the centers of axially adjacent heat transfer protrusions 160 in adjacent radially extending rows can be in the range of 0.3 mm to 0.9 mm. The axial distance AD ​​between adjacent radially extending rows 176 of heat transfer protrusions 160 can be in the range of 0.8 mm to 1.3 mm. The angular offset distance OF between heat transfer protrusions 160 can be, for example, in the range of 0.9 mm to 1.4 mm. Although a particular radially staggered columnar pattern has been described herein, the heat transfer protrusions 160 can be arranged in an alternative staggered columnar pattern to achieve the desired heat transfer. In other embodiments, portions of the innermost width W2 of adjacent heat transfer protrusions 160 may intersect or overlap.

[0040] During operation, coolant is drawn from the impact sleeve 154 ( Figure 4 The heat transfer protrusion 160 leaves and impacts the inner surface 152 of the airfoil 150. In its presence near the leading edge 136, the heat transfer protrusion 160 induces turbulence in the coolant flow, thereby increasing its heat transfer capacity. The heat transfer protrusion 160 can extend to any radial extent and any chord percentage to provide the desired heat transfer and cooling along the leading edge 136 and in the regions near the leading edge 136 on the pressure side 134 and suction side 132.

[0041] Embodiments of this disclosure provide a cast turbine nozzle, a turbine nozzle section, and a turbine. This teaching is particularly applicable to certain second-stage nozzles with small-radius leading edges. Heat transfer protrusions provide improved cooling effects to maintain part life, turbine efficiency, and power output at product specifications. More specifically, heat transfer protrusions, or “bulges,” increase the surface area inside the airfoil compared to a flat, unreinforced surface and provide additional heat transfer effects by turbulent airflow, thereby increasing energy exchange (heat transfer). Because the heat transfer protrusions are applied only to a portion of the airfoil body, this arrangement prevents overheating downstream of the leading edge of narrower airfoils.

[0042] 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 otherwise specified by context or language, 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.

[0043] 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 cast turbine (108) nozzle (112), the cast turbine nozzle comprising: An airfoil (130) having a body (128) and a cooling cavity (150) defined by an inner surface (152) of the body (128), the body including a suction side (132), a pressure side (134) opposite to the suction side (132), a leading edge (136) spanning between the pressure side (134) and the suction side (132), and a trailing edge (138) opposite to the leading edge (136) and spanning between the pressure side (134) and the suction side (132). At least one end wall (120, 122) is connected to the airfoil (130) along the suction side (132), the pressure side (134), the trailing edge (138) and the leading edge (136); and Multiple heat transfer protrusions (160) extend inwardly from the inner surface (152) of the body (128) within the cooling cavity (150). The multiple heat transfer protrusions (160) extend from the leading edge (136) along the suction side (132) and along the pressure side (134) in a radially staggered columnar pattern. The inner surface (152) includes a flat surface (164) extending between adjacent heat transfer protrusions (160). The plurality of heat transfer protrusions (160) extend along the suction side (132) in the range of 28% to 32% of the arch length and along the pressure side (134) in the range of 9% to 13% of the arch length, the arch length representing the distance from the leading edge (136) to the trailing edge (138) through the center of the airfoil (130), and are equidistant between the suction side (132) and the pressure side (134).

2. The casting turbine (108) nozzle (112) according to claim 1, wherein the turbine (108) nozzle (112) includes a second-stage nozzle.

3. The casting turbine (108) nozzle (112) according to claim 1, wherein each of the plurality of heat transfer protrusions (160) has a truncated conical cross section over the entire height of each heat transfer protrusion.

4. The casting turbine (108) nozzle (112) according to claim 3, wherein each of the plurality of heat transfer protrusions (160) has an innermost surface (170) parallel to the inner surface (152) of the cooling cavity (150) between adjacent heat transfer protrusions (160).

5. The casting turbine (108) nozzle (112) according to claim 3, wherein each heat transfer protrusion (160) has a circular cross-section over the entire width of each heat transfer protrusion.

6. The casting turbine (108) nozzle (112) according to claim 1, the casting turbine nozzle further comprising an impact sleeve (154) within the cooling chamber (150), the impact sleeve (154) including a plurality of holes (156) configured to guide coolant against the inner surface (152) and around the plurality of heat transfer protrusions (160).

7. The casting turbine (108) nozzle (112) according to claim 1, wherein the at least one end wall (120, 122) comprises an inner end wall (122) or an outer end wall (120).

8. The casting turbine (108) nozzle (112) according to claim 1, wherein the ratio of the innermost width of each heat transfer protrusion (160) to the outermost width of each heat transfer protrusion (160) relative to the inner surface (152) is in the range of 0.2 to 0.

9.

9. The casting turbine (108) nozzle (112) according to claim 8, wherein the innermost width of each heat transfer protrusion (160) is in the range of 0.2 mm to 0.8 mm.

10. The cast turbine (108) nozzle (112) according to claim 1, wherein the height of each heat transfer protrusion (160) from the inner surface (152) is in the range of 0.5 mm to 1.0 mm.

11. The casting turbine (108) nozzle (112) according to claim 1, wherein the radially staggered columnar pattern of the plurality of heat transfer protrusions (160) comprises a plurality of radially extending rows (176) that are radially staggered relative to each other, wherein a first radial distance between the centers of the heat transfer protrusions (160) in the same radially extending row (176) is in the range of 0.9 mm to 1.4 mm, and a second radial distance between the centers of axially adjacent heat transfer protrusions (160) in adjacent radially extending rows (176) is in the range of 0.3 mm to 0.9 mm, and The axial distance between adjacent radially extending rows (176) of the heat transfer protrusion (160) is in the range of 0.8 mm to 1.3 mm.

12. A nozzle (112) section for a turbine (108), the nozzle (112) section comprising: A set of nozzles (112), the set of nozzles (112) including at least one casting nozzle, the at least one casting nozzle having: An airfoil (130) having a body (128) and a cooling cavity (150) having an inner surface (152) defined within the body (128), the body including a suction side (132), a pressure side (134) opposite to the suction side (132), a leading edge (136) spanning between the pressure side (134) and the suction side (132), and a trailing edge (138) opposite to the leading edge (136) and spanning between the pressure side (134) and the suction side (132). At least one end wall (120, 122) is connected to the airfoil (130) along the suction side (132), the pressure side (134), the trailing edge (138) and the leading edge (136); and Multiple heat transfer protrusions (160) extend inward from the inner surface (152) within the body (128), the multiple heat transfer protrusions (160) extending from the leading edge (136) along the suction side (132) and along the pressure side (134) in a radially staggered columnar pattern. The inner surface (152) includes a flat surface (164) extending between adjacent heat transfer protrusions (160). The plurality of heat transfer protrusions (160) extend along the suction side (132) in the range of 28% to 32% of the arch length and along the pressure side (134) in the range of 9% to 13% of the arch length, the arch length representing the distance from the leading edge (136) to the trailing edge (138) through the center of the airfoil (130), and are equidistant between the suction side (132) and the pressure side (134).

13. The nozzle (112) section according to claim 12, wherein the stationary nozzle section is a second-stage nozzle section.

14. The nozzle (112) section according to claim 12, wherein each of the plurality of heat transfer protrusions (160) has a truncated conical cross section over the entire height of each heat transfer protrusion.

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