阅读说明：本技术 在前缘的内表面上具有热传递突起部的铸造涡轮喷嘴 (Cast turbine nozzle with heat transfer protrusions on the inner surface of the leading edge ) 是由 S·W·纽曼 B·D·路易斯 D·达斯 于 2021-05-08 设计创作，主要内容包括：本发明题为“在前缘的内表面上具有热传递突起部的铸造涡轮喷嘴”。本发明公开了一种铸造涡轮(108)喷嘴(112),该铸造涡轮喷嘴包括翼型件(130),该翼型件(130)具有主体(128)和由该主体(128)的内表面(152)限定的冷却腔(150),该主体包括吸力侧(132)、与该吸力侧(132)相对的压力侧(134)、跨接在该压力侧(134)与该吸力侧(132)之间的前缘(136)、与该前缘(136)相对并且跨接在该压力侧(134)与该吸力侧(132)之间的后缘(138)。该喷嘴(112)还包括以径向交错的柱状图案从该前缘(136)在该主体(128)内向内延伸的多个热传递突起部(160)。(The invention provides a cast turbine nozzle with heat transfer protrusions on the inner surface of the leading edge. A cast turbine (108) nozzle (112) includes an airfoil (130), the 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 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 the leading edge (136) and spanning between the pressure side (134) and the suction side (132). The nozzle (112) also 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.)

1. A cast turbine (108) nozzle (112), comprising:

an airfoil (130), the 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 the suction side (132), a leading edge (136) spanning between the pressure side (134) and the suction side (132), a trailing edge (138) opposite the leading edge (136) and spanning between the pressure side (134) and the suction side (132);

at least one endwall (120, 122) connected with the airfoil (130) along the suction side (132), the pressure side (134), the trailing edge (138), and the leading edge (136); and

a plurality of heat transfer protrusions (160) extending inwardly from the inner surface (152) of the body (128) within the cooling cavity (150), the plurality of heat transfer protrusions (160) extending in a radially staggered columnar pattern from the leading edge (136) along the suction side (132) and along the pressure side (134),

wherein the inner surface (152) includes a planar surface (164) extending between adjacent heat transfer protrusions (160).

2. The cast turbine (108) nozzle (112) of claim 1, wherein the turbine (108) nozzle (112) comprises a second stage nozzle (112).

3. The cast turbine (108) nozzle (112) of claim 1, wherein each heat transfer protrusion (160) of the plurality of heat transfer protrusions (160) has a frustoconical cross-section throughout a height of the each heat transfer protrusion.

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

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

6. The cast turbine (108) nozzle (112) of claim 1, further comprising an impingement sleeve (154) within the cooling cavity (150), a plurality of holes (156) included in the impingement sleeve (154) configured to direct a coolant against the inner surface (152) and around the plurality of heat transfer protrusions (160).

7. The cast turbine (108) nozzle (112) of claim 1, wherein the at least one endwall (120, 122) includes an inner endwall (122) or an outer endwall (120).

8. The cast turbine (108) nozzle (112) of claim 1, wherein a ratio of an innermost width of each heat transfer protrusion (160) to an outermost width of each heat transfer protrusion (160) relative to the inner surface (152) is in a range of 0.2 to 0.9.

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

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

11. The cast turbine (108) nozzle (112) of 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 one another, wherein a first radial distance between centers of heat transfer protrusions (160) in a same radially extending row (176) is in a range of 0.9 millimeters to 1.4 millimeters and a second radial distance between centers of axially adjacent heat transfer protrusions (160) in adjacent radially extending rows (176) is in a range of 0.3 millimeters to 0.9 millimeters, and

wherein an axial distance between adjacent radially extending rows (176) of the heat transfer protrusions (160) is in a range of 0.8 to 1.3 millimeters.

12. The cast turbine (108) nozzle (112) of claim 1, wherein the plurality of heat transfer protrusions (160) extend along a suction side (132) in a range of 28% to 32% of an arcuate length and along a pressure side (134) in a range of 9% to 13% of the arcuate length.

13. A nozzle (112) section for a turbine (108), the nozzle (112) section comprising:

a set of nozzles (112), the set of nozzles (112) comprising at least one casting nozzle (112), the at least one casting nozzle (112) having:

an airfoil (130) having a body (128) and a cooling cavity (150) having an inner surface (152) defined by the body (128), the body including a suction side (132), a pressure side (134) opposite the suction side (132), a leading edge (136) spanning between the pressure side (134) and the suction side (132), a trailing edge (138) opposite the leading edge (136) and spanning between the pressure side (134) and the suction side (132);

at least one endwall (120, 122) connected with the airfoil (130) along the suction side (132), the pressure side (134), the trailing edge (138), and the leading edge (136); and

a plurality of heat transfer protrusions (160) extending inwardly from the inner surface (152) within the body (128), the plurality of heat transfer protrusions (160) extending in a radially staggered columnar pattern from the leading edge (136) along the suction side (132) and along the pressure side (134),

wherein the inner surface (152) includes a planar surface (164) extending between adjacent heat transfer protrusions (160).

14. The nozzle (112) segment of claim 13, wherein the stationary nozzle (112) segment is a second stage nozzle (112) segment.

15. The nozzle (112) section of claim 13, wherein each heat transfer protrusion (160) of the plurality of heat transfer protrusions (160) has a frustoconical cross-section over an entire height of the each heat transfer protrusion.

Technical Field

The present disclosure relates generally to turbomachines and, more particularly, to cast turbine nozzles having heat transfer protrusions on the inner surface of the leading edge of a cooling cavity in an airfoil.

Background

The turbine nozzle includes a cooling cavity in the airfoil body to direct a coolant to cool the airfoil. The cooling cavity provides space for an impingement cooling sleeve that directs coolant against an inner surface of the airfoil body that defines the cooling cavity. In certain nozzle stages, it is advantageous to make the radius of the leading edge of the turbine nozzle smaller, which narrows the airfoil. Narrower airfoils make it more difficult to maintain cooling with conventional impingement cooling.

Disclosure of Invention

A first aspect of the present 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 the suction side, a leading edge spanning between the pressure side and the suction side, a trailing edge opposite the leading edge and spanning between the pressure side and the suction side; at least one endwall connected to the airfoil along a suction side, a pressure side, a trailing edge, and a 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 in a radially staggered columnar pattern from the leading edge along the suction side and along the pressure side, wherein the inner surface includes a planar surface extending between adjacent heat transfer protrusions.

A second aspect of the present disclosure provides a nozzle segment for a turbine, the nozzle segment having a set of nozzles including at least one casting 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 the suction side, a leading edge spanning between the pressure side and the suction side, a trailing edge opposite the leading edge and spanning between the pressure side and the suction side; at least one endwall connected to the airfoil along a suction side, a pressure side, a trailing edge, and a 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 in a radially staggered columnar pattern from the leading edge along the suction side and along the pressure side, wherein the inner surface includes a planar surface extending between adjacent heat transfer protrusions.

A third aspect of the present disclosure provides a turbine comprising 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 the suction side, a leading edge spanning between the pressure side and the suction side, a trailing edge opposite the leading edge and spanning between the pressure side and the suction side; at least one endwall connected to the airfoil along a suction side, a pressure side, a trailing edge, and a 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 in a radially staggered columnar pattern from the leading edge along the suction side and along the pressure side, wherein the inner surface includes a planar surface extending between adjacent heat transfer protrusions.

Exemplary aspects of the present disclosure are designed to solve the problems herein described and/or other problems not discussed.

Drawings

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

FIG. 1 is a schematic view of an exemplary turbomachine in the form of a combustion turbine or Gas Turbine (GT) system according to various embodiments of the present disclosure;

FIG. 2 is a cross-sectional illustration of an example gas turbine assembly having four stages of turbines that may be used with the turbomachine of FIG. 1;

FIG. 3 shows schematic perspective views of a pair of exemplary turbine nozzles including airfoils having heat transfer protrusions according to various embodiments of the present disclosure;

FIG. 4 illustrates a perspective view of an exemplary impingement sleeve for use with a turbine nozzle, according to an embodiment of the present disclosure;

FIG. 5 illustrates a top perspective view of a pair of cast turbine nozzles in a turbine nozzle section according to an embodiment of the present disclosure;

FIG. 6 illustrates a slightly enlarged top perspective view of a cast turbine nozzle according to an embodiment of the present disclosure;

FIG. 7 shows a perspective view of a plurality of heat transfer protrusions according to an embodiment of the present disclosure;

FIG. 8 shows a plan view of the inner surface of the cooling cavity looking into the top of the heat transfer protrusion, according to an embodiment of the present disclosure; and is

FIG. 9 illustrates a cross-sectional side view of a heat transfer protrusion taken along line 9-9 of FIG. 8, according to an embodiment of the present disclosure.

It should be noted that the drawings of the present disclosure are not necessarily drawn to scale. The drawings are intended to depict only typical aspects of the disclosure, and therefore should not be considered as limiting the scope of the disclosure. In the drawings, like numbering represents like elements between the drawings.

Detailed Description

First, in order to clearly describe the presently disclosed subject matter, it will be necessary to select certain terms when referring to and describing the relevant machine components within the turbine. To the extent possible, the general industry terminology will be used and employed in a manner consistent with the accepted meaning of that term. Unless otherwise indicated, such terms should be given a broad interpretation consistent with the context of the application and the scope of the appended claims. One of ordinary skill in the art will appreciate that often several different or overlapping terms may be used to refer to a particular component. An object that may be described herein as a single part may comprise multiple components and in another context be referred to as being made up of multiple components. Alternatively, an object that may be described herein as comprising a plurality of components may be referred to elsewhere as a single part.

Furthermore, several descriptive terms may be used regularly herein, and it should prove helpful to define these terms at the beginning of this section. Unless otherwise indicated, these terms and their definitions are as follows. As used herein, "downstream" and "upstream" are terms that indicate a direction relative to a fluid flow (such as coolant in an impingement space in an airfoil, or air flow through a combustor, for example). The term "downstream" corresponds to the direction of fluid flow, and the term "upstream" refers to the direction opposite to flow. Without any further details, the terms "forward" and "aft" refer to directions, wherein "forward" refers to the forward or compressor end of the engine and "aft" refers to the aft section of the turbine.

It is often desirable to describe components that are disposed at different radial positions relative to a central axis. The term "radial" refers to movement or position perpendicular to an axis. For example, if a first component is closer to an axis than a second component, the first component will be described herein as being "radially inward" or "inboard" of the second component. On the other hand, if the first component resides farther from the axis than the second component, it may be described herein that the first component is "radially outward" or "outboard" of the second component. The term "axial" refers to movement or position parallel to an axis. Finally, the term "circumference" refers to movement or position about an axis. It should be understood that such terms may apply with respect to a central axis of the turbine.

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

The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the disclosure. As used herein, the singular forms "a", "an" and "the" are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms "comprises" and/or "comprising," when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and/or groups thereof. "optional" or "optionally" means that the subsequently described event or circumstance may or may not occur, or that the subsequently described element or feature may or may not be present, and that the description includes instances where the event occurs (or the feature is present) and instances where it does not occur (or is not present).

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

Embodiments of the present disclosure provide a cast turbine nozzle, a turbine nozzle segment, and a turbine. The turbine nozzle includes a plurality of heat transfer protrusions on an inner surface of a cooling cavity in an airfoil thereof. The heat transfer protrusions provide improved cooling effects to maintain part life, turbine efficiency, and power output. More specifically, the heat transfer protrusions (or "lobes") increase the surface area inside the airfoil relative to a flat, non-enhanced surface, and provide additional heat transfer effects by disrupting airflow and "tripping" boundary layer flow, thereby increasing energy exchange (heat transfer). The heat transfer protrusions are applied to only a portion of the airfoil body, i.e., the area including and surrounding the leading edge, to prevent overheating downstream of the leading edge of the narrower airfoil.

Referring to the drawings, FIG. 1 is a schematic view of an exemplary turbomachine 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-end assembly 106 that includes one or more fuel nozzles. The GT system 100 also includes a turbine 108 and a common compressor/turbine shaft 110 (hereinafter "rotor 110"). In one embodiment, the GT system 100 is a 6f.03fl18 engine, available from General Electric Company, Greenville, s.c. The present disclosure is not limited to any one particular GT system and may be implanted with other engines, including, for example, the generic electric company class HA, F, B, LM, GT, TM, and E engines, and other company engine models. Furthermore, the teachings of the present disclosure are not necessarily applicable only to GT systems, and may be applied to blades and/or nozzles of other types of turbines (e.g., steam turbines, jet engines, compressors, etc.).

In operation, air flows through compressor 102 and compressed air is supplied to combustor 104. Specifically, the compressed air is supplied to fuel nozzles integrated with combustor 104 in head-end assembly 106. Head end assembly 106 is in fluid communication with combustion zone 105. The fuel nozzles in head-end assembly 106 are also in fluid communication with a fuel source (not shown in FIG. 1), and the fuel nozzles pass fuel and air to combustion zone 105. The combustor 104 ignites and combusts fuel to generate combustion products. In the exemplary embodiment, there is a plurality of combustors 104 having head-end assemblies 106, with each head-end assembly 106 having one or more fuel nozzles. The combustor 104 is in fluid communication with a turbine 108, where the gas steam heat energy from the combustion products is converted to mechanical rotational energy.

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, an electrical generator, a load compressor, and/or another turbine.

Fig. 2 shows a cross-sectional view of an illustrative portion of a turbine 108 having four stages L0-L3 that may be used with the GT system 100 of fig. 1. The four stages are referred to as L0, L1, L2, and L3. The stage L0 is the first stage and is the smallest stage (in the radial direction) of the four stages. Stage L1 is a second stage after the first stage in the axial direction. Stage L2 is the third stage and is the next stage after the second stage in the axial direction. Stage L3 is the fourth, last stage in the axial direction, and has its vanes largest (in the radial direction). It should be understood that four stages are shown as an example only, and that each turbine may have more or less than four stages.

A set of stationary vanes or nozzles 112 cooperate with a set of rotating vanes 114 to form each stage L0-L3 of turbine 108 and define a portion of a flow path through turbine 108. The rotating blades 114 in each set are coupled to a respective rotor wheel 116 that couples them circumferentially to the rotor 110 (FIG. 1). That is, a plurality of rotating blades 114 are mechanically coupled to each rotor wheel 116 in a circumferentially spaced manner. The stationary nozzle section 115 includes a plurality of stationary nozzles 112 circumferentially spaced around the rotor 110. Each nozzle 112 may include at least one endwall (or platform) 120, 122 connected with an airfoil 130. In the example shown, the nozzle 112 includes a radially outer end wall 120 and a radially inner end wall 122. Radially outer endwall 120 couples nozzle 112 to a casing 124 of turbine 108. In certain embodiments, the stationary nozzle section 115 is a second stage nozzle section, i.e., stage L1 in fig. 2.

Turning to FIG. 3, a schematic perspective view of a cast turbine nozzle (or simply nozzle) 112 is shown to better illustrate components of the nozzle, in accordance with various embodiments. In fig. 3, two nozzles 112 are shown as part of a stationary nozzle section 115. As such, each nozzle 112 is a stationary nozzle that forms part of a stationary nozzle section 115 (FIG. 2) and forms part of an annulus of the stationary nozzle that is a stage of a turbine (e.g., turbine 108), as previously described. During operation of a turbine (e.g., turbine 108), nozzle 112 may remain stationary in order to direct a flow of a working fluid (e.g., combustion gases, which may also be steam) to one or more movable blades (e.g., blades 114), thereby causing those movable blades to initiate rotation of rotor 110. It should be appreciated that the nozzle 112 is configured to be coupled (mechanically coupled via fasteners, welds, slots/grooves, etc.) with a plurality of similar or different nozzles (e.g., the nozzle 112 or other nozzles) to form an annulus that is a nozzle of stages L0-L3 of the turbine 108.

Each turbine nozzle 112 may include a body 128 having an airfoil 130 with a convex suction side 132 and a concave pressure side 134 (blocked in FIG. 3) opposite the suction side 132. The nozzle 112 may further include a leading edge 136 spanning between the pressure side 134 and the suction side 132, and a trailing edge 138 opposite the leading edge 136 and spanning between the pressure side 134 and the suction side 132. As shown, and as previously described, the nozzle 112 may also include at least one endwall 120, 122 (two shown) connected to the airfoil 130 along a suction side 132, a pressure side 134, a trailing edge 138, and a 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 endwall 120 is configured to align radially outward of the stationary nozzle segment 115 (FIG. 2) and to couple the respective nozzle 112 to a casing 124 (FIG. 2) of the turbine 108 (FIG. 2). The radially inner endwall 122 is configured to align on a radially inner side of the stationary nozzle segment 115 (fig. 2).

In various embodiments, each nozzle 112 includes fillets 140, 142 connecting the airfoil 130 and each respective endwall 120, 122. The fillet 140 may comprise a welded or brazed fillet, which may be formed via conventional metal-inert gas (MIG) welding, tungsten-inert gas (TIG) welding, brazing, or the like. The fillets 140, 142 may overlap a portion of the airfoil 130. The degree of overlap may vary from nozzle to nozzle, stage to stage, and/or turbine to turbine.

Each nozzle 112 according to embodiments of the present disclosure is cast, e.g., formed from a molten material poured into a casting and hardened. Nozzle 112 may include any now known or later developed metal or metal alloy, such as a superalloy, capable of withstanding the environment within turbine 108.

Each nozzle 112 may also include a cooling cavity 150 having an inner surface 152 defined within the body 128. FIG. 4 illustrates a perspective view of an exemplary impingement insert or sleeve 154 inserted into each cooling cavity 150. That is, in operation, the impingement sleeve 154 is positioned within the cooling cavity 150. As shown, a plurality of holes 156 are included in the impingement sleeve 154 and are configured to direct the coolant against the inner surface 152 and around a plurality of heat transfer protrusions 160 (see FIGS. 5-7). As understood in the art, the cooling cavity 150 is fluidly coupled to a source of coolant, such as pressurized air from the compressor 102. The coolant passes through the holes 156 in the impingement insert 150 to impinge on the inner surface 152, thereby cooling the nozzle 112. A locator 158 may space the impingement sleeve 154 from the inner surface 152 to form an impingement cooling zone therebetween.

In certain commercial embodiments of the turbine 108, it has been found advantageous to scale the nozzle 112 for use on the turbine 108 of a different (e.g., smaller) gas turbine 100. Accordingly, the nozzle 112 (and specifically the airfoil 130) is made smaller and/or narrower in size, which results in the radius of the leading edge 136 becoming smaller and smaller. The narrower airfoil 130 makes it more difficult to cool the leading edge 136 with conventional impingement cooling. For example, the turbine nozzle 112 may include a second stage nozzle for a 6-series gas turbine.

Embodiments of the present disclosure 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 protrusion 160 is integral with the airfoil 130. FIG. 5 illustrates a perspective view of a cast turbine nozzle 112 including heat transfer protrusions 160, and FIG. 6 illustrates a slightly enlarged perspective view of the cast turbine nozzle, and FIG. 7 illustrates an enlarged perspective view of a plurality of heat transfer protrusions 160. Heat transfer protrusions 160 extend in a radially staggered columnar pattern from inner surface 152 at leading edge 136 along suction side 132 and along pressure side 134. The heat transfer protrusions 160 do not extend as conventionally along the entire chordal length of each side 132, 134, as it has been found that doing so with a narrower airfoil 130 causes overheating in the downstream region closer to the trailing edge 138. Instead, the plurality of heat transfer protrusions extend along the suction side 132 in a range of 28% to 32% of the arch length and along the pressure side 134 in a 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 arcuate length CL is shown in fig. 5. The extent of the arcuate length of the heat transfer protrusions 160 based on the percentage will be defined on each side 132, 134 at a location perpendicular to the arcuate length. In any event, only a portion of the inner surface along each side 132, 134 is covered by the heat transfer protrusions 160, and the inner surface 152 downstream of the heat transfer protrusions 160 is free of protrusions or other structures that cause turbulence in the coolant flow in the aft direction toward the trailing edge 138. The heat transfer protrusions 160 may extend any radial extent 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 certain embodiments, the heat transfer protrusions 160 may extend radially, but stop in a range of 8 to 13 millimeters from one or more of the end walls 120, 122.

FIG. 8 shows a plan view looking at the inner surface 152 of the top of the heat transfer protrusion 160, and FIG. 9 shows a cross-sectional side view of the heat transfer protrusion 160 taken along line 9-9 of FIG. 8. As shown in fig. 8 and 9, the inner surface 152 includes a flat surface 164 extending between adjacent heat transfer protrusions 160. That is, the flat surfaces 164 separate adjacent heat transfer protrusions 160, wherein the inner surface 152 is free of inward or outward curves other than inward or outward curves to form the airfoil 130. Further, as shown in fig. 9, each heat transfer protrusion 160 may have a frustoconical cross-section throughout the height of each heat transfer protrusion. Each heat transfer protrusion 160 has an innermost surface 170 that is parallel to the inner surface 152 of the cooling cavity 150 (fig. 5-6) between adjacent heat transfer protrusions 160. As used herein, "innermost" indicates a portion of the structure closest to the center of the airfoil 130, and "outermost" indicates a portion 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 millimeters to 1.0 millimeter.

The innermost width W1 of the heat transfer protrusion 160 may be in the range of 0.2 millimeters to 0.8 millimeters. The outermost width W2 of the heat transfer protrusion 160 may be in the range of 0.6 millimeters to 1.2 millimeters. 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. As shown in fig. 9, each heat transfer protrusion 160 may have a circular cross-section over the entire width of each heat transfer protrusion. However, other non-elongated shapes are also possible. The heat transfer protrusions 160 extend from the inner surface 152 at a substantially perpendicular angle a (i.e., substantially 90).

As shown in fig. 6 and 8, the heat transfer protrusions 160 are arranged in a radially staggered columnar pattern. As best shown in fig. 8, the radially staggered columnar pattern of the plurality of heat transfer protrusions 160 includes a plurality of radially extending rows 176 (three shown in fig. 8) that are radially staggered (perpendicular on the page) relative to each other. Any number of rows necessary to cover a desired percentage of chord length on each side 132, 134 may be used. A first radial distance R1 between centers of the heat transfer protrusions 160 in the same radially extending row 176 may be in a range of 0.9 millimeters to 1.4 millimeters. A second radial distance R2 between centers of axially adjacent heat transfer protrusions 160 in adjacent radially extending rows may be in a range of 0.3 millimeters to 0.9 millimeters. The axial distance AD between adjacent radially extending rows 176 of heat transfer protrusions 160 may be in the range of 0.8 millimeters to 1.3 millimeters. The angular offset distance OF between the heat transfer protrusions 160 may be in the range OF, for example, 0.9 millimeters to 1.4 millimeters. Although specific radially staggered columnar patterns have been described herein, the heat transfer protrusions 160 may be arranged in alternative staggered columnar patterns to achieve the desired heat transfer. In other embodiments, portions of the innermost widths W2 of adjacent heat transfer protrusions 160 may intersect or overlap.

In operation, coolant exits from the impingement sleeve 154 (FIG. 4) and impinges against the inner surface 152 of the airfoil 150. In the presence of the leading edge 136, the heat transfer protrusions 160 induce turbulence in the coolant flow, thereby increasing its heat transfer capability. The heat transfer protrusions 160 may extend any radial extent and any chord percentage to provide desired heat transfer and cooling along the leading edge 136 and in areas of the pressure side 134 and the suction side 132 proximate the leading edge 136.

Embodiments of the present disclosure provide a cast turbine nozzle, a turbine nozzle segment, and a turbine. The teachings are particularly applicable to certain second stage nozzles having a smaller radius leading edge. The heat transfer protrusions provide improved cooling effects to maintain part life, turbine efficiency, and power output for product specifications. More specifically, the heat transfer protrusions or "lobes" increase the surface area inside the airfoil relative to a flat, non-enhanced surface and provide an additional heat transfer effect by perturbing the airflow, thereby increasing the energy exchange (heat transfer). This arrangement prevents overheating downstream of the leading edge of the narrower airfoil, since the heat transfer protrusions are applied to only a portion of the airfoil body.

Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms (such as "about", "about" and "substantially") is not to be limited to the precise value specified. In at least some cases, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations may be combined and/or interchanged. Unless context or language indicates otherwise, such ranges are identified and include all sub-ranges subsumed therein. "about" as applied to a particular value of a range applies to both end values, which may indicate +/-10% of the value unless otherwise dependent on the accuracy of the instrument measuring the value.

The corresponding structures, materials, acts, and equivalents of all means or step plus function elements in the claims below are intended to include any structure, material, or act for performing the function in combination with other claimed elements as specifically claimed. The description of the present disclosure has been presented for purposes of illustration and description, but is not intended to be exhaustive or limited to the disclosure in the form disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the disclosure. The embodiments were chosen and described in order to best explain the principles of the disclosure and the practical application, and to enable others of ordinary skill in the art to understand the disclosure for various embodiments with various modifications as are suited to the particular use contemplated.