Patent Description:
In a turbine, a number of stationary nozzles are arranged in an annulus to direct a working fluid towards a rotating blade stage. A working fluid, such as combustion gases or steam, is directed by the stationary vanes to impart rotation to the rotating blade stage to generate power. The pairwise throat area is an area between an inner endwall, an outer endwall and wing portions of adjacent nozzles. The pairwise throat area is typically selected to provide ideal turbine performance, for example, by directing the working fluid in a manner to impart the most power to the rotating blade stage. The pairwise throat area that provides the best turbine performance can change during operation of the turbine and can change over time as the turbine ages. It is therefore advantageous to periodically change the pairwise throat area of a turbine to improve or maintain performance levels.

One approach to change a pairwise throat area includes replacing an entire nozzle set, for example, during a periodic maintenance of a turbine. Current nozzle set replacement includes completely replacing each nozzle in the set including the endwalls and airfoil. This process is expensive because each new nozzle has to be built in its entirety. In another approach, variable nozzle assemblies allow minimal changes in geometry between wing portions to adjust the pairwise throat area, for example, by rotating the wing portions of the nozzles during operation of the turbine. Variable nozzle assemblies require complicated and expensive mechanical systems to allow for movement of the wing portions and to maintain sufficient mechanical strength for the working environment of the turbine. Further, variable nozzle assemblies can only provide a limited amount of adjustment, which may be insufficient to provide all desired throat area adjustments over a lifetime of a turbine.

<CIT>, <CIT> and <CIT> each disclose a turbine nozzle assembly system including a circumferential row of a plurality of nozzles that collectively form an annulus of a particular nozzle stage of the turbine, wherein adjacent nozzles have a different airfoil wing shape and/or are spaced differently from each other, thereby providing varying throat areas between adjacent nozzles that differ amongst pairs of adjacent nozzles in the circumferential row of the plurality of nozzles that collectively form the annulus of the particular turbine nozzle stage.

<CIT> discloses a stator assembly for a turbine diaphragm including inner and outer endwalls provided with slots for receiving the stator vanes, wherein the stator vanes are inserted through the slots of the outer endwall and into the slots of the inner endwall and secured thereto by welding or brazing.

<CIT> discloses a vane cooling system for a gas turbine engine comprising a vane arranged on a stator and having a cooling chamber extending continuously from a radially inner end to a radially outer end of the vane, the vane including a radially inner and a radially outer inlet for supplying a cooling fluid into the cooling chamber and an array of outlet holes arranged adjacent a trailing edge of the vane.

<CIT> discloses gas turbine nozzles including fillets between the airfoils and the inner and outer endwalls and including cooling passages passing through the outer endwall and into the vane.

An aspect of the invention provides a turbine nozzle assembly system, comprising: a plurality of nozzle sets, each nozzle set including a plurality of nozzles that collectively form an annulus of a particular nozzle stage of a turbine, each nozzle of a respective nozzle set including: an airfoil having: an inner endwall mount end, an outer endwall mount end, and a wing portion between the inner endwall mount end and the outer endwall mount end, an inner endwall including a first joint opening configured to receive the inner endwall mount end of the airfoil; an outer endwall including a second joint opening configured to receive the outer endwall mount end of the airfoil; wherein the inner endwall mount ends amongst the plurality of nozzle sets are identical to each other, wherein the outer endwall mount ends amongst the plurality of nozzle sets are identical to each other, wherein the wing portions of a respective nozzle set of the plurality of nozzle sets have a wing shape that is identical for each airfoil within the respective nozzle set but different than the wing shapes of the wing portions of the other nozzle sets of the plurality of nozzle sets, wherein for each nozzle set: adjacent nozzles of the plurality of nozzles define a pairwise throat area created by the respective inner endwalls, the respective outer endwalls and the respective wing portions of the adjacent airfoils in the annulus of the particular nozzle stage (L0-L3) of the turbine (<NUM>), and wherein the pairwise throat area (<NUM>) of each nozzle set of the plurality of nozzle sets is identical for each pairwise throat area (<NUM>) within the respective nozzle set (<NUM>) but different than the pairwise throat areas of the other nozzle sets of the plurality of nozzle sets.

Another aspect of the invention is that for each nozzle: the outer endwall is mounted to the outer endwall mount end of the airfoil by a first fillet, and the inner endwall is mounted to the inner endwall mount end of the airfoil by a second fillet.

Another aspect of the invention is that for each nozzle: at least one of the outer endwall mount end and the inner endwall mount end further includes a cooling passage therein, the cooling passage positioned adjacent at least a portion of the respective first or second fillet.

Another aspect of the invention is that for each nozzle: the cooling passage includes a radially facing inlet.

Another aspect of the invention is that for each nozzle: the inlet is adjacent a leading edge of the wing portion of the airfoil.

Another aspect of the invention is that for each nozzle: a trailing edge cooling passage in the wing portion and in fluid communication with the cooling passage, the trailing edge cooling passage includes a plurality of passages exiting a trailing edge of the wing portion of the airfoil.

Another aspect of the invention is that for each nozzle: the cooling passage includes a radially facing inlet, and an outlet facing into an interior cooling chamber of the airfoil.

Another aspect of the invention is for each nozzle: a plurality of cooling passages within the at least one of the outer endwall mount end and the inner endwall mount end.

Another aspect of the invention is that each of the plurality of wing shapes that are different amongst the plurality of nozzle sets has a similar radius of curvature distribution at each of a plurality of spanwise cross-sectional locations.

Another aspect relates to a method of modifying a turbine nozzle assembly using the turbine nozzle assembly system described above, the method comprising: providing a plurality of nozzles of a first nozzle set of the plurality of nozzle sets forming an annulus of a particular nozzle stage of the turbine, each pair of adjacent nozzles of the first nozzle set defining an identical first pairwise throat area; for each nozzle in the first nozzle set: removing an inner endwall mount end of the airfoil from the inner endwall of the nozzle and removing the outer endwall mount end of the airfoil from the outer endwall of the nozzle; and for each nozzle in a second nozzle set: coupling the inner endwall mount end of the airfoil of the second nozzle set to the inner endwall and coupling the outer endwall mount end of the airfoil to the outer endwall, each pair of adjacent nozzles of the second nozzle set defining an identical second pairwise throat area, wherein the second pairwise throat area of the second nozzle set is different than the first pairwise throat area of the first nozzle set.

Another aspect not encompassed by the wording of the claims is that coupling the outer endwall to the outer endwall mount end of each nozzle of the second nozzle set includes brazing to create a first fillet; and wherein the coupling the inner endwall to the inner endwall mount end of the of each nozzle of the second nozzle set includes brazing to create a second fillet.

Another aspect of the invention is that the first and second wing portions have a similar radius of curvature distribution at each of a plurality of spanwise cross-sectional locations between the inner endwall mount end and the outer endwall mount end.

The drawings are intended to depict only typical aspects of the invention and therefore should not be considered as limiting the scope of the invention.

As an initial matter, in order to clearly describe the current technology, it will become necessary to select certain terminology when referring to and describing relevant machine components within a turbomachine. To the extent possible, common industry terminology will be used and employed in a manner consistent with its accepted meaning. Unless otherwise stated, such terminology should be given a broad interpretation consistent with the context of the present invention and the scope of the appended claims. Those of ordinary skill in the art will appreciate that often a particular component may be referred to using several different or overlapping terms. What may be described herein as being a single part may include and be referenced in another context as consisting of multiple components. Alternatively, what may be described herein as including multiple components may be referred to elsewhere as a single part.

In addition, several descriptive terms may be used regularly herein, and it should prove helpful to define these terms at the onset of this section. These terms and their definitions, unless stated otherwise, are as follows. As used herein, "downstream" and "upstream" are terms that indicate a direction relative to the flow of a fluid, such as the working fluid through the turbine and by turbine blades, or, for example, the flow of air through the combustor or coolant through one of the turbine's component systems. The term "downstream" corresponds to the direction of flow of the fluid, and the term "upstream" refers to the direction opposite to the flow. Components, such as airfoils, positioned within the flow of fluids through a gas turbine may be described as having a "leading edge," which is the foremost edge of the component that first encounters the oncoming flow of fluids, and a "trailing edge" opposite the leading edge. The terms "forward" and "aft," without any further specificity, refer to directions, with "forward" referring to the front or compressor end of the engine, and "aft" referring to the rearward or turbine end of the engine.

It is often required to describe parts that are disposed at different radial positions with regard to a center axis. The term "radial" refers to movement or position perpendicular to an axis.

For example, if a first component resides closer to the axis than a second component, it will be stated herein that the first component is "radially inward" or "inboard" of the second component.

If, on the other hand, the first component resides further from the axis than the second component, it may be stated 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 A, e.g., rotor shaft <NUM>. Finally, the term "circumferential" refers to movement or position around an axis. It will be appreciated that such terms may be applied in relation to the center axis of the turbine.

The terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

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 to, connected to or coupled to the other element or layer, or intervening elements or layers may be present.

As used herein, the term "identical" relative to parts indicates the intent to have the parts (or portions thereof) have the same design and manufacturing specifications. For example, the mount ends of nozzles from different nozzle sets are intended to be exactly the same - identical. However, as-made mount end dimensions may be slightly different because of manufacturing tolerances, yet both nozzle ends would still fit satisfactorily within the corresponding endwall mount. Thus, "identical" may be construed as "retro-fittable" or "also fits.

Various aspects of the disclosure are directed toward a turbine nozzle assembly system that includes a plurality of nozzle sets. Each nozzle set includes a plurality of nozzles that collectively form a circumferential ring or annulus, i.e., of a nozzle stage in a turbomachine. The nozzles in each set include an inner endwall including a first joint opening configured to receive an inner endwall mount end of an airfoil, and an outer endwall including a second joint opening configured to receive the outer endwall mount end of the airfoil. The airfoils of the nozzles have an inner endwall mount end and an outer endwall mount end that are identical amongst the plurality of nozzle sets, so any of them can be used with the same inner and outer endwall.

A wing portion of the airfoil between the inner endwall mount end and the outer endwall mount end has a wing shape selected from a plurality of wing shapes that are identical within the respective nozzle set but different amongst each of the plurality of nozzle sets. A pairwise throat area is created by the inner endwall, the outer endwall and the wing portions of adjacent airfoils in the annulus. Rather than replace an entire nozzle to change the throat area, for each nozzle in a given nozzle set, the endwalls can be removed from the original airfoil, and the original airfoil can be replaced with an airfoil having a different wing shape. The replacement airfoil may provide a different pairwise throat area and thus a different overall throat area for the nozzle set. That is, each nozzle set of the plurality of nozzle sets provides a different wing shape that provides a different throat area (pairwise and overall) compared to each other nozzle set of the plurality of nozzle sets. The system allows changing of an overall throat area for a nozzle stage without replacing the entirety of each nozzle - only the airfoil is changed. The system is less expensive to implement than an adjustable vane system or a total replacement of all nozzles within a nozzle stage.

Referring to the drawings, <FIG> is a schematic view of an illustrative turbomachine <NUM> in the form of a combustion turbine or gas turbine (GT) system <NUM> (hereinafter "GT system <NUM>"), which may be used for electrical power generation. GT system <NUM> includes a compressor <NUM> and a combustor <NUM>. Combustor <NUM> includes a combustion region <NUM> and a fuel nozzle assembly <NUM>. GT system <NUM> also includes a turbine <NUM> and a common rotor shaft <NUM> (hereinafter referred to as "rotor shaft <NUM>").

In one non-limiting embodiment, GT system <NUM> may be a 9F. <NUM> engine, commercially available from General Electric Company, Greenville, S. However, the present invention is not limited to any one particular GT system and may be implemented in connection with other engines including, for example, other F, HA, B, LM, GT, TM and E-class engine models of General Electric Company, and engine models of other companies. Further, the teachings of the invention are not necessarily applicable to only a GT system and may be applied to other types of turbomachines, e.g., steam turbines, jet engines, compressors, etc..

<FIG> shows a cross-section view of an illustrative portion of turbine <NUM> with four stages L0-L3 that may be used with GT system <NUM> in <FIG>. The four stages are referred to as L0, L1, L2, and L3. Stage L0 is the first stage and is the smallest (in a radial direction) of the four stages. Stage L1 is the second stage and is the next stage in an axial direction, which is adjacent to and downstream of stage L0. Stage L2 is the third stage and is the next stage in an axial direction, which is adjacent and downstream of stage L1. Stage L3 is the fourth, last stage and is the largest (in a radial direction). It is to be understood that four stages are shown as one non-limiting example only, and each turbine <NUM> may have more or less than four stages.

A set of stationary nozzles <NUM> includes a plurality of nozzles <NUM> that collectively form a circumferential ring or annulus for a particular stage of turbine <NUM>. That is, set of nozzles <NUM> includes stationary nozzles <NUM> circumferentially spaced around rotor shaft <NUM>. Nozzle sets <NUM> cooperate with respective sets of rotating blades <NUM> to form each stage L0-L3 of turbine <NUM> and to define a portion of a flow path through turbine <NUM>. Rotating blades <NUM> in each set are coupled to a respective rotor wheel <NUM> that couples them circumferentially to rotor shaft <NUM> (<FIG>). That is, a plurality of rotating blades <NUM> is mechanically coupled in a circumferentially spaced manner to each rotor wheel <NUM>. As will be described in greater detail herein, each nozzle <NUM> includes outer and inner endwalls (or platforms) <NUM>, <NUM>. In the example shown, nozzle <NUM> includes a radially outer endwall <NUM> and a radially inner endwall <NUM>. Radially outer endwall <NUM> couples nozzle <NUM> to a casing <NUM> of turbine <NUM>.

Referring to <FIG>, in operation, air flows through compressor <NUM>, and compressed air is supplied to combustor <NUM>. Specifically, the compressed air is supplied to fuel nozzle assembly <NUM> that is integral to combustor <NUM>. Fuel nozzle assembly <NUM> is in flow communication with combustion region <NUM>. Fuel nozzle assembly <NUM> is also in flow communication with a fuel source (not shown in <FIG>) and channels fuel and air to combustion region <NUM>. Combustor <NUM> ignites and combusts fuel. Combustor <NUM> is in flow communication with turbine <NUM> within which gas stream thermal energy is converted to mechanical rotational energy. Turbine <NUM> is rotatably coupled to and drives rotor shaft <NUM>. Compressor <NUM> may also be rotatably coupled to rotor shaft <NUM>. In the illustrative embodiment, there are a plurality of combustors <NUM> and fuel nozzle assemblies <NUM>. In the following discussion, unless otherwise indicated, only one of each component will be discussed. At least one end of rotating rotor shaft <NUM> may extend axially away from either compressor <NUM> or turbine <NUM> 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.

<FIG> shows a schematic and expanded three-dimensional view of a turbine nozzle assembly system <NUM> (hereinafter "system <NUM>") including a nozzle <NUM>. <FIG> shows a schematic three-dimensional view of an illustrative nozzle <NUM> in an assembled form, according to embodiments of the invention. <FIG> also includes an exploded view of nozzle <NUM> in <FIG>. As described above, system <NUM> includes a plurality of nozzle sets <NUM>. For purposes of description, four nozzle sets 115A-D are schematically illustrated in <FIG>. However, any number of sets can be provided. Each nozzle set 115A-D is shown schematically as a different airfoil 128A-D with each airfoil having a different wing portion 130A-D. Each nozzle set 115A-D includes a plurality of nozzles <NUM> that collectively form a circumferential ring or annulus for a stage L0-L3 (<FIG>) of turbine <NUM> (<FIG>).

As noted, during operation of turbine <NUM> (<FIG>), nozzles <NUM> will remain stationary in order to direct the flow of working fluid (e.g., gas or steam) to one or more movable blades (e.g., blades <NUM>), causing those movable blades to initiate rotation of rotor shaft <NUM>. It is understood that nozzle <NUM> may be configured to couple (mechanically couple via fasteners, welds, slot/grooves, etc.) with a plurality of similar or distinct nozzles (e.g., nozzles <NUM> or other nozzles) to form an annulus of nozzles in a stage L0-L3 (<FIG>) of turbine <NUM> (<FIG>).

According to embodiments of the invention, each nozzle set 115A-D of the plurality of nozzle sets <NUM> provides a different pairwise throat area compared to each other nozzle set of the plurality of nozzle sets. Consequently, each nozzle set <NUM> also has a different overall throat area, i.e., the sum of all pairwise throat areas in the nozzle set. The different pairwise throat area is created by using airfoils 128A-D having identical inner endwall mount ends <NUM> and identical outer endwall mount ends <NUM>, but different wing portions 130A-D between mount ends <NUM>, <NUM>.

As shown in <FIG>, each nozzle <NUM> includes an airfoil <NUM> having a wing portion <NUM>. Wing portion <NUM> has a concave pressure side <NUM> and (obstructed in <FIG>) an opposing convex suction side <NUM>. Wing portion <NUM> can also include a leading edge <NUM> spanning between pressure side <NUM> and suction side <NUM>, and a trailing edge <NUM> opposing leading edge <NUM> and spanning between pressure side <NUM> and suction side <NUM>.

As shown in <FIG>, airfoil <NUM> also includes an outer endwall mount end <NUM> for coupling airfoil <NUM> to outer endwall <NUM>. Nozzle sets 115A-D have identical outer endwall mount ends <NUM>, i.e., outer endwall mount ends <NUM> are identical amongst the plurality of nozzle sets 115A-D. Airfoil <NUM> also includes an inner endwall mount end <NUM> for coupling airfoil <NUM> to inner endwall <NUM>. Nozzle sets 115A-D have identical inner endwall mount ends <NUM>, i.e., inner endwall mount ends <NUM> are identical amongst the plurality of nozzle sets 115A-D. Mount ends <NUM>, <NUM> may include any structure capable of coupling to endwalls <NUM>, <NUM>, e.g., by brazing. In one example, mount ends <NUM>, <NUM> each include a rim member <NUM> that forms respective ends of wing portion <NUM>. Mount ends <NUM>, <NUM> extend a relatively small portion of a radial height of airfoil <NUM>, e.g., <NUM>-<NUM>% each.

As previously noted, and as shown best in <FIG>, nozzle <NUM> also includes outer endwall <NUM> and inner endwall <NUM> connected with airfoil <NUM> along suction side <NUM>, pressure side <NUM>, leading edge <NUM>, and trailing edge <NUM>. Outer endwalls <NUM> are configured to align on the radially outer side of nozzle set <NUM> (<FIG>) and to couple respective nozzle(s) <NUM> to casing <NUM> (<FIG>) of turbine <NUM> (<FIG>). As shown in <FIG>, outer endwall <NUM> includes a joint opening <NUM> configured to receive outer endwall mount end <NUM> of airfoil <NUM>. Inner endwalls <NUM> are configured to align on the radially inner side of nozzle set <NUM> (<FIG>). As shown in <FIG>, inner endwall <NUM> includes a joint opening <NUM> configured to receive inner endwall mount end <NUM> of airfoil <NUM>. Each joint opening <NUM>, <NUM> is configured to receive the respective mount end <NUM>, <NUM> by being sized and shaped to allow receipt of mount ends <NUM>, <NUM> therein in a manner allowing joining, e.g., by welding or brazing.

In various embodiments, shown in <FIG>, each nozzle <NUM> includes a fillet <NUM>, <NUM> connecting airfoil <NUM> and each respective endwall <NUM>, <NUM>. That is, outer endwall <NUM> is mounted to outer endwall mount end <NUM> (<FIG>) of airfoil <NUM> by a fillet <NUM>, and inner endwall <NUM> is mounted to inner endwall mount end <NUM> (<FIG>) of airfoil <NUM> by a fillet <NUM>. Fillets <NUM>, <NUM> can include a weld or braze fillet creating a joint, which may be formed via conventional metal-inert gas (MIG) welding, tungsten-inert gas (TIG) welding, brazing, etc. As shown best in <FIG>, fillets <NUM>, <NUM> can overlap a portion of airfoil <NUM>. The extent of overlap can vary from nozzle to nozzle, stage to stage, and/or turbine to turbine.

<FIG> shows a schematic cross-sectional view of two different sets of nozzles, e.g., 115A, 115B, superimposed to illustrate different pairwise throat areas, according to embodiments of the invention.

The cross-sectional view of <FIG> may be taken, for example, along line <NUM>-<NUM> in <FIG>, but may be at any spanwise cross-sectional location along wing portions <NUM> excepting mount ends <NUM>, <NUM>. A pairwise throat area is created by outer endwall <NUM>, inner endwall <NUM> and wing portions <NUM> of adjacent airfoils <NUM> of the annulus of nozzles <NUM>. Changing the pairwise throat area can change the overall throat area of the annulus of a particular nozzle set <NUM>.

In <FIG>, different pairwise throat areas are illustrated by different throat widths TW1, TW2 at a particular cross-section of wing portions <NUM> of airfoils 128A, 128B. Throat width may be defined as a minimum distance between adjacent airfoils 128A, 128B. In the example shown, throat width is illustrated between a trailing edge <NUM> of one airfoil and a closest point of convex suction side <NUM> of the other airfoil. However, throat width is not necessarily identified at those particular points in all instances. In any event, a throat width TW1 of wing portions 130A of airfoils 128A in nozzle set 115A is different than a throat width TW2 of wing portions 130B of airfoils 128B in nozzle set 115B (i.e., TW1≠TW2). The difference in throat width can occur at any number of spanwise cross-sectional locations of wing portions <NUM>. In this manner, across a radial length of wing portions 130A-D of airfoils 128A-D, the wing portion can be shaped and sized to create a wide variation in pairwise throat area for adjacent airfoils 128A-D in a particular nozzle set 115A-D.

Of note, because mount ends <NUM>, <NUM> of each airfoil <NUM> are identical regardless of airfoil <NUM> in which employed, any airfoil 128A-D that provides a different wing portion <NUM> with a different pairwise throat area can be mounted to outer and inner endwall <NUM>, <NUM>. In this manner, only airfoils <NUM> need to be changed to adjust a pairwise throat area of a nozzle set <NUM>, i.e., for a stage of turbine <NUM> (<FIG>). Any number of nozzle sets <NUM> can be created as part of system <NUM> to allow for replacement of an airfoil 128A having a first pairwise throat area with another airfoil, e.g., 128B, C or D, having a second, different pairwise throat area.

With further reference to <FIG>, while different pairwise throat areas are provided by wing portions 130A-B of different airfoils 128A-B, each of the plurality of wing shapes that are different amongst the plurality of nozzle sets 115A-D (<FIG>) may have a similar, or re-optimized, radius of curvature distribution at each of a plurality of spanwise cross-sectional locations. That is, while a throat width changes between adjacent airfoils, each airfoil <NUM> can have a similar radius of curvature distribution if that is desirable for the application, e.g., within <NUM>% of the previous radius of curvature. In this manner, the aerodynamic performance of a nozzle set 115A-D (<FIG>) created by a wing shape for a wing portion <NUM> is maintained despite the changing of airfoils <NUM>.

Referring to <FIG> and <FIG>, a method according to embodiments of the invention will be described. As shown in <FIG>, for each nozzle in first nozzle set 115A including a plurality of nozzles <NUM> that collectively form a circumferential ring or annulus, i.e., of a stage L0-L3 (<FIG>), inner endwall mount end <NUM> of first wing portion 130A of first airfoil 128A is removed from inner endwall <NUM> of nozzle <NUM>. Similarly, outer endwall mount end <NUM> of first wing portion 130A of first airfoil 128A is removed from outer endwall <NUM> of nozzle <NUM>. As noted, each first wing portion 130A has a first wing shape providing a first pairwise throat area with wing portions of adjacent first airfoils 130A of first nozzle set 115A. Endwalls <NUM>, <NUM> may be removed using any now known or later developed technique, e.g., heating to melt fillets <NUM>, <NUM>, cutting and then removing remnants of airfoil 130A, etc..

Inner endwall mount ends <NUM> of first wing portion 130A and a selected second wing portion 130B, C or D are identical, and outer endwall mount ends <NUM> of first wing portion 130A and second wing portion 130B, C or D are identical. As a result, any of airfoils 128A-D can be readily substituted for one another to change a pairwise throat area of the set of nozzles.

Embodiments of the method continue, as shown in <FIG>, with coupling inner endwall mount end <NUM> of a replacement, or second, airfoil 128B, C or D (of a second nozzle set 115B, C or D) to inner endwall <NUM>. The method further includes coupling outer endwall mount end <NUM> of the respective second airfoil 128B, C, or D to outer endwall <NUM>. This process is repeated for each nozzle in a nozzle set <NUM>. As explained relative to <FIG>, each second wing portion <NUM> (e.g., 130B) has a second wing shape providing a second pairwise throat area with adjacent wing portions 130B of adjacent second airfoils 128B in second nozzle set 115B. The second pairwise throat area of second nozzle set 115B is different from the first pairwise throat area of first nozzle set 115A. Likewise, the pairwise throat area of third nozzle set 115C and the pairwise throat area of fourth nozzle set 115D are each different from the first pairwise throat area of first nozzle set 115A and the second pairwise throat area of second nozzle set 115B.

The coupling of outer endwall <NUM> to outer endwall mount end <NUM> of second wing portion 130B may include, for example, brazing to create first fillet <NUM> (<FIG>). Similarly, coupling inner endwall <NUM> to inner endwall mount end <NUM> of second wing portion 130B may include, for example, brazing to create a second fillet <NUM> (<FIG>). As noted, first and second wing portions 130A, 130B may have a similar radius of curvature distribution at each of a plurality of spanwise cross-sectional locations between inner endwall mount end <NUM> and outer endwall mount end <NUM>, e.g., within <NUM>% of the previous radius of curvature.

Referring to <FIG>, alternative embodiments of nozzle <NUM> are illustrated. In various embodiments, nozzle <NUM> may include one or more cooling passages <NUM> to cool, among other things, mount ends <NUM>, <NUM> of airfoil <NUM>. More particularly, at least one of outer endwall mount end <NUM> and inner endwall mount end <NUM> may further include a cooling passage <NUM>. <FIG> shows a schematic cross-sectional view of a rim member <NUM> of mount end <NUM> or <NUM> within respective inner or outer endwalls <NUM> or <NUM>. As shown, regardless of mount end <NUM>, <NUM>, cooling passage <NUM> may be positioned adjacent at least a portion of respective first or second fillet <NUM>, <NUM>. In this manner, a coolant passing through cooling passage <NUM> can cool rim member <NUM>, fillet <NUM> or <NUM>, and/or inner or outer endwall <NUM>, <NUM>.

Cooling passage <NUM> may extend through nozzle <NUM> in a number of different ways to deliver coolant, where desired. <FIG> shows a rear perspective view of airfoil <NUM> with inner and outer endwalls removed. Here, cooling passage <NUM> passes lengthwise through rim members <NUM> of outer and inner endwall mount ends <NUM>, <NUM>, e.g., along pressure side <NUM> of wing portion <NUM> of airfoil <NUM>. (<FIG> also shows view line <NUM>-<NUM> of <FIG>. ) Coolant may exit cooling passage(s) <NUM> through outlets <NUM> in rim members <NUM> near trailing edge <NUM> of wing portion <NUM>. In other embodiments, cooling passage <NUM> may be provided in only one of rim members <NUM>, i.e., rim member <NUM> of either mount end <NUM> or <NUM>, but not both.

<FIG> shows a rear perspective view of airfoil <NUM> according to other embodiments of the invention. <FIG> shows a nozzle <NUM> with inner and outer endwalls <NUM>, <NUM> (<FIG>) removed. Here, cooling passage <NUM> passes lengthwise through part of one rim member <NUM> (e.g., of outer endwall mount end <NUM>), and along part of pressure side <NUM> of wing portion <NUM> of airfoil <NUM>. Nozzle <NUM> also includes a trailing edge cooling passage <NUM> in wing portion <NUM> and in fluid communication with cooling passage <NUM>. Trailing edge cooling passage <NUM> may include a plurality of passages <NUM> exiting trailing edge <NUM> of wing portion <NUM> of airfoil <NUM>. Any number of passages <NUM> may be used. Coolant can thus enter cooling passage <NUM> near leading edge <NUM>, pass along pressure side <NUM> in rim member <NUM>, and then exit through passages <NUM> in trailing edge <NUM>. In either of the <FIG> and <FIG> embodiments, cooling passage <NUM> may include a radially facing inlet(s) <NUM> through which coolant can enter cooling passage(s) <NUM>. Inlet(s) <NUM> may be adjacent a leading edge <NUM> of wing portion <NUM> of airfoil <NUM>. In this manner, coolant can enter the typically hotter region(s) of airfoil <NUM> and then pass towards relatively cooler regions of the airfoil to cool other parts.

<FIG> shows a schematic cross-sectional view of a cooling passage <NUM>, according to other embodiments of the invention.

In <FIG>, cooling passage <NUM> is in a rim member <NUM> of outer endwall mount end <NUM> within outer endwall <NUM>. (Although not shown, it will be recognized that this cooling passage <NUM> arrangement can also be applied to rim member <NUM> of inner endwall mount end <NUM> within inner endwall <NUM>. ) A nozzle cap <NUM> is shown sealing a radial outer end <NUM> of a radially extending, interior cooling chamber <NUM> within wing portion <NUM> of airfoil <NUM>. Here, cooling passage <NUM> includes a radially facing inlet <NUM> (facing radially outward in this example), and an outlet <NUM> facing into interior cooling chamber <NUM> of airfoil <NUM>. Coolant can pass from an area <NUM> radially outside of outer endwall mount end <NUM> through rim member <NUM> thereof to cool rim member <NUM>, fillet <NUM>, and outer endwall <NUM>. Subsequently, coolant enters interior cooling chamber <NUM> where it can be used to provide further cooling radially inward of rim member <NUM>.

For example, although not necessary in all cases, an impingement cooling member <NUM> may be positioned within interior cooling chamber <NUM>. Impingement cooling member <NUM> may include any now known or later developed impingement cooling structure such as a sleeve <NUM> with a plurality of openings <NUM> therein that allow coolant within interior cooling chamber <NUM> to exit and impinge on an inner surface <NUM> of wing portion <NUM> of airfoil <NUM>. Coolant from interior cooling chamber <NUM>, including coolant from cooling passage <NUM>, can impinge inner surface <NUM> to cool wing portion <NUM>.

<FIG> shows a perspective, partial cross-sectional view of outer endwall mount end <NUM> of nozzle <NUM> including a plurality of cooling passages <NUM> from <FIG>. Cooling passages <NUM> can be arranged within rim member <NUM> of outer endwall mount end <NUM> in any manner, e.g., spaced evenly, located where hot spots are expected, etc. As illustrated, any number of cooling passages <NUM>, as shown in <FIG>, may be positioned about rim member <NUM> of outer endwall mount end <NUM>.

Coolant can be provided to cooling passage(s) <NUM> from any now known or later developed coolant source. For example, for a radially outer end, coolant can be provided from area <NUM> (<FIG> and <FIG>), which is within casing <NUM> (<FIG>) and is filled with compressed air, e.g., from compressor <NUM> (<FIG>). In another example, for a radially inner end, coolant can be provided from an internal cooling chamber <NUM> of wing portion <NUM> of airfoil <NUM>, or an internal wheel space (not shown) between stages of turbine <NUM>.

While particular embodiments of cooling passage(s) <NUM> have been provided herein, other embodiments can include any variety of cooling passages developed for nozzles. It will be recognized that with replacement of airfoils <NUM>, cooling passages <NUM> can be provided that are the same as in the original airfoils, or the cooling passages <NUM> can be adjusted from the original airfoil to improve, for example, cooling, turbine performance and nozzle longevity.

Embodiments of the invention provide a system that allows for adjusting of a throat area (i.e., pairwise throat area and overall throat area) of a set of nozzles of a turbine without the expense of replacing the entirety of each nozzle in the set. In this manner, aerodynamic performance of a turbine <NUM> can be maintained or improved despite the aging of the turbine. The different airfoils have the same mount ends <NUM>, <NUM> that allow coupling to used endwalls <NUM>, <NUM>, thus eliminating the need to replace the endwalls. Thus, system <NUM> allows airfoils <NUM> to be made separately, e.g., by casting or additive manufacture, from endwalls <NUM>, <NUM>, which is easier and less expensive than forming a one-piece nozzle and replacing each nozzle in a set. Cooling passages <NUM> can be provided in the replacement airfoils to maintain or improve cooling.

Accordingly, a value modified by a term or terms, such as "about," "approximately" and "substantially," are not to be limited to the precise value specified. Here and throughout the specification and claims, range limitations may be combined and/or interchanged; such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. "Approximately," as applied to a particular value of a range, applies to both end values and, unless otherwise dependent on the precision of the instrument measuring the value, may indicate +/- <NUM>% of the stated value(s).

Claim 1:
A turbine nozzle assembly system (<NUM>), comprising:
a plurality of nozzle sets (<NUM>, 115A-D)), each nozzle set (<NUM>, 115A-D) including a plurality of nozzles (<NUM>) that collectively form an annulus of a particular nozzle stage (L0-L3) of a turbine (<NUM>), characterised by
each nozzle (<NUM>) of a respective nozzle set (<NUM>) including
an airfoil (<NUM>, 128A-D) having:
an inner endwall mount end (<NUM>),
an outer endwall mount end (<NUM>), and
a wing portion (<NUM>, 130A-B) between the inner endwall mount end (<NUM>) and the outer endwall mount end (<NUM>);
an inner endwall (<NUM>) including a first joint opening (<NUM>, <NUM>) configured to receive the inner endwall mount end (<NUM>) of the airfoil (<NUM>, 128A-D);
an outer endwall (<NUM>) including a second joint opening (<NUM>, <NUM>) configured to receive the outer endwall mount end (<NUM>) of the airfoil (<NUM>, 128A-D);
wherein the inner endwall mount ends (<NUM>) amongst the plurality of nozzle sets (<NUM>, 115A-D) are identical to each other,
wherein the outer endwall mount ends (<NUM>) amongst the plurality of nozzle sets (<NUM>, 115A-D) are identical to each other,
wherein the wing portions (<NUM>, 130A-B) of a respective nozzle set (<NUM>, 115A-D) of the plurality of nozzle sets (<NUM>, 115A-D) have a wing shape that is identical for each airfoil (<NUM>, 128A-D) within the respective nozzle set (<NUM>) but different than the wing shapes of the wing portions of the other nozzle sets (<NUM>, 115A-D) of the plurality of nozzle sets (<NUM>, 115A-D),
wherein for each nozzle set (<NUM>, 115A-D): adjacent nozzles (<NUM>) of the plurality of nozzles (<NUM>) define a pairwise throat area (<NUM>) created by the respective inner endwalls (<NUM>), the respective outer endwalls (<NUM>) and the respective wing portions (<NUM>) of the adjacent airfoils (<NUM>, 128A-B) in the annulus of the particular nozzle stage (L0-L3) of the turbine (<NUM>), and
wherein the pairwise throat area (<NUM>) of each nozzle set (<NUM>) of the plurality of nozzle sets (<NUM>, 115A-D) is identical for each pairwise throat area (<NUM>) within the respective nozzle set (<NUM>) but different than the pairwise throat areas (<NUM>) of the other nozzle sets (<NUM>) of the plurality of nozzle sets (<NUM>, 115A-D).