Patent Description:
Turbomachines, such as gas turbine systems, are utilized to generate power for electric generators. In general, gas turbine systems generate power by passing a fluid (e.g., hot gas) through a turbine component of the gas turbine system. More specifically, inlet air may be drawn into a compressor and may be compressed. Once compressed, the inlet air is mixed with fuel, which may be ignited by a combustor of the gas turbine system to form the operational fluid (e.g., hot gas) of the gas turbine system. The fluid may then flow through a fluid flow path for rotating a plurality of rotating blades and a rotor or shaft of the turbine component for generating the power. The fluid may be directed through the turbine component via the plurality of rotating blades and a plurality of stationary nozzles or vanes positioned between the rotating blades. As the plurality of rotating blades rotate the rotor of the gas turbine system, a generator, which is coupled to the rotor, may generate power from the rotation of the rotor.

To improve operational efficiencies, rotating blades may include tip shrouds on radially outer ends thereof. The tip shrouds interact with an inner surface of a stationary casing to direct the operational fluid. The tip shrouds include a mass of material that presents a number of mechanical integrity issues. One challenge is addressing creep, or the gradual deformation of the tip shroud under the stress of rotational forces exerted on the rotating blades. Creep can create a number of mechanical issues such as deformation of the airfoil or the tip shroud. In addition, excessive bending moments created by a mass imbalance in the tip shroud can exacerbate creep issues. Hence, another challenge is ensuring mass balance in the tip shroud to provide ideal aerodynamic, heat transfer, mechanical and aeromechanic performance.

<CIT> discloses a turbine blade, comprising a root section, an airfoil coupled to the root section, the airfoil including a pressure side and a suction side, and a tip shroud. The tip shroud comprises a body coupled to a radial outer end of the airfoil and including a leading circumferential-facing edge and a trailing circumferential-facing edge, and at least one tip rail extending radially outwardly from the body and extending along a circumferential length of the body. The tip shroud further comprises a first plurality of cooling passages defined in the body and extending circumferentially therein, and at least one first edge wall arrangement along at least one of the leading circumferential-facing edge and the trailing circumferential-facing edge of the body. The first edge wall arrangement includes a first edge wall extending axially and radially outwardly from the body along the at least one of the leading circumferential-facing edge and the trailing circumferential-facing edge of the body, the first edge wall including a first circumferentially facing surface, and an exit surface adjacent the first edge wall. The exit surface having an exit opening defined therein through which at least one of the first plurality of cooling passages exits the body. The exit surface is angled relative to the first circumferentially facing surface of the first edge wall. <CIT> discloses another prior art turbine tip shroud.

All aspects, examples, and features mentioned below can be combined in any technically possible way.

An aspect of the invention is a tip shroud for a turbine blade, according to claim <NUM>.

Another aspect of the disclosure includes any of the preceding aspects, and the first edge wall and the exit surface extend axially between a pair of axially opposing, radially extending walls in the body.

Another aspect of the disclosure includes any of the preceding aspects, and the first edge wall includes at least one opening therethrough.

Another aspect of the disclosure includes any of the preceding aspects, and the exit opening for the at least one of the first plurality of cooling passages is linearly aligned with the at least one opening in the first edge wall.

Another aspect of the disclosure includes any of the preceding aspects, and further includes a curved opening defined in a trailing edge of the body adjacent a trailing edge of the airfoil, wherein the curved opening is not filled by the body of an adjacent tip shroud.

Another aspect of the disclosure includes any of the preceding aspects, and the curved opening defines a plane angled in a range of <NUM>° and <NUM>° relative to a radial direction.

Another aspect of the disclosure includes any of the preceding aspects, and further includes at least one second edge wall arrangement along at least one of the leading circumferential-facing edge and the trailing circumferential-facing edge of the body, each second edge wall arrangement including: a second edge wall extending axially and radially outwardly from the body along the at least one of the leading circumferential-facing edge and the trailing circumferential-facing edge of the body; and an inner wall extending axially and radially outwardly from the body, the inner wall parallel to and circumferentially spaced from the second edge wall to create a radially extending pocket therebetween, the inner wall having an exit opening defined therein for at least one of a second plurality of cooling passages defined in the body and extending circumferentially therein, the exit opening circumferentially facing into the radially extending pocket.

Another aspect of the disclosure includes any of the preceding aspects, and the second edge wall includes at least one opening therethrough.

Another aspect of the disclosure includes any of the preceding aspects, and the at least one tip rail includes a plurality of tip rails, and wherein the at least one first edge wall arrangement is axially positioned between a pair of the plurality of tip rails.

Another aspect of the disclosure includes any of the preceding aspects, and the exit surface is not planar.

Another aspect of the invention provides a turbine blade, according to claim <NUM>, comprising: a root section; an
airfoil coupled to the root section, the airfoil including a pressure side and a suction side; and a tip shroud as described before.

Another aspect of the invention includes a gas turbine, according to claim <NUM>, comprising the turbine blade of
any of the preceding aspects.

Two or more aspects described in this disclosure, including those described in this summary section, may be combined to form implementations not specifically described herein.

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.

As an initial matter, in order to clearly describe the subject matter of the current disclosure, 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 application 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 engine 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 (i.e., the direction from which the flow originates). 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 section of the turbomachine.

It is often required to describe parts that are disposed at differing 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. Finally, the term "circumferential" refers to a direction perpendicular to an axis, i.e., a plane in which movement or position may be around the axis. It will be appreciated that such terms may be applied in relation to the center axis of the turbine.

The terms such as "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.

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 describe component or element may or may not be present, and that the description includes instances where the event occurs or the component is present and instances where it does not 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 to, connected to, or coupled to the other element or layer, or intervening elements or layers may be present.

As indicated above, the disclosure provides a tip shroud for a turbine blade of a gas turbine system. The tip shroud may include a body coupled to a radial outer end of an airfoil of the turbine blade. The airfoil includes a pressure side and a suction side. The body of the tip shroud includes a leading circumferential-facing edge and a trailing circumferential-facing edge. The tip shroud may include at least one tip rail extending radially outwardly from the body and extending generally along a circumferential length of the body. Cooling passages are defined in the body and extend circumferentially in the body to cool an area near a first edge wall. The tip shroud also includes at least one first edge wall arrangement along at least one of the leading circumferential-facing edge and the trailing circumferential-facing edge of the body. The first edge wall arrangement(s) may include a first edge wall extending axially and radially outwardly from the body of the tip shroud along the leading and/or trailing circumferential-facing edges of the body. The first edge wall includes a circumferentially facing surface.

Where the tip shroud would otherwise include an inner wall spaced from the first edge wall and through which the cooling passages exit the body, the tip shroud in embodiments of the disclosure includes an exit surface adjacent the first edge wall. The exit surface includes an exit opening defined therein through which at least one of the cooling passages exits the body. The exit surface is angled relative to the circumferentially facing surface of the first edge wall in a range of <NUM>° to <NUM>°. Hence, the exit surface allows less material to be used at any selected circumferential-facing edge of the tip shroud to enhance mass balance, while also providing adequate cooling.

Referring to the drawings, <FIG> is a schematic view of an illustrative turbomachine <NUM> in the form of a gas turbine (GT) system <NUM> (hereinafter "GT system <NUM>"). 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 compressor/turbine rotor shaft <NUM> (hereinafter referred to as "rotor shaft <NUM>"). In one non-limiting embodiment, GT system <NUM> may be a GT26 engine, commercially available from General Electric Company, Greenville, S. The present disclosure is not limited to any one particular GT system and may be implanted in connection with other engines including, for example, other HA, F, B, LM, GT, TM and E-class engine models of General Electric Company, and engine models of other companies. Further, the teachings of the disclosure 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 (that is, downstream from Stage L0). Stage L2 is the third stage and is the next stage in an axial direction (that is, downstream from Stage L1). Stage L3 is the fourth, last stage (downstream from Stage L2) 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 may have more or less than four stages.

In turbine <NUM>, a set of stationary vanes or nozzles <NUM> cooperate with a set of rotating turbine blades <NUM> to form each stage L0-L3 of turbine <NUM> and to define a portion of a flow path through turbine <NUM>. Rotating turbine blades <NUM> in each set are coupled to a respective rotor wheel <NUM> that couples them circumferentially to rotor shaft <NUM>. That is, set of rotating turbine blades <NUM> is mechanically coupled in a circumferentially spaced manner to each rotor wheel <NUM>. A static blade section <NUM> includes stationary nozzles <NUM> circumferentially spaced around rotor shaft <NUM>. Each nozzle <NUM> may include at least one endwall (or platform) <NUM>, <NUM> connected with an airfoil <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>.

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 is 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 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 an enlarged perspective view of an illustrative turbine blade <NUM> in detail. For purposes of description, a legend may be provided in the drawings in which the X-axis extends generally axially (same as arrow A), the Y-axis extends generally perpendicular to axis A of rotor shaft <NUM> (<FIG>) (indicating a circumferential plane or direction), and the Z-axis extends radially, relative to an axis A of rotor shaft <NUM> (<FIG>). The Z-axis is perpendicular to the X-axis and the Y-axis. Blade <NUM> is a rotatable (dynamic) blade, which is part of set of turbine blades <NUM> circumferentially dispersed about rotor shaft <NUM> (<FIG>) in a stage of a turbine (e.g., turbine <NUM>).

During operation of turbine <NUM>, as a working fluid (e.g., gas in GT system <NUM>, or steam in a steam turbine) is directed across the blade's airfoil, blade <NUM> will initiate rotation of a rotor shaft (e.g., rotor shaft <NUM>) and rotate about axis A defined by rotor shaft <NUM>. It is understood that blade <NUM> is configured to couple (mechanically via fasteners, welds, slot/grooves, etc.) with a plurality of similar or distinct blades (e.g., blades <NUM> or other blades) to form set of turbine blades <NUM> (<FIG>) in a stage of turbine <NUM> (<FIG>). Referring to <FIG>, blade <NUM> can be located in any stage (L0-L3).

Returning to <FIG>, blade <NUM> can include an airfoil <NUM> having a pressure side <NUM> (obstructed in this view) and a suction side <NUM> opposing pressure side <NUM>. Blade <NUM> can also include a leading edge <NUM> spanning between pressure side <NUM> and suction side <NUM>, and a trailing edge <NUM> opposing a leading edge <NUM> and spanning between pressure side <NUM> and suction side <NUM>. As noted, pressure side <NUM> of airfoil <NUM> generally faces upstream, and suction side <NUM> generally faces downstream.

As shown, blade <NUM> can also include a root section <NUM> connected with airfoil <NUM> and a turbine blade tip shroud <NUM> (hereinafter "tip shroud <NUM>") on a radial outer end <NUM> of airfoil <NUM>. Root section <NUM> can be connected with airfoil <NUM> along pressure side <NUM>, suction side <NUM>, leading edge <NUM> and trailing edge <NUM>. In various embodiments, blade <NUM> can include a fillet <NUM> proximate a radially inner end <NUM> of airfoil <NUM>, fillet <NUM> connecting airfoil <NUM> and root section <NUM>. Fillet <NUM> can include a weld or braze fillet, which may be formed via conventional MIG welding, TIG welding, brazing, etc. Root section <NUM> is illustrated in <FIG> as including a dovetail <NUM>, but root section <NUM> can have any suitable configuration to connect to rotor shaft <NUM>. Specifically, root section <NUM> is configured to fit into a mating slot (e.g., dovetail slot) in the turbine rotor shaft (e.g., a rotor wheel of rotor shaft <NUM>) and to mate with adjacent components of other blades <NUM>. Root section <NUM> is intended to be located radially inboard of airfoil <NUM> and be formed in any complementary configuration to the rotor shaft.

Tip shroud <NUM> can be connected with airfoil <NUM> along pressure side <NUM>, suction side <NUM>, leading edge <NUM> and trailing edge <NUM>. In various embodiments, blade <NUM> can include a fillet <NUM> proximate radially outer end <NUM> of airfoil <NUM>. Fillet <NUM> connects airfoil <NUM> and tip shroud <NUM>. Fillet <NUM> can include a weld or braze fillet, which may be formed via conventional MIG welding, TIG welding, brazing, etc. Tip shroud <NUM> is configured to interact with an inner surface of casing <NUM> (<FIG>) and/or a casing shroud therein (not shown).

<FIG> shows a radially inward, perspective view of a pair of adjacent turbine blades <NUM> and, in particular, adjacent tip shrouds <NUM>; and <FIG> shows a perspective view of a single tip shroud <NUM>. As will be apparent from the description that follows, embodiments of the disclosure may include repeating structure. For example, where tip shroud <NUM> includes more than two tip rails <NUM>, a first edge wall arrangement <NUM> according to embodiments of the disclosure, as will be described herein, may be repeated between the different pairs of tip rails <NUM>. To differentiate between repeating structure, where necessary, numeric references may be accompanied by a letter reference, e.g., A, B, C, etc. Where a reference letter is omitted from repeating structures, depending on the apparent context, a single instant of the structure is being referenced alone or a number of the structures are being collectively referenced.

With reference to <FIG>, tip shroud <NUM> for turbine blade <NUM> may include a body <NUM> coupled to radial outer end <NUM> of airfoil <NUM> of turbine blade <NUM>. As noted, airfoil <NUM> includes pressure side <NUM> and suction side <NUM>. In <FIG> and <FIG>, an airfoil cooling chamber <NUM> has the general shape of airfoil <NUM> and is mostly radially inward of tip shroud <NUM>. It is understood that one or more cooling chambers (not shown) within airfoil <NUM> deliver a coolant to tip shroud <NUM> and airfoil cooling chamber <NUM>. Some of the coolant exits cooling chamber <NUM> through openings <NUM> therein and/or may be directed to tip rails <NUM>, but other coolant is directed through one or more sets (pluralities) of cooling passages <NUM> defined in body <NUM> and extending circumferentially therein (along the Y axis). Cooling passages <NUM> appear as rounded ribs in a radial outer surface <NUM> (<FIG>) of body <NUM>. Tip shroud <NUM> has a leading edge <NUM>, a trailing edge <NUM>, a leading circumferentially-facing edge <NUM>, and a trailing circumferentially-facing edge <NUM>. Leading circumferentially-facing edge <NUM> is so termed because it is on pressure side <NUM> of airfoil <NUM>, and trailing circumferentially-facing edge <NUM> is so termed because it is on suction side <NUM> of airfoil <NUM>.

As understood in the field, different extents of tip shroud <NUM> overhanging from airfoil <NUM> can cause mass imbalance. For example, in the illustrative drawings, leading circumferentially-facing edge <NUM> may extend slightly farther from airfoil <NUM> than trailing circumferentially-facing edge <NUM>, creating an imbalance.

Tip shroud <NUM> may also include at least one tip rail <NUM>. Each tip rail <NUM> extends radially outwardly from body <NUM> and extends generally along a circumferential length of body <NUM>. As used herein, "generally" indicates within +/-<NUM>° relative to the direction stated, such as generally along a circumferential length of body <NUM>, or generally parallel to tip rails <NUM>. For purposes of description, tip shroud <NUM> will be mainly illustrated with three axially spaced tip rails 250A-C. Here, tip shroud <NUM> may include a first tip rail 250A extending radially outwardly from body <NUM> and extending generally along a circumferential length of body <NUM>, and a second tip rail 250B extending radially outwardly from body <NUM> and extending generally along the circumferential length of the body. In the non-limiting example shown, tip shroud <NUM> also includes a third tip rail 250C extending radially outwardly from body <NUM> and extending generally along a circumferential length of body <NUM>. It is emphasized that the teachings of the disclosure can be applied to tip shroud <NUM> having any number of tip rails, e.g., one (<FIG>), two (<FIG>) and more than three. Where two or more tip rails <NUM> are used, each tip rail <NUM> is axially spaced (X-axis) from an adjacent tip rail. In the example shown, second tip rail 250B is axially spaced from first tip rail 250A, and third tip rail 250C is axially spaced from second tip rail 250B. As illustrated, a plurality of cooling passages <NUM> extends between circumferentially adjacent tip rail(s) <NUM>, e.g., between each pair of tip rails <NUM> and generally parallel to tip rail(s) <NUM>. A first plurality of cooling passages 246A extends between first and second tip rails 250A, 250B, and a second plurality of cooling passages 246B extends between second and third tip rails 250B, 250C.

<FIG> shows an enlarged cross-sectional view along view line <NUM>-<NUM> in <FIG> of a first edge wall arrangement <NUM>, according to embodiments of the disclosure. As will be described, each first edge wall arrangement <NUM> may be selectively positioned along at least one of leading circumferential-facing edge <NUM> and trailing circumferential-facing edge of the body <NUM>, to reduce mass. In the examples of <FIG> and <FIG>, first edge wall arrangements <NUM> are positioned along leading circumferential-facing edge <NUM> of tip shroud(s) <NUM>, e.g., to address an imbalance in that direction. Hence, in the examples shown in <FIG>, first edge wall arrangements <NUM> are on pressure side <NUM> of airfoil <NUM>.

First edge wall arrangements <NUM> include a first edge wall <NUM> extending axially and radially outwardly from body <NUM> along at least one of leading circumferential-facing edge <NUM> (shown) and trailing circumferential-facing edge <NUM> (e.g., <FIG>) of body <NUM>. As will be described further, tip shroud <NUM> may include more than one first edge wall arrangement <NUM>. In this case, a first edge wall 252A may extend axially and radially outwardly from body <NUM> along at least one of leading and trailing circumferential-facing edges <NUM> (shown), <NUM> (e.g., <FIG>) of body <NUM>, and another first edge wall 252B may extend axially and radially outwardly from body <NUM> along at least one of leading and trailing circumferential-facing edges <NUM>, <NUM> of body <NUM>. First edge wall(s) <NUM> each includes a circumferentially facing surface <NUM>, i.e., facing towards airfoil <NUM>. First edge wall(s) <NUM> acts as a circumferentially outer edge wall of tip shroud <NUM> on pressure side <NUM> and/or suction side <NUM> of airfoil <NUM>. In some instances, first edge walls <NUM> may act as a stiffener for tip shroud <NUM>. Where pairs of tip rails <NUM> are present, first edge wall arrangements <NUM> may extend axially between pairs of tip rails <NUM>. For example, a first edge wall 252A may extend axially between first tip rail 250A and second tip rail 250B, and a first edge wall 252B may extend axially between second and third tip rails 250B, 250C (<FIG>).

<FIG> shows an enlarged cross-sectional view along view line <NUM>-<NUM> in <FIG> of a second edge wall arrangement <NUM> of tip shroud <NUM>. Second edge wall arrangement(s) <NUM> are positioned along at least one of leading circumferential-facing edge <NUM> and trailing circumferential-facing edge <NUM> of body <NUM>, i.e., where first edge wall arrangement <NUM> is not present and where mass reduction may not be necessary. Second edge wall arrangement(s) <NUM> include a second edge wall <NUM> extending axially and radially outwardly from body <NUM> along at least one of leading circumferential-facing edge <NUM> and trailing circumferential-facing edge <NUM> of body <NUM>. In the examples shown in <FIG>, second edge wall arrangements <NUM> are on suction side <NUM> of airfoil <NUM>.

Second edge wall arrangement(s) <NUM> also include an inner wall <NUM> extending axially and radially outwardly from body <NUM>. Where pairs of tip rails <NUM> are present, each second edge wall arrangement <NUM> may extend axially between pairs of tip rails <NUM>. For example, a second edge wall 254A may include an inner wall 258A extending axially between first tip rail 250A and second tip rail 250B, and a second edge wall 254B may include an inner wall 258B extending axially between second and third tip rails 250B, 250C (<FIG>) and radially outwardly from body <NUM>. Inner wall(s) <NUM> is/are parallel to and circumferentially spaced from a respective second edge wall <NUM> to define a radially extending pocket <NUM> therebetween. Radially extending pocket <NUM> is open in a radial outward direction, but closed at a radial inward direction. Inner wall(s) <NUM> has an exit opening <NUM> defined therein for cooling passage(s) <NUM> defined in body <NUM> and extending circumferentially therein. Exit opening(s) <NUM> circumferentially face into radially extending pocket <NUM>. Second edge wall <NUM> may also include at least one opening <NUM> therethrough. Exit opening(s) <NUM> for at least one of cooling passages <NUM> may be linearly aligned with opening(s) <NUM> through second edge wall <NUM>. Hence, coolant exiting cooling passage(s) <NUM> may exit through radially extending pocket <NUM> or through opening(s) <NUM> in second edge wall <NUM>.

Referring to <FIG>, in conventional tip shrouds, both edge walls <NUM>, <NUM> have second edge wall arrangements <NUM>, i.e., the configuration illustrated relative to second edge wall <NUM> in <FIG>. In accordance with embodiments of the disclosure, it has been discovered that inner wall <NUM> provides a mass of material near circumferentially-facing edge(s) <NUM> and/or <NUM> that is not necessary and that may create a mass imbalance that causes bending moments that exacerbate creep issues. As shown in <FIG>, tip shroud <NUM> and, in particular, first edge wall arrangement <NUM>, according to embodiments of the disclosure, includes an exit surface <NUM> adjacent first edge wall <NUM>, rather than inner wall <NUM> (<FIG>). Exit surface <NUM> includes an exit opening <NUM> defined therein through which at least one of plurality of cooling passages <NUM> exits body <NUM>. In <FIG>, two exit openings <NUM> are shown, but any number may be employed - see <FIG> and <FIG>. Exit surface <NUM> is angled (see angle α) relative to circumferentially facing surface <NUM> of first edge wall <NUM> in a range of <NUM>° to <NUM>°. In this manner, in a first edge wall arrangement <NUM>, coolant is projected toward first edge wall <NUM> to cool with similar efficacy as a second edge wall arrangement <NUM> including inner wall <NUM> and pocket <NUM> arrangement (<FIG>), but without the mass of inner wall <NUM>.

Tip shroud <NUM> can be initially manufactured with exit surface <NUM> therein using any now known or later developed manufacturing process, e.g., casting, additive manufacture, etc. Alternatively, exit surface <NUM> may be formed in a tip shroud <NUM> manufactured with inner wall <NUM> (<FIG>) on first edge wall <NUM> and may be machined to remove inner wall <NUM> (<FIG>), e.g., grinding, cutting, or otherwise physically removing inner wall <NUM>, to create exit surface <NUM> (<FIG>). In any event, as shown best in <FIG>, first edge wall <NUM> and exit surface <NUM> may extend axially between a pair of axially opposing, radially extending walls <NUM> in body <NUM>. Radially extending walls <NUM> may be spaced from, for example, respective tip rails, such as tip rails 250A, 250B. One of the radially extending walls <NUM> is also shown in <FIG>.

As shown in <FIG> and <FIG>, first edge wall <NUM> may include at least one opening <NUM> therethrough, so coolant exiting exit opening(s) <NUM> in exit surface <NUM> can cool first edge wall <NUM> and other downstream structure. In one non-limiting example, exit opening(s) <NUM> for at least one of plurality of cooling passages <NUM> may be linearly aligned with opening(s) <NUM> in first edge wall <NUM>, e.g., there could be one-to-one alignment of openings. However, this is not necessary in all instances. Any number of exit openings <NUM> and/or openings <NUM> may be employed.

As shown in <FIG>, where tip shroud <NUM> includes more than two tip rails <NUM>, i.e., 250A-C, the above-described arrangement may be repeated between the different pairs of tip rails <NUM>. More particularly, as shown between tip rails 250B, 250C, tip shroud <NUM> may include another first edge wall arrangement <NUM> including first edge wall 252B extending axially and radially outwardly from body <NUM>, e.g., between tip rails 250B-C, on pressure side <NUM> of airfoil <NUM>. First edge wall 252B includes circumferentially facing surface <NUM> (<FIG>). Between tip rails 250B, 250C, another plurality of cooling passages 246B are defined in body <NUM> and extend circumferentially therein. Here, a second exit surface 270B is adjacent first edge wall 252B. Second exit surface 270B has the same arrangement as shown in <FIG>. That is, exit surface 270B (<FIG>) has an exit opening 272B (<FIG>) defined therein through which at least one of cooling passages 246B exits body <NUM>. Second exit surface 270B is angled (angle α) relative to circumferentially facing surface <NUM> of first edge wall 252B in a range of <NUM>° to <NUM>°. Any number of first edge wall arrangements <NUM> can be employed.

<FIG> shows a radially outward perspective view of trailing edge <NUM> of airfoil <NUM> and leading circumferential-facing edge <NUM> of tip shroud <NUM>; and <FIG> shows a radially inward perspective view of trailing edge <NUM> of airfoil <NUM> and leading circumferential-facing edge <NUM> of tip shroud <NUM>. Referring to <FIG>, <FIG> and <FIG>, to further reduce mass imbalance, tip shroud <NUM> may include a curved opening <NUM> defined in a trailing edge <NUM> of body <NUM> adjacent trailing edge <NUM> of airfoil <NUM>. In the example shown, curved opening <NUM> is on pressure side <NUM> of airfoil <NUM>. Curved opening <NUM> is formed in an area of body <NUM> that normally extends circumferentially forward of aft-most tip rail 250C and trailing edge <NUM> of airfoil <NUM>. In some situations, adjacent tip shrouds <NUM> have interlocking surfaces, sometimes referred to as Z-notches for their Z-like shape. Here, as shown in <FIG>, curved opening <NUM> is not filled by body <NUM> of an adjacent tip shroud <NUM> and does not interlock with an adjacent surface. In one embodiment, curved opening <NUM> defines a plane angled (angle β) in a range of <NUM>° and <NUM>° relative to a radial direction Z.

<FIG> shows a perspective view of tip shroud <NUM>, according to alternative embodiments of the disclosure. <FIG> shows an embodiment including first edge wall arrangements <NUM> along both leading circumferentially-facing edge <NUM> and trailing circumferentially-facing edge <NUM>. Here, a mass imbalance may not be present, but mass reduction is still desired.

<FIG> shows a perspective view of tip shroud <NUM>, according to further alternative embodiments of the disclosure. <FIG> shows an embodiment including one first edge wall arrangement <NUM> along leading circumferentially-facing edge <NUM> at only one location (e.g., between tip rails 250A-B). Here, a mass imbalance may be present only in tip shroud <NUM> near leading edge <NUM> of airfoil <NUM>, such that less mass reduction is desired than the <FIG> embodiment.

<FIG> shows a perspective view of tip shroud <NUM>, according to other embodiments of the disclosure. <FIG> shows an embodiment including first edge wall arrangements <NUM> as in <FIG>, but without tip rails 250B-C. That is, only one tip rail 250A is present. As illustrated, where not confined by tip rails <NUM>, first and second wall arrangements <NUM>, <NUM> may have any desired axial length.

<FIG> shows a perspective view of tip shroud <NUM>, according to other embodiments of the disclosure. <FIG> shows an embodiment including first edge wall arrangements <NUM> along leading circumferentially-facing edge <NUM>, but with only two tip rails 250A-B. It will be recognized that first edge wall arrangement <NUM> can be used in any location on tip shroud <NUM> in which mass reduction is desired, e.g., to address a mass imbalance. Other arrangements, not illustrated, may also be possible.

As illustrated, for example, in <FIG>, exit surface <NUM> has a planar surface. In other embodiments, exit surface <NUM> may generally have the angle with circumferentially facing surface <NUM> of first edge wall <NUM>, but may not be planar. <FIG> show cross-sectional views of exit surface <NUM> of tip shroud <NUM>, according to alternative embodiments of the disclosure. In <FIG>, exit surface <NUM> is stepped. Here, exit surface <NUM> may be made by a number of machining steps, and angle α may be defined by a consistent element of the steps, e.g., outer corners thereof. In <FIG>, exit surface <NUM> may be slightly arced, e.g., inwardly, while generally retaining angle α, e.g., from a point where exit surface <NUM> meets first edge wall <NUM> and a radially outermost point of exit surface <NUM> over exit openings <NUM> of cooling passages <NUM>. Exit surface <NUM> can have a number of other shapes within the scope of the disclosure.

Embodiments of the disclosure provide a tip shroud with a first edge wall arrangement including an exit surface that includes exit opening(s) defined therein, through which at least one of the tip shroud's cooling passages exits the body. The creation of the exit surface is accomplished by removal of mass that otherwise contributes to mass imbalance, while retaining the cooling efficacy of a second wall arrangement. Embodiments of the disclosure can also provide the trailing edge of the tip shroud body with a curved opening near the trailing edge of the airfoil to remove additional mass.

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 tip shroud (<NUM>) for a turbine blade (<NUM>), comprising:
a body (<NUM>) coupled to a radial outer end (<NUM>) of an airfoil (<NUM>) of the turbine blade (<NUM>), the airfoil (<NUM>) including a pressure side (<NUM>) and a suction side (<NUM>) and the body (<NUM>) including a leading circumferential-facing edge (<NUM>) and a trailing circumferential-facing edge (<NUM>);
at least one tip rail (<NUM>) extending radially outwardly from the body (<NUM>) and extending along a circumferential length of the body (<NUM>);
a first plurality of cooling passages (246A) defined in the body (<NUM>) and extending circumferentially therein; and
at least one first edge wall arrangement (<NUM>) along at least one of the leading circumferential-facing edge (<NUM>) and the trailing circumferential-facing edge (<NUM>) of the body (<NUM>), each first edge wall arrangement (<NUM>) including:
a first edge wall (<NUM>) extending axially and radially outwardly from the body (<NUM>) along the at least one of the leading circumferential-facing edge (<NUM>) and the trailing circumferential-facing edge (<NUM>) of the body (<NUM>), the first edge wall (<NUM>) including a first circumferentially facing surface (<NUM>) facing towards the airfoil (<NUM>), and, characterized in that the tip shroud further comprises
an exit surface (<NUM>) adjacent the first edge wall (<NUM>) and facing the first circumferentially facing surface (<NUM>), the exit surface (<NUM>) having an exit opening (<NUM>) defined therein through which at least one of the first plurality of cooling passages (246A) exits the body (<NUM>) to allow coolant exiting the exit opening (<NUM>) in the exit surface (<NUM>) to cool the first edge wall (<NUM>),
wherein the exit surface (<NUM>) is angled relative to the first circumferentially facing surface (<NUM>) of the first edge wall (<NUM>) in a range of <NUM>° to <NUM>°.