Various embodiments of the invention include turbine nozzles and systems employing such nozzles. Various particular embodiments include a turbine nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including a non-axisymmetric contour proximate a junction between the endwall and the leading edge of the airfoil.

FIELD OF THE INVENTION

The subject matter disclosed herein relates to turbomachines. More particularly, the subject matter disclosed herein relates to components within turbomachines such as gas and/or steam turbines.

BACKGROUND OF THE INVENTION

Some aircraft and/or power plant systems, for example certain jet aircraft, nuclear, simple cycle and combined cycle power plant systems, employ turbines (also referred to as turbomachines) in their design and operation. Some of these turbines employ airfoils (e.g., turbine blades, blades, airfoils, etc.) which during operation are exposed to fluid flows. These airfoils are configured to aerodynamically interact with the fluid flows and generate energy (e.g., creating thrust, turning kinetic energy to mechanical energy, thermal energy to mechanical energy, etc.) from these fluid flows as part of power generation. As a result of this interaction and conversion, the aerodynamic characteristics and losses of these airfoils have an impact on system and turbine operation, performance, thrust, efficiency, and power.

BRIEF DESCRIPTION OF THE INVENTION

Various embodiments of the invention include turbine nozzles and systems employing such nozzles. Various particular embodiments include a turbine nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including a non-axisymmetric contour proximate a junction between the endwall and the leading edge of the airfoil.

A first aspect of the invention includes a turbine nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including a non-axisymmetric contour proximate a junction between the endwall and the leading edge of the airfoil.

A second aspect of the invention includes a static nozzle section having: a set of static nozzles, the set of static nozzles including at least one nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including a non-axisymmetric contour proximate a junction between the endwall and the leading edge of the airfoil, wherein at least one of the suction side or the pressure side of the airfoil includes a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in TABLE I, wherein the coordinate values are non-dimensional values of from 0 to 1 convertible to distances by multiplying the values by a trailing edge height expressed in units of distance, and wherein X and Y values connected by smooth continuing arcs define airfoil profile sections at each distance Z along the airfoil, the profile sections at the Z distances being joined smoothly with one another to form the airfoil profile, wherein the Cartesian coordinate values have an origin at a root of the leading edge of the airfoil.

A third aspect of the invention includes a turbine nozzle comprising: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, wherein at least one of the pressure side or the suction side of the airfoil includes a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in TABLE I, wherein the coordinate values are non-dimensional values of from 0 to 1 convertible to distances by multiplying the values by a trailing edge height expressed in units of distance, and wherein X and Y values connected by smooth continuing arcs define airfoil profile sections at each distance Z along the airfoil, the profile sections at the Z distances being joined smoothly with one another to form the airfoil profile, wherein the Cartesian coordinate values have an origin at a root of the leading edge of the airfoil.

It is noted that the drawings of the invention are not necessarily to scale. 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. It is understood that elements similarly numbered between the FIGURES may be substantially similar as described with reference to one another. Further, in embodiments shown and described with reference toFIGS. 1-6, like numbering may represent like elements. Redundant explanation of these elements has been omitted for clarity. Finally, it is understood that the components ofFIGS. 1-6and their accompanying descriptions may be applied to any embodiment described herein.

DETAILED DESCRIPTION OF THE INVENTION

As noted herein, various aspects of the invention are directed toward turbine nozzles. Particular aspects of the invention include turbine nozzles having at least one endwall with a non-axisymmetric contour.

In contrast to conventional turbine nozzles, aspects of the invention include a turbine nozzle (e.g., a static nozzle for directing a working fluid such as gas or steam) having a non-axisymmetric contour at the leading edge of its endwall. This non-axisymmetric contour can provide for enhanced performance, efficiency and/or durability of the nozzle (and associated turbine stages and turbine machines) when compared with conventional nozzles.

As used herein, the terms “axial” and/or “axially” refer to the relative position/direction of objects along axis A, which is substantially parallel to the axis of rotation of the turbomachine (in particular, the rotor section). As further used herein, the terms “radial” and/or “radially” refer to the relative position/direction of objects along axis (r), which is substantially perpendicular with axis A and intersects axis A at only one location. Additionally, the terms “circumferential” and/or “circumferentially” refer to the relative position/direction of objects along a circumference which surrounds axis A but does not intersect the axis A at any location. Further, the terms leading edge/pressure side refer to components and/or surfaces which are oriented predominately upstream relative to the fluid flow of the system, and the terms trailing edge/suction side refer to components and/or surfaces which are oriented predominately downstream relative to the fluid flow of the system.

Referring to the drawings,FIG. 1shows a perspective partial cut-away illustration of a turbine10(e.g., a gas or steam turbine and/or an aviation jet engine) according to various embodiments of the invention. Turbine10includes a rotor12that includes a rotating shaft14and a plurality of axially spaced rotor wheels18. A plurality of rotating buckets20are mechanically coupled to each rotor wheel18. More specifically, buckets20are arranged in rows that extend circumferentially around each rotor wheel18. A static nozzle section21is shown including a plurality of stationary nozzles22that circumferentially around shaft14, and the nozzles22are axially positioned between adjacent rows of buckets20. Stationary nozzles22cooperate with buckets20to form a stage of the turbine10, and to define a portion of a flow path through turbine10. As shown, the static nozzle section21at least partially surrounds the rotor12(shown in this cut-away view). It is understood that the turbine10shown is a dual-flow turbine10that includes an axially centered inlet mouth which feeds two sets of turbine stages. It is understood that various teachings can be applied to axial turbines, e.g., axial inlet gas turbines that inlet a combustion gas from a first axial end and outlet that combustion gas to a second axial end after the gas has performed mechanical work on the turbine.

Returning toFIG. 1, in operation, gas24enters an inlet26of turbine10and is channeled through stationary nozzles22. Nozzles22direct gas24against blades20. Gas24passes through the remaining stages imparting a force on buckets20causing shaft14to rotate. At least one end of turbine10may extend axially away from rotating shaft12and may be attached to a load or machinery (not shown) such as, but not limited to, a generator, and/or another turbine.

In one embodiment, turbine10may include five stages. The five stages are referred to as L0, L1, L2, L3and L4. Stage L4is the first stage and is the smallest (in a radial direction) of the five stages. Stage L3is the second stage and is the next stage in an axial direction. Stage L2is the third stage and is shown in the middle of the five stages. Stage L1is the fourth and next-to-last stage. Stage L0is the last stage and is the largest (in a radial direction). It is to be understood that five stages are shown as one example only, and each turbine may have more or less than five stages. Also, as will be described herein, the teachings of the invention do not require a multiple stage turbine. In another embodiment, turbine10may comprise an aircraft engine used to produce thrust.

Turning toFIG. 2, a schematic three-dimensional depiction of a turbine nozzle (or simply, nozzle)200is shown according to various embodiments. The nozzle200is a stationary nozzle which forms an annulus of stationary nozzles in a stage of a turbine (e.g., turbine10). That is, during operation of a turbine (e.g., turbine10), the nozzle200will remain stationary in order to direct the flow of working fluid (e.g., gas or steam) to one or more movable buckets (e.g., buckets20), causing those movable buckets to initiate rotation of a rotor shaft (e.g., shaft14). It is understood that nozzle200is configured to couple (mechanically couple via fasteners, welds, slot/grooves, etc.) with a plurality of similar or distinct nozzles (e.g., nozzles200or other nozzles) to form an annulus of nozzles in a stage of the turbine.

Returning toFIG. 2, the turbine nozzle200can include an airfoil202having a suction side204, and a pressure side206(obstructed in this view) opposing the suction side204. The nozzle200can also include a leading edge208spanning between the pressure side206and the suction side204, and a trailing edge210opposing the leading edge208and spanning between the pressure side206and the suction side204.

As shown, the nozzle200can also include at least one endwall212(two shown) connected with the airfoil202. The nozzle200can be connected with the airfoil202along the suction side204, pressure side206, trailing edge210and the leading edge208. In various embodiments, the nozzle includes a fillet214connecting the nozzle210and each endwall212. The fillet214can include a weld or braze fillet, which may be formed via conventional MIG welding, TIG welding, brazing, etc.

As described herein, and in contrast to conventional turbine nozzles, the turbine nozzle200can include at least one endwall212with a non-axisymmetric contour218proximate a junction220between the endwall212and the leading edge208of the airfoil202. That is, the nozzle200includes an endwall212with a contour218proximate the junction220between the endwall212and the leading edge208of the airfoil202that improves the flow area around the airfoil202when compared with conventional nozzles.

In various embodiments, the contour218allows for more efficient fluid flow across the airfoil202than conventional nozzles200, allowing for fewer heat load-related failures, and improving the efficiency of fluid flow within a turbine utilizing such a nozzle200.

With reference toFIG. 1, in various embodiments, the nozzle200can include a first stage nozzle (L4) or second stage nozzle (L3). In particular embodiments, the nozzle200is a second stage nozzle (L3), and the improved flow profile across the airfoil200and endwall212interface allows that second stage nozzle (L3) to withstand the high-temperature gas entering the turbine10at that first stage. In various embodiments, the turbine10can include a set of nozzles200in only the second stage (L3) of the turbine10, or in only the first stage (L4) and the second stage (L3) of the turbine10.

In various embodiments, at least one of the endwalls212including the contour218can include an inner endwall, e.g., a radially inner endwall configured to align on the radially inner side of a the static nozzle section. In other embodiments, at least one of the endwalls212including the contour218can include an outer endwall, e.g., a radially outer endwall configured to align on the radially outer side of the static nozzle section. In some cases, both endwalls212include the contour218, and in other cases, only one of the endwalls212includes the contour218.

According to various embodiments, the non-axisymmetric contour218includes a first surface222along the endwall212on the suction side204of the leading edge208, and a second surface224along the endwall212on the pressure side206of the leading edge208. The second surface224and the first surface222can have distinct slopes, e.g., distinct radial v. circumferential ratios. In various embodiments, the first surface222has a distinct profile from the second surface224. In some cases, the distinct profile includes distinct endwall features (e.g., bump(s), trough(s), etc.) in the first surface222as compared with the second surface224(having its own endwall features (e.g., bump(s), trough(s), etc.). In some cases, the second surface224has a substantially flat, or unsloped gradient, and the first surface222has a gradient distinct from the gradient of the second surface224, e.g., a gradient that is positive or negative, but not equal to zero.

According to various particular embodiments, the first surface222has a first length L1measured from a junction228of the suction side204and the leading edge208of the airfoil202along the endwall212to an outer edge230of the endwall212. In these embodiments, the second surface224has a second length L2measured from a junction232of the pressure side206and the leading edge208of the airfoil202along the endwall212to an inner edge234of the endwall212. In various embodiments, the first length L1is distinct from the second length L2, and in particular embodiments, the second length L2is greater than the first length L1.

Turning toFIG. 3, a radial contour map of a portion of the endwall212(excluding the nozzle200) is shown according to various embodiments. As can be seen in this example radial contour map, the contour218(shown in three parts as a leading edge pressure side feature218A, a leading edge suction side feature218B and a mid-passage feature218C) can aid in improving aerodynamic efficiency proximate the leading edge208of the nozzle200. The radial contour map also shows a passage trough250on the pressure side206of the nozzle200.

With reference toFIG. 2andFIG. 3, according to various particular embodiments, the non-axisymmetric contour218includes at least one bump (thickened area)260and a passage trough250on the pressure side206of the nozzle200. In particular embodiments, the leading edge pressure side feature218A can include at least one bump (thickened area)260, and the leading edge suction side feature218B and mid passage feature218C can include at least one bump (thickened area)260. It is understood that according to various embodiments, the bump (thickened area)260along the leading edge pressure side218A can be thinner than the bump260along the leading edge suction side218B (extending a lesser distance toward the opposite endwall212from the second surface224than the bump260extends toward the opposite endwall from the first surface222).

In various particular embodiments, each bump (thickened area)260can extend across approximately at least 10 percent of the axial length LA(along axis A) of the endwall212. In some particular cases, each thickened area260can extend across approximately 20-30 percent of the axial length LAof the endwall212.

In various particular embodiments, at least one bump260(within218A) can have an apex at approximately 0% axial chord upstream of the leading edge208(+/−5%). In these cases, the bump260can span approximately 10-20% pitch as measured from the pressure side206.

In other particular embodiments, at least one bump260(e.g., within218B) has an apex at approximately 5% of the axial cord length upstream of the leading edge208(+/−5%). In these cases, the bump260can span approximately 0-10% pitch as measured from the suction side204.

In another embodiment, at least one bump260(within218C) has an apex at approximately 40-60% pitch as measured from pressure side206. This bump260may span approximately 10-20% pitch.

In various embodiments, the passage trough250includes a depression having an apex at approximately 60% (+/−10%) of the chord length of the pressure side206(+/−10%). In this case, the depression in the pressure trough250can span approximately 40-60% pitch as measured from the pressure side206.

It is understood that in various embodiments, other apex locations and pitches are possible, and those values given herein are merely illustrative of several of the many possible embodiments in accordance with the disclosure.

With reference toFIG. 4(and continuing reference toFIGS. 2-3), a plurality of points270-278along span S, including root215and tip217, can correspond to Z coordinate values of chord lines, and a cross section of airfoil202at each point can be described by a respective set of X and Y coordinates. For example, 100 points can be listed for each cross section270-278, though it should be apparent that more or fewer points can be used for each cross section, and more or fewer cross sections can be used, as may be desired and/or appropriate. The X, Y, and Z coordinate values in TABLE I have been expressed in normalized or non-dimensionalized form in values of from 0 to 1, but it should be apparent that any or all of the coordinate values could instead be expressed in distance units so long as the proportions are maintained. To convert an X, Y or Z value of TABLE I to a respective X, Y or Z coordinate value in units of distance, such as inches or meters, the non-dimensional X, Y or Z value given in TABLE I can be multiplied by a trailing edge height of airfoil202in such units of distance. By connecting the X and Y values with smooth continuing arcs, each profile cross section at each distance Z can be fixed, and the airfoil profiles of the various surface locations between the distances Z can be determined by smoothly connecting adjacent profile sections to one another, thus forming the airfoil profile.

The values in TABLE I are generated and shown to four decimal places for determining the profile of at least one of a suction side or a pressure side of a nominal airfoil202at ambient, non-operating, or non-hot conditions, and do not take any coatings or fillets into account, though embodiments could account for other conditions, coatings, and/or fillets. To allow for typical manufacturing tolerances and/or coating thicknesses, ±values can be added to the values listed in TABLE I, particularly to the X and Y values therein. For example, a tolerance of about 10-20percent of a thickness of the trailing edge in a direction normal to any surface location along the airfoil profile can define an airfoil profile envelope for a nozzle airfoil design at cold or room temperature. In other words, a distance of about 10-20 percent of a thickness of the trailing edge in a direction normal to any surface location along the airfoil profile can define a range of variation between measured points on an actual airfoil surface and ideal positions of those points, particularly at a cold or room temperature, as embodied by the invention. The nozzle airfoil design, as embodied by the invention, is robust to this range of variation without impairment of mechanical and aerodynamic functions. Likewise, the profile and/or design can be scaled up or down, such as geometrically, without impairment of operation, and such scaling can be facilitated by use of normalized coordinate values, i.e. multiplying the normalized values by a scaling factor, or a larger or smaller number of distance units than might have originally been used. For example, the values in TABLE I, particularly the X and Y values, could be multiplied by a scaling factor of 2, 0.5, or any other desired scaling factor. Alternatively, the values could be multiplied by a larger or smaller desired span. As referenced herein, the origin of the X, Y, Z coordinate system is the root of the leading edge (junction232) of the airfoil202.

According to various embodiments, and as a result of endwall contour218, a region of a passage trough250between two airfoils202proximate endwall212can be affected. For example, a bottom edge of a passage trough250between a pair of nozzles200can vary radially, whereas an endwall without a contour would leave a bottom edge of such a throat as at least a straight line, if not a substantially constant radial distance.

Turning toFIG. 5, a schematic view of portions of a multi-shaft combined cycle power plant900is shown. Combined cycle power plant900may include, for example, a gas turbine980operably connected to a generator970. Generator970and gas turbine980may be mechanically coupled by a shaft915, which may transfer energy between a drive shaft (not shown) of gas turbine980and generator970. Also shown inFIG. 5is a heat exchanger986operably connected to gas turbine980and a steam turbine992. Heat exchanger986may be fluidly connected to both gas turbine980and a steam turbine992via conventional conduits (numbering omitted). Gas turbine980and/or steam turbine992may include one or more nozzles200as shown and described with reference toFIG. 2and/or other embodiments described herein. Heat exchanger986may be a conventional heat recovery steam generator (HRSG), such as those used in conventional combined cycle power systems. As is known in the art of power generation, HRSG986may use hot exhaust from gas turbine980, combined with a water supply, to create steam which is fed to steam turbine992. Steam turbine992may optionally be coupled to a second generator system970(via a second shaft915). It is understood that generators970and shafts915may be of any size or type known in the art and may differ depending upon their application or the system to which they are connected. Common numbering of the generators and shafts is for clarity and does not necessarily suggest these generators or shafts are identical. In another embodiment, shown inFIG. 6, a single shaft combined cycle power plant990may include a single generator970coupled to both gas turbine980and steam turbine992via a single shaft915. Steam turbine992and/or gas turbine980may include one or more nozzles200shown and described with reference toFIG. 2and/or other embodiments described herein.

The apparatus and devices of the present disclosure are not limited to any one particular engine, turbine, jet engine, generator, power generation system or other system, and may be used with other aircraft systems, power generation systems and/or systems (e.g., combined cycle, simple cycle, nuclear reactor, etc.). Additionally, the apparatus of the present invention may be used with other systems not described herein that may benefit from the increased efficiency of the apparatus and devices described herein.

In various embodiments, components described as being “coupled” to one another can be joined along one or more interfaces. In some embodiments, these interfaces can include junctions between distinct components, and in other cases, these interfaces can include a solidly and/or integrally formed interconnection. That is, in some cases, components that are “coupled” to one another can be simultaneously formed to define a single continuous member. However, in other embodiments, these coupled components can be formed as separate members and be subsequently joined through known processes (e.g., fastening, ultrasonic welding, bonding).