Patent Description:
Turbomachines include a plurality of turbine rotor blades coupled to a rotor. A working fluid such as steam or a combusted fuel is forced against the blades to force them to turn the rotor. Turbine rotor blades operate in extremely hot conditions and require cooling. Cooling features can be provided in a number of ways.

One mechanism to provide cooling is an impingement insert. An impingement insert or sleeve includes a hollow body having cooling passages in a wall thereof that allow delivery of a coolant through the cooling passages to impact or impinge on a surface to be cooled. Impingement inserts are used, for example, in a variety of hot gas path (HGP) components in turbomachinery such as a turbine rotor blade to increase cooling performance of cooling circuitry therein. One challenge with impingement inserts is positioning the impingement insert within a tapered or curved cavity in an HGP component in a sufficiently close manner to allow for high cooling performance, but not so close that cooling is ineffective. One indicator of cooling performance of an impingement insert is the Z/D parameter, which is a ratio of a standoff distance Z between the insert and interior surface of the HGP component and a diameter D of the cooling passages (holes) in the impingement insert. The Z/D parameter value of an insert is typically designed to be within a desired range that results in better cooling performance.

Impingement cooling is typically not provided by an insert where the necessary standoff distance cannot be created. For example, where the cavity in the HGP component curves too significantly that an impingement insert cannot be made sufficiently thin or curved to respect the necessary standoff distance, impingement cooling cannot be provided. One approach to address this challenge provides the impingement insert in a number of flexible, longitudinal sections to make it easier to insert them into the HGP component. However, having to sequentially position and couple a number of insert sections together or couple them to the HGP component, increases the complexity, time and costs of manufacture. Flexible impingement insert sections also do not provide a contiguous element about their periphery, i.e., laterally (cross-section), which can detract from cooling performance where they are discontinuous.

Impingement cooling has been applied in a limited manner to rotating turbine rotor blades in turbomachines, e.g., for leading edges thereof. However, impingement cooling has not been applied more broadly across an entirety of an inner surface of a turbine rotor blades because the centrifugal forces experienced by the rotating blades forces the coolant to the radially outer tip end of the blade as it rotates, making impingement cooling less effective.

Another cooling feature includes cooling passages through a part of the turbine rotor blade to be cooled. For example, the turbine rotor blades include a platform that extends laterally to form part of a working fluid path through the turbomachine in cooperation with a platform of an adjacent turbine rotor blade. Due to the high temperatures of the working fluid, the platforms typically include a cooling circuit therein that feeds a number of cooling passages that exit through a slash face of the platform. Some platforms include a damping pin seat in the slash face that receives an axially extending pin therein that mates with an adjacent damping pin seat in an adjacent platform to seal the working fluid path. The cooling passages are typically drilled into the slash face to fluidly couple the passages to the cooling circuit. Consequently, the cooling passages have a linear configuration that may not adequately cool all of the platform. For example, the cooling passages may pass through an extension that forms a damping pin seat, but inadequately cool other portions of the slash face.

Cooling features may also be employed with angel wings. In this regard, another cooling feature includes cooling passages that either deliver coolant into the angel wing, or radially around an angel wing. Mounts for turbine rotor blades may also include cooling features therein. <CIT> relates to an engine component having a platform with a passageway. <CIT> relates to a platform cooling arrangement in a turbine rotor blade. <CIT> relates to an apparatus for cooling a platform of a turbine component.

A first aspect of the invention provides a turbine rotor blade, according to claim <NUM>.

A second aspect of the invention provides a turbine rotor blade according to claim <NUM>.

A third aspect of the invention provides a turbine rotor blade according to claim <NUM>.

The illustrative aspects of the present invention are designed to solve the problems herein described and/or other problems not discussed.

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 invention, it will become necessary to select certain terminology when referring to and describing relevant machine components within, for example, a turbomachine. When doing this, if 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 turbomachine or, for example, the flow of air through the combustor or coolant through one of the turbine's components. The term "downstream" corresponds to the direction of flow of the fluid, and the term "upstream" refers to the direction opposite to the flow. The terms "forward" and "aft," without any further specificity, refer to directions, with "forward" referring to the front end of the turbomachine (i.e., compressor end) or a component thereof, and "aft" referring to the rearward end of the turbomachine (i.e., turbine end) or component thereof. Forward and aft generally denoted by an X direction in the drawings. It is often required to describe parts that are at differing radial positions with regard to a center axis. The term "radial" refers to movement or position perpendicular to an axis, e.g., a turbomachine rotor axis. In cases such as this, 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. Radial direction is generally denoted by a Z direction in the drawings. The term "axial" refers to movement or position parallel to an axis, i.e., a turbomachine rotor axis. Finally, the term "circumferential" refers to movement or position around an axis. Although not shown as curved in the legends in the drawings, the circumferential direction is generally denoted by a Y direction in the drawings. It will be appreciated that such terms may be applied in relation to the rotor axis of a turbomachine.

Where an element or layer is referred to as being "on," "engaged to," "connected to," or "coupled to" another element or layer, it may be directly on, engaged, connected or coupled to the other element or layer, or intervening elements or layers may be present. In contrast, when an element is referred to as being "integral," "directly on," "directly engaged to," "directly connected to," or "directly coupled to" another element or layer, there may be no intervening elements or layers present.

As indicated above, the invention provides a turbine rotor blade or turbine rotor blade root including a number of integral features which are made possible through additive manufacturing of the blade and/or root. The additive manufacturing allows formation of structures that provide cooling where not previously allowed, improve cooling compared to conventional systems, provide additional structural strength and/or lower the weight of the blade.

<FIG> shows a schematic view of an illustrative turbomachine <NUM> that may include a turbine rotor blade including integral features according to various embodiments of the invention. In the example shown, turbomachine <NUM> includes a gas turbine (GT) system <NUM> that 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 shaft (sometimes referred to as a rotor) <NUM>. In one embodiment, GT system <NUM> is a 7HA or 9HA GT system, commercially available from General Electric Company, Greenville, S. The present invention is not limited to any one particular GT system and may be employed in connection with other engines including, for example, the other HA, F, B, LM, GT, TM and E-class engine models of General Electric Company, and engine models of other companies. Further, a turbine rotor blade, as described herein, may find application in other forms of turbomachines, e.g., steam turbines, jet engines, compressors, etc..

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>. 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> for which gas stream thermal energy is converted to mechanical rotational energy. Turbine <NUM> is rotatably coupled to and drives rotor <NUM>. Compressor <NUM> also is rotatably coupled to rotor <NUM>. In the illustrative embodiment, there is a plurality of combustors <NUM> and fuel nozzle assemblies <NUM>.

<FIG> shows a cross-sectional view of an illustrative turbine <NUM> with three stages that may be used with GT system <NUM> in <FIG>. Each stage includes sets of stationary vanes or nozzles <NUM> and turbine rotor blades <NUM>. Stationary nozzles <NUM> may be held in turbine <NUM> by a radially outer platform <NUM> and a radially inner platform <NUM>. Stationary nozzles <NUM> may include one or more circumferentially spaced airfoils <NUM> (<FIG>). Turbine rotor blades <NUM> are coupled to rotor <NUM> and extend between rows of stationary nozzles <NUM>. Combustion gases are directed by stationary nozzles <NUM> against turbine rotor blades <NUM> to turn rotor <NUM> (<FIG>).

<FIG> shows a perspective view of an illustrative turbine rotor blade <NUM> of GT system <NUM> in which integral features according to various embodiments of the invention may be employed.

<FIG> shows an axial cross-sectional view of turbine rotor blade <NUM> including an integral feature in the form of an integral impingement cooling structure <NUM> according to various embodiments of the invention.

Turbine rotor blade <NUM> includes an airfoil body <NUM> including a concave pressure side outer wall <NUM> and a convex suction side outer wall <NUM> that connect along leading and trailing edges <NUM>, <NUM>. As shown in <FIG>, outer walls <NUM>, <NUM> have an airfoil inner surface <NUM> defining a radially extending chamber <NUM> for receiving a coolant flow <NUM>. Referring again to <FIG>, turbine rotor blade <NUM> may also include a tip end <NUM> at a radial outer end <NUM> of airfoil body <NUM>. Turbine rotor blade <NUM> may also include a turbine rotor blade root <NUM> (hereinafter "root <NUM>") by which turbine rotor blade <NUM> attaches to rotor <NUM> (<FIG>), e.g., by a rotor wheel <NUM> (<FIG>).

For purposes of this invention, root <NUM> may include any portion of turbine rotor blade <NUM> including, and radially inward of, platform <NUM>. Root <NUM> may include a blade mount <NUM> configured for mounting in a corresponding slot in the perimeter of a rotor wheel <NUM> (<FIG>). Blade mount <NUM> may have any now known or later developed outer configuration for mounting to rotor disk <NUM> (<FIG>) such as but not limited to a dovetail or fir tree arrangement. Turbine rotor blade <NUM>, i.e., root <NUM> thereof, may further include a shank <NUM> that extends between blade mount <NUM> and a platform <NUM>. Platform <NUM> is disposed at the junction of airfoil body <NUM> and shank <NUM> and defines a portion of the inboard boundary of the flow path through turbine <NUM> (<FIG>). Shank <NUM> is thus located at a radial inner end <NUM> of airfoil body <NUM>, and blade mount <NUM> is located radially inward of shank <NUM>. Platform <NUM> extends laterally outward relative to shank <NUM>. As will be described further herein, radially extending chamber <NUM> may extend at least partially into shank <NUM> to define a shank inner surface <NUM> (<FIG>). Outer walls <NUM> and <NUM> of airfoil body <NUM> extend in the radial (Z) direction from platform <NUM> to tip end <NUM>. It will be appreciated that airfoil body <NUM> is the active component of turbine rotor blade <NUM> that intercepts the flow of working fluid and induces the rotor to rotate.

In certain embodiments, turbine rotor blade <NUM> may include, inter alia, airfoil body <NUM> and an integral feature in the form of an integral impingement cooling structure <NUM> therein. The impingement cooling structure is not an insert, but is made integrally through, e.g., additive manufacture, with the rest of the blade. As noted herein, airfoil body <NUM> may include concave pressure side outer wall <NUM> and convex suction side outer wall <NUM> that connect along leading and trailing edges <NUM>, <NUM>. The outer walls <NUM>, <NUM> have airfoil inner surface <NUM> defining radially extending chamber <NUM> for receiving coolant flow <NUM>. Turbine rotor blade <NUM> may also include tip end <NUM> at radial outer end <NUM> of airfoil body <NUM>, and shank <NUM> at radial inner end <NUM> of airfoil body <NUM>. Radially extending chamber <NUM> may extend at least partially into shank <NUM> to define shank inner surface <NUM>. Integral impingement cooling structure <NUM> is within radially extending chamber <NUM>, and may include hollow body <NUM> including a first end <NUM>, a second end <NUM>, an interior surface <NUM> and an exterior surface <NUM>. A plurality of cooling passages <NUM> extend through hollow body <NUM> and are in fluid communication with the radially extending chamber <NUM> to allow the coolant flow to pass from interior surface <NUM> of hollow body <NUM> to impinge on the inner surface of at least airfoil body <NUM>. In contrast to conventional impingement inserts, first end <NUM> of hollow body <NUM> is integrally formed to the shank inner surface <NUM>, i.e., via the additive manufacturing. Consequently, exterior surface <NUM> of hollow body <NUM> may be made to be uniformly spaced from airfoil inner surface <NUM> between first end <NUM> and second end <NUM> of the hollow body <NUM>, regardless of the curvature of the airfoil inner surface <NUM>. In another embodiment, non-uniform, but custom-built, standoff spacing may be employed for purposes of providing different impingement cooling, heat pick-up and/or re-use. For example, closer standoff spacing may be employed where increased impingement cooling is required, and wider standoff spacing used where less impingement cooling is required. Further, the hollow body <NUM> may have cooling passages about an entirety of its periphery and radial span to provide impingement cooling throughout the blade, not just at a leading edge thereof. Hence, the integral impingement cooling structure allows for maximum coverage of impingement with limited sacrifices normally associated with impingement inserts, and can have a number of variable cooling features. For example, the turbine rotor blade may have: variable chordwise width for the impingement cooling structure or a pin bank aft thereof, tailored impingement cooling structure wall thickness for different cooling loads, and varied supports to address different coefficients of thermal expansion (CTE) between the airfoil body and impingement cooling structure.

As shown in <FIG>, and the radial, circumferential cross-sectional view of <FIG>, turbine rotor blade <NUM> may include impingement cooling structure <NUM> within radially extending chamber <NUM>. Impingement cooling is typically provided by one or more impingement inserts that are inserted into a radially extending chamber <NUM> and coupled to airfoil body <NUM>, e.g., by fasteners or welding. The impingement inserts are typically linear, but may include some curvature. Where radially extending chamber <NUM> has a curved airfoil inner surface <NUM> as in <FIG>, it is impossible to have impingement inserts uniformly spaced from the inner surface along an entire radial span of the blade. In order to address this challenge, impingement cooling structure <NUM> according to embodiments of the invention is integrally formed with the rest of turbine rotor blade <NUM>, via additive manufacturing.

As shown in <FIG>, impingement cooling structure <NUM> includes hollow body <NUM> including first end <NUM>, second end <NUM>, interior surface <NUM> and exterior surface <NUM>. Impingement cooling structure <NUM> also includes plurality of cooling passages <NUM> through hollow body <NUM> and in fluid communication with radially extending chamber <NUM> to allow coolant flow <NUM> to pass from interior surface <NUM> of the hollow body to impinge on at least airfoil inner surface <NUM>, e.g., it may impinge inner surfaces in, inter alia, airfoil body <NUM>, tip end <NUM>, shank <NUM>, and/or platform <NUM>. In contrast to conventional turbine rotor blades, and as shown in <FIG> and especially <FIG>, exterior surface <NUM> of hollow body <NUM> is uniformly spaced from airfoil inner surface <NUM> between first end <NUM> and second end <NUM> of hollow body <NUM>. That is, using additive manufacturing rather than mechanical insertion, impingement cooling structure <NUM> can be formed (simultaneously with, e.g., airfoil inner surface <NUM>) to have the same curvature, bends, twists and any other shape or dimension, to match that of the inner surface adjacent thereto. Notably, impingement cooling structure <NUM> can be uniformly spaced from airfoil inner surface <NUM> along an entire radial span that it covers, ensuring the desired Z/D parameter over all of turbine rotor blade <NUM>. The Z/D parameter is a ratio of a standoff distance Z between exterior surface <NUM> and an inner surface (e.g., airfoil inner surface <NUM>, shank inner surface <NUM>, etc.) of turbine rotor blade <NUM> and a diameter D of cooling passages <NUM> (holes) in impingement cooling structure <NUM>. In one example, Z/D ranges from approximately <NUM> to approximately <NUM>. In another example, Z/D may range from approximately <NUM> to approximately <NUM>. A standoff distance Z may be smaller than conventionally available for castings, e.g., less than approximately <NUM> millimeters (<NUM> inches). Smaller diameter D cooling passages <NUM> than conventional castings may also be employed, e.g., as a function of clogging from debris. Advantageously, cooling passages <NUM> may extend around an entire peripheral extent of hollow body <NUM> such that coolant flow <NUM> exits hollow body <NUM> in all directions to provide impingement cooling to all of airfoil inner surface <NUM> of airfoil body <NUM>. Alternatively, cooling passages <NUM> may be omitted in areas where impingement cooling of inner surfaces <NUM>, <NUM> is not desired or required. Cooling passages <NUM> may extend along any desired radial extent of hollow body <NUM>.

As shown in <FIG>, first end <NUM> of hollow body <NUM> is integrally formed to shank inner surface <NUM>. First end <NUM> meets shank inner surface <NUM> at a meeting location <NUM>, which extends about the entire periphery of first end <NUM>, i.e., there are no openings between first end <NUM> and shank inner surface <NUM> at meeting location (other than perhaps a cooling passage <NUM>). In certain embodiments, first end <NUM> of hollow body <NUM> is integrally formed to shank inner surface <NUM> radially inward of platform <NUM>. However, this particular meeting location <NUM> may not be necessary in all instances, e.g., in some cases, meeting location <NUM> may be radially outward of platform <NUM>. As shown in <FIG>, although not necessary in all cases, second end <NUM> of hollow body <NUM> may also be integrally formed to inner surface <NUM> of tip end <NUM>. Cooling passages <NUM> may optionally provide impingement cooling of tip end <NUM>, or pass coolant to tip end <NUM> for other forms of cooling.

<FIG> shows an enlarged cross-sectional view of meeting location <NUM> (<FIG>) of impingement cooling structure <NUM> and shank <NUM> of turbine rotor blade <NUM>, according to various embodiments of the invention.

As shown in <FIG>, first end <NUM> of hollow body <NUM> may extend substantially in a radial direction (arrow Z) relative to meeting location <NUM> of first end <NUM> of hollow body <NUM> and shank inner surface <NUM>. As used herein, "substantially in a radial direction" indicates that first end <NUM> extends radially away from rotor <NUM> (<FIG>) with some degree of tolerance, e.g., +/-<NUM>°. In contrast, at least a portion of the shank inner surface <NUM> extends at an angle α relative to radial direction Z from meeting location <NUM> of first end <NUM> of hollow body <NUM> and shank <NUM> (<FIG>). In another embodiment, shank inner surface <NUM> is aligned substantially in a radial direction and first end <NUM> of hollow body gradually curves or transitions towards meeting location <NUM> to keep angle α, for example, < <NUM>°. In a further embodiment, both shank inner surface <NUM> and first end <NUM> of hollow body <NUM> gradually curve or transition towards meeting location <NUM>. Angle α may be any angle desired and within the range of additive manufacture, e.g., < <NUM>° from vertical. To maintain structural integrity, it is desirable that angle α be as small as possible, e.g., < <NUM>°, < <NUM>° or < <NUM>°. As shown only in <FIG>, in certain embodiments, a support structure <NUM> may be positioned between first end <NUM> of hollow body <NUM> and shank inner surface <NUM>, e.g., radially outward of meeting location <NUM> and radially inward of platform <NUM>. In further embodiments, support structures <NUM> may be positioned at any location between exterior surface <NUM> of hollow body <NUM> and airfoil inner surface <NUM>, shank inner surface <NUM>, etc. In still further embodiments, at least portions of support structure <NUM> include hollow support elements (e.g., lattice)to enable cooling flow directly from chamber <NUM> to outer walls <NUM> and/or <NUM> (<FIG>) of airfoil body <NUM>. For example, it may be desirable to provide film cooling to certain areas of airfoil body <NUM>, such as leading and/or trailing edges <NUM>, <NUM>, directly from radially extending chamber <NUM>. Support structure <NUM> may include any now known or later developed element(s) capable of positioning first end <NUM> of hollow body <NUM> relative to shank inner surface <NUM>. Support structure <NUM> may include, but is not limited to: a lattice structure, straight or arced bar(s), etc. Support structure <NUM> may also be integrally formed via additive manufacture.

Impingement cooling structure <NUM> may also include a variety of optional alternative integral cooling features. In one example, impingement cooling structure <NUM> may be optionally formed with varying wall thicknesses. Varying wall thicknesses may be advantageous to, for example, accommodate varying CTEs between impingement cooling structure <NUM> and hotter airfoil body <NUM>, shank <NUM> and/or platform <NUM>. As observed in <FIG>, airfoil body <NUM>, shank <NUM> and/or platform <NUM> may have a wide variety of wall thicknesses, and may have varying thicknesses over an extent thereof. <FIG> shows an enlarged, partial cross-sectional view of a portion of impingement cooling structure <NUM> adjacent airfoil body <NUM>, platform <NUM> or shank <NUM>. As noted, in certain embodiments, as shown in <FIG>, hollow body <NUM> may include at least one first portion <NUM> having a first wall thickness W1 between interior surface <NUM> and exterior surface <NUM> thereof, and at least one second portion <NUM> having a second wall thickness W2 between interior surface <NUM> and exterior surface <NUM> thereof. In the example shown, first wall thickness W1 is greater than second wall thickness W2. Any number of thicker and/or thinner portions <NUM>, <NUM> may be provided in impingement cooling structure <NUM>. The thickness of portions <NUM>, <NUM> may be any dimension desired to address the structural and/or thermal requirements of the location.

In another example optional structure, additional support may be desired and/or required to support integral impingement cooling structure <NUM> relative to inner surfaces <NUM>, <NUM>. For example, additional support may be desired and/or required at thinner wall thickness portions <NUM> (<FIG>) of impingement cooling structure <NUM>. To this end, as shown in <FIG>, turbine rotor blade <NUM> may also include a support <NUM> on exterior surface <NUM> of hollow body <NUM> in at least one portion <NUM> having thinner wall thickness W2. Support <NUM> may be integrally formed with hollow body <NUM> (and rest of turbine rotor blade <NUM>) to space exterior surface <NUM> of hollow body <NUM> from, e.g., airfoil inner surface <NUM>, between first end <NUM> (<FIG>) and second end <NUM> of hollow body <NUM>. Any number of supports <NUM> may be provided in thinner wall portion(s) <NUM>. Support(s) <NUM> may include a passage(s) <NUM> therethrough in fluid communication with one of the plurality of cooling passages <NUM>, i.e., to allow coolant flow <NUM> to pass therethrough and impinge inner surface(s) <NUM>, <NUM>. In certain embodiments, regardless of wall thickness, turbine rotor blade <NUM> may include a support(s) <NUM> on exterior surface <NUM> of hollow body <NUM>. Support <NUM> may be integrally formed with hollow body <NUM> (and rest of turbine rotor blade <NUM>) to space exterior surface <NUM> of hollow body <NUM> from, e.g., airfoil inner surface <NUM>, between first end <NUM> (<FIG>) and second end <NUM> of hollow body <NUM>. Supports <NUM>, <NUM> may take any form that allows: reduction of stress between hotter outer walls <NUM>, <NUM> of airfoil body <NUM> and cooler impingement cooling structure <NUM>, provide any necessary thermal expansion, provide structural support, and/or desired spacing of hollow body <NUM> from inner surface(s) <NUM>, <NUM>. Supports <NUM>, <NUM> can have any desired dimension and/or shape such as but not limited to: tubes, bars, etc..

<FIG> shows an enlarged cross-sectional view of another alternative embodiment including a reinforcement member <NUM> surrounding at least one of cooling passages <NUM>. Reinforcement member <NUM> may include any structurally strengthening member such as a thicker wall, etc. Certain embodiments, as shown in <FIG>, may also include a stiffener rib <NUM> integrally formed to interior surface <NUM> of hollow body <NUM>. Any number of stiffener ribs <NUM> may be provided, and each may extend any desired radial extent of hollow body <NUM>. Supports <NUM>, <NUM>, reinforcement member <NUM> and/or stiffener rib <NUM> may be integrally formed with the rest of turbine rotor blade <NUM> via additive manufacture.

<FIG> shows a cross-sectional view of turbine rotor blade <NUM> including integral impingement cooling structure <NUM> and additional optional alternative integral cooling features. In one alternative embodiment, impingement cooling structure <NUM> may be optionally formed with varying spacing Z from inner surface <NUM>, <NUM>. The spacing Z may be customized to provide the desired Z/D parameter and desired cooling at various locations. For example, turbine rotor blade <NUM> may have a number of high heat load regions <NUM>, i.e., regions that experience higher temperatures and require more cooling compared to other regions of the blade. In the example shown, high heat load regions <NUM> include regions: near leading edge <NUM> (195A), pressure side outer wall <NUM> near trailing edge <NUM> (195B), and suction side wall <NUM> downstream of leading edge <NUM> (195C). At high heat load regions <NUM>, a first spacing Z1 may be employed between integral impingement cooling structure <NUM> and inner surface <NUM>, <NUM> at high heat load regions <NUM>, while a second, larger spacing Z2 is used at other locations that do not have such a high heat load. In this manner, more cooling can be provided where necessary, i.e., at high heat load regions <NUM>, using first spacing Z1 with the spacing increasing between impingement cooling structure <NUM> and inner surface <NUM>, <NUM> to second, larger spacing Z2 for lower heat load regions. As shown in <FIG>, the larger second spacing Z2 may allow coolant flow <NUM> to limit or reduce heat absorption as it moves downstream toward trailing edge <NUM>, allowing coolant flow <NUM> to be cooler and have more heat absorbing capacity for downstream regions, e.g., serpentine cooling passage <NUM> and/or pin bank <NUM> (described herein). A transition between spacings Z1 and Z2 can be at any desired rate, e.g., gradual over a relatively long distance, abrupt at a particular location, or at any rate therebetween. Second spacing Z2 may be anywhere from, for example, <NUM> to <NUM> times first spacing Z1. The Z/D parameter can be customized for each region of concern. As noted, in one example, Z/D ranges from approximately <NUM> to approximately <NUM>. In another example, Z/D may range from approximately <NUM> to approximately <NUM>. Diameter D of cooling passages <NUM> can also be configured to customize the Z/D parameter for different regions.

<FIG> also shows turbine rotor blade <NUM> including one or more post-impingement target features <NUM> on inner surface <NUM>. Post-impingement target features <NUM> may include any now known or later developed structure on inner surface <NUM> to promote cooling. In the example shown, impingement target features <NUM> include bumps, but they could include any structure. In one scenario, hollow body <NUM> may include a local bulge <NUM> to match impingement target feature(s) <NUM> contour, and thus maintain spacing Z (i.e., Z1 as shown). While two target features <NUM> and bulge <NUM> pairs are shown, any number may be employed. In one embodiment, post-impingement target features <NUM> may also optionally include additional integral cooling features, such as but not limited to film cooling holes <NUM>. Film cooling holes <NUM> direct coolant flow <NUM>, post-impingement with post-impingement target features <NUM>, i.e., inner surface <NUM> thereof, to create a cooling film <NUM> over sidewall(s) <NUM>, <NUM>. Any number of film cooling holes <NUM> may be applied within each post-impingement cooling features <NUM>.

<FIG> shows a first radial, cross-sectional view along view line <NUM>-<NUM> in <FIG>, and <FIG> shows a second, radial cross-sectional view along view line <NUM>-<NUM> in <FIG>, the latter of which is in a slightly different plane than <FIG> and in the opposite direction. As shown in <FIG>, in certain embodiments, hollow body <NUM> has a chordwise width WC1 that is smaller near tip end <NUM> than shank <NUM>. Most conventional impingement inserts have an opposite chordwise width arrangement to allow them to be inserted through an open tip end of the airfoil body. Further, hollow body <NUM> may have alternating wider and narrow chordwise widths WC1 over its radial span (up and down page in <FIG>). Consequently, an axially aft end <NUM> of hollow body <NUM> may vary in a chordwise location along a radial span of hollow body <NUM>. In this manner, impingement cooling structure <NUM> can have a shape (chordwise width WC1) that curves over its radial span to be uniformly spaced from airfoil inner surface <NUM> and/or shank inner surface <NUM> (<FIG>), regardless of the latter's shape. Conventional impingement inserts cannot provide such features.

As shown in <FIG>, <FIG> and <FIG>, airfoil body <NUM> further includes at least one chordwise extending, serpentine cooling passage <NUM> extending from airfoil inner surface <NUM> aft of hollow body <NUM> toward trailing edge <NUM>. As shown best in <FIG> and <FIG>, each chordwise extending, serpentine cooling passage <NUM> may have a same chordwise width WC2, e.g., shorter than chordwise width WC1 of hollow body <NUM>. Airfoil body <NUM> may also include a plurality of radially spaced, trailing edge cooling passages <NUM> extending through trailing edge <NUM>, i.e., from serpentine cooling passage(s) <NUM>. Each of the plurality of trailing edge cooling passages <NUM> has a same chordwise width WC3, i.e., along a radial span of turbine rotor blade <NUM>. As illustrated in <FIG> and <FIG>, a space <NUM> between trailing edge cooling passages <NUM> and serpentine cooling passage(s) <NUM> has a varying chordwise width WC4, i.e., along a radial span of turbine rotor blade <NUM>. Turbine rotor blade <NUM> may further include a pin bank <NUM> between a forward end of plurality of trailing edge cooling passages <NUM> (right side thereof in <FIG>, left side thereof in <FIG>) and an aft end of chordwise extending, serpentine cooling passage(s) <NUM> (left side thereof in <FIG>, right side thereof in <FIG>). Consequently, as illustrated, in <FIG>, pin bank <NUM> may have the varying chordwise width WC4 along a radial span thereof. <FIG> and <FIG> illustrate a variety of cooling features including but not limited to: film cooling via openings <NUM>, main chord impingement cooling via impingement cooling structure <NUM>, near trailing edge <NUM> cooling via serpentine cooling passage(s) <NUM>, and trailing edge <NUM> pin bank cooling via pin bank <NUM>.

Additive manufacturing (AM) includes a wide variety of processes of producing a component through the successive layering of material rather than the removal of material. As such, additive manufacturing can create complex geometries, such as those described herein relative to turbine rotor blade <NUM>, without the use of any sort of tools, molds or fixtures, and with little or no waste material. Instead of machining components from solid billets of material, much of which is cut away and discarded, the only material used in additive manufacturing is what is required to shape the component. Additive manufacturing techniques typically include taking a three-dimensional computer aided design (CAD) file of the component (e.g., turbine rotor blade <NUM>) to be formed, electronically slicing the component into layers, e.g., <NUM>-<NUM> micrometers thick, and creating a file with a two-dimensional image of each layer, including vectors, images or coordinates. The file may then be loaded into a preparation software system that interprets the file such that the component can be built by different types of additive manufacturing systems. In 3D printing, rapid prototyping (RP), and direct digital manufacturing (DDM) forms of additive manufacturing, material layers are selectively dispensed, sintered, formed, deposited, etc., to create the component. While other manufacturing processes such as casting may also be employed, turbine rotor blade <NUM> may be advantageously made by additive manufacturing.

In metal powder additive manufacturing techniques, such as direct metal laser melting (DMLM) (also referred to as selective laser melting (SLM)), direct metal laser sintering (DMLS), selective laser sintering (SLS), electron beam melting (EBM), and perhaps other forms of additive manufacturing, metal powder layers are sequentially melted together to form the component. More specifically, fine metal powder layers are sequentially melted after being uniformly distributed using an applicator on a metal powder bed. Each applicator includes an applicator element in the form of a lip, brush, blade or roller made of metal, plastic, ceramic, carbon fibers or rubber that spreads the metal powder evenly over the build platform. The metal powder bed can be moved in a vertical axis. The process takes place in a processing chamber having a precisely controlled atmosphere. Once each layer is created, each two dimensional slice of the component geometry can be fused by selectively melting the metal powder. The melting may be performed by a high powered melting beam, such as a <NUM> Watt ytterbium laser, to fully weld (melt) the metal powder to form a solid metal. The melting beam moves in the X-Y direction using scanning mirrors, and has an intensity sufficient to fully weld (melt) the metal powder to form a solid metal. The metal powder bed may be lowered for each subsequent two dimensional layer, and the process repeats until the component is completely formed. In order to create certain larger blades faster, some metal additive manufacturing systems employ a pair of high powered lasers that work together to form a blade. Here, a method of making turbine rotor blade <NUM> may include sequentially creating a layer of material and applying a heat source to sinter the layer of materials to form the structure described herein. Thus, additive manufacturing results in airfoil body <NUM>, tip end <NUM>, shank <NUM> and impingement cooling structure <NUM> including a plurality of integral material layers.

Turbine rotor blade <NUM> may be made of a metal which may include a pure metal or an alloy, capable of withstanding the environment in which employed. In one example, the metal may include practically any non-reactive metal powder, i.e., non-explosive or non-conductive powder, such as but not limited to: a cobalt chromium molybdenum (CoCrMo) alloy, stainless steel, an austenite nickel-chromium based alloy such as a nickel-chromium-molybdenum-niobium alloy (NiCrMoNb) (e.g., Inconel <NUM> or Inconel <NUM>), a nickel-chromium-iron-molybdenum alloy (NiCrFeMo) (e.g., Hastelloy® X available from Haynes International, Inc. ), or a nickel-chromium-cobalt-molybdenum alloy (NiCrCoMo) (e.g., Haynes <NUM> available from Haynes International, Inc. ), etc. In another example, the metal may include practically any metal such as but not limited to: tool steel (e.g., H13), titanium alloy (e.g., Ti<NUM>Al<NUM>V), stainless steel (e.g., <NUM>) cobalt-chrome alloy (e.g., CoCrMo), and aluminum alloy (e.g., AlSi<NUM>Mg).

In contrast to conventional impingement inserts, first end <NUM> of hollow body <NUM> is integrally formed to shank inner surface <NUM>, i.e., via additive manufacturing. Consequently, exterior surface <NUM> of hollow body <NUM> may be made to be uniformly spaced from inner surface(s) between first end <NUM> and second end <NUM> of hollow body <NUM>, regardless of the curvature of, for example, airfoil inner surface <NUM> and/or shank inner surface <NUM>. Further, hollow body <NUM> may have cooling passages <NUM> about an entirety of its periphery and radial span to provide impingement cooling throughout the blade, not just at a leading edge thereof. Hence, the integral impingement cooling structure <NUM> allows for maximum coverage of impingement with no sacrifices normally associated with impingement inserts, and can have a number of variable cooling features. For example, turbine rotor blade <NUM> may have: a variable chordwise width for the impingement cooling structure <NUM> (i.e., width WC1) or a pin bank <NUM> (i.e., WC4) aft of structure <NUM>, a tailored impingement cooling structure <NUM>, different wall thicknesses for different cooling and/or structural loads, and varied supports <NUM>, <NUM> to address different coefficients of thermal expansion (CTE) between airfoil body <NUM> and impingement cooling structure <NUM>. Additional cooling features, such as turbulators (not shown), may also be provided and customized around each cooling passage <NUM> to optimize impingement cooling. Turbine rotor blade <NUM> also may include axial venting through trailing edge <NUM>, as described relative to <FIG>.

Referring to <FIG>, another integral feature according to embodiments of the invention is illustrated. Similar to the previous embodiment, turbine rotor blade <NUM> may include airfoil body <NUM> with radially extending chamber <NUM> for receiving coolant flow <NUM>. As shown best in <FIG> and <FIG>, platform <NUM> extends laterally outward relative to airfoil body <NUM> and terminates at at least one slash face <NUM> (e.g., <FIG>, <FIG> shows a perspective, transparent view of a pressure side <NUM> of platform <NUM>, and <FIG> shows a top down, transparent view of a suction side <NUM> of platform <NUM>. As shown in <FIG> and <FIG>, a cooling circuit <NUM> is located within platform <NUM> and is in fluid communication with a source of coolant <NUM>. Source of coolant <NUM> may take any of a variety of forms. In one example, where turbine rotor blade <NUM> includes impingement cooling structure <NUM> in radially extending chamber <NUM>, source of coolant <NUM> to cooling circuit <NUM> may provide the coolant after passing through impingement cooling structure <NUM>, i.e., the coolant is post-impingement coolant. In another embodiment, source of coolant <NUM> may be radially extending chamber <NUM>. For example, where impingement cooling structure <NUM> is not provided, or it is provided radially outward of platform <NUM>, source of the coolant <NUM> to cooling circuit <NUM> may provide the coolant directly from radially extending chamber <NUM>. Other sources of coolant <NUM> may also be used, e.g., wheel space portion between shanks <NUM> of adjacent turbine rotor blades <NUM>. Cooling circuit <NUM> may take any now known or later developed form. In the example shown in <FIG>, cooling circuit <NUM> includes a sinusoidal path through platform <NUM>. In contrast, in <FIG>, cooling circuit <NUM> includes an elbow path. Cooling circuit <NUM> may have a less complex path or a more complex path, and may extend where necessary to cool platform <NUM>.

Turbine rotor blade <NUM> also includes cooling passage(s) <NUM> from cooling circuit <NUM> through a surface <NUM> of slash face(s) <NUM>, i.e., to cool slash faces <NUM> and other structure. Cooling passage(s) <NUM> are in platform <NUM> and in fluid communication with cooling circuit <NUM>. In contrast to conventional, linear cooling passages, cooling passage(s) <NUM> extend in a non-linear configuration from cooling circuit <NUM> to exit through at least one slash face <NUM> of platform <NUM>, providing improved cooling compared to linear cooling passages. In <FIG>, cooling passage(s) <NUM> have a (gently) curved shaped. Any number of cooling passage(s) <NUM> may be employed, to provide the desired cooling. Further, they may have any uniform or non-uniform cross-sectional shape desired, and may be uniformly or non-uniformly spaced, to provide the desired cooling. The non-linear configuration is made possible by, for example, additive manufacturing. As noted, airfoil body <NUM> and platform <NUM>, including the parts that define cooling passages(s) <NUM>, may include a plurality of integral material layers.

<FIG> shows an enlarged cross-sectional view of a slash face <NUM>. In some embodiments, as shown in <FIG>, slash face(s) <NUM> may include an extension member <NUM>. Extension member <NUM> may define a damper pin seat <NUM> configured to receive a damper pin <NUM> (shown in phantom) that seals with the damper pin seat of an adjacent turbine rotor blade (not shown). Where provided, cooling passage(s) <NUM> may extend through extension member <NUM>. In this regard, cooling passage(s) <NUM> may have a non-linear configuration that extends from cooling circuit <NUM> radially outward and about damper pin seat <NUM> to outer surface <NUM> of slash face <NUM>, i.e., in a curved shape that is sharper or more turning than <FIG>.

Cooling passage(s) <NUM> may take any of a number of non-linear configurations to provide the desired cooling. The non-linear configuration, e.g., curved shape, may extend in any desired direction within platform <NUM>, e.g., radially (inward or outward), axially (aft or forward) or circumferentially (clockwise or counterclockwise), or a combination of the directions. Cooling passage(s) <NUM> may all have the same shape to provide the same cooling attributes at each location where provided, or they may vary in shape within platform <NUM> to provide custom cooling for each location where they are provided. In addition to the curved shape shown in <FIG> and <FIG>, in an embodiment shown in <FIG>, cooling passage(s) <NUM> may have a helical (corkscrew) shape, i.e., with a number of helical coils <NUM>. Any number of helical coils <NUM> may be used for each cooling passage <NUM>. As shown in <FIG> and <FIG>, cooling passage(s) <NUM> may have at least one first turn <NUM> (<FIG>) in a first direction FD, and at least one second turn <NUM> (<FIG>) in a second, opposite direction SD, creating a generally zig-zag path. Any number of first and second turns <NUM>, <NUM> (<FIG>) may be used for each cooling passage <NUM>. As shown on the left side of <FIG>, an amplitude A of each turn <NUM>, <NUM> can be consistent so as to form a sinusoidal shape with the at least one first and second turns of equal amplitude A. Alternatively, as shown on the right side of <FIG>, the amplitude of each turn <NUM>, <NUM> can be inconsistent so as to form a more random zig-zag path with turns <NUM>, <NUM>. As also shown in the right side of <FIG>, an input <NUM> and an exit <NUM> of each cooling passage <NUM> need not be aligned. In another embodiment, shown in <FIG>, cooling passage(s) <NUM> may have a plurality of branches <NUM>, e.g., like a tree. Any branching configuration may be employed.

<FIG> shows an embodiment in which cooling passage(s) <NUM> have a curved shape, e.g., more planar within platform <NUM> than in <FIG>. <FIG> also illustrates exit <NUM> of cooling passage(s) <NUM> may meets slash face <NUM> of platform <NUM> at an angle α less than <NUM>°. Angle α may be customized to provide the desired cooling to platform <NUM> and/or film cooling to slash face <NUM>. While shown separately, any of the cooling passage examples or aspects thereof may be combined with the other examples.

Cooling circuit <NUM> and cooling passage(s) <NUM> may be provided in pressure side <NUM> of platform alone (<FIG> alone), in suction side <NUM> of platform <NUM> alone (<FIG> alone), or in both sides <NUM>, <NUM> of platform <NUM> (<FIG>). If provided on only one side of platform <NUM>, any other conventional structure may be provided in the other side of the platform. In the latter case, as shown collectively in <FIG> and <FIG>, cooling circuit <NUM> may include a first portion 234SS in suction side <NUM> of platform <NUM> and a second portion 234PS on pressure side <NUM> of platform <NUM>. Portions 234SS and 234PS may be separated or fluidly coupled. In this regard, slash face <NUM> includes a suction side slash face 230SS and a pressure side slash face 230PS, Here, cooling passage(s) <NUM> in platform <NUM> may include: at least one first cooling passage <NUM> in fluid communication with first portion 234SS of the cooling circuit and exiting suction side slash face 230SS, and at least one second cooling passage <NUM> in fluid communication with second portion 234PS of cooling circuit <NUM> and exiting pressure side slash face 230PS,.

Non-linear cooling passages <NUM> allow coolant to be directed where needed in platform <NUM>, in contrast to conventional, drilled linear coolant passages. The additive manufacture of coolant passages <NUM> allow them to have a wide variety of non-linear configurations that direct cooling where necessary and provide enhanced cooling through their shape.

Referring to <FIG> and <FIG>, another integral feature according to embodiments of the invention includes an angel wing <NUM> having a coolant transfer passage therein. <FIG> shows a radial cross-section through turbine rotor blade <NUM> including an angel wing <NUM>, <FIG> shows a transparent perspective view of turbine rotor blade <NUM> including angel wing <NUM>, <FIG> shows an axial view of a set of turbine rotor blades 120A-C including angel wing(s) <NUM>, and <FIG> shows a top down view of a turbine rotor blade <NUM> including angel wings <NUM>. With further respect to <FIG>, a set of turbine rotor blades includes: a first turbine rotor blade 120A, a second turbine rotor blade 120B and a third turbine rotor blade 120C (collectively or individually, turbine rotor blade <NUM>). First turbine rotor blade 120A is positioned between second and third turbine rotor blades 120B, 120C. In this embodiment, as shown in <FIG>, each turbine rotor blade <NUM> may include airfoil body <NUM> including concave pressure side outer wall <NUM> (<FIG>) and convex suction side outer wall <NUM> (<FIG>) that connect along leading and trailing edges <NUM>, <NUM> (<FIG>). Turbine rotor blade <NUM> may also include shank <NUM> at radial inner end <NUM> of airfoil body <NUM>. In addition, turbine rotor blade <NUM> includes at least one angel wing <NUM> extending laterally from at least one side <NUM>, <NUM> of shank <NUM>.

As shown for one nozzle-blade interface in <FIG>, an opening <NUM> exists at the interface between adjacent nozzles <NUM> and turbine rotor blades <NUM> that can allow hot working fluid to exit the hot gas path and enter a wheel space <NUM> of turbine <NUM>. In order to limit this leakage of hot gas, turbine rotor blade <NUM> typically includes axially projecting angel wing seals <NUM>, also simply referred to as 'angel wings'. Angel wings <NUM> cooperate with projecting segments or 'discouragers' <NUM> which extend from nozzle <NUM>. Angel wings <NUM> and discouragers <NUM> overlap (or nearly overlap), but do not touch each other, thus restricting fluid flow.

Turning to <FIG>, in accordance with embodiments of the disclosure, turbine rotor blade <NUM> may also include a coolant transfer passage <NUM> defined through the at least one angel wing <NUM>. Coolant transfer passage <NUM>, e.g., for a first turbine rotor blade 120A (<FIG>), fluidly couples a first wheel space portion <NUM> defined between shank 148A (<FIG>) and a first adjacent shank 148B (<FIG>) of a first adjacent turbine rotor blade 120B and a second wheel space portion <NUM> defined between shank 148A (<FIG>) and a second adjacent shank 148C of a second adjacent turbine rotor blade 120C (<FIG>). As shown best by observing <FIG>, <FIG> and <FIG>, each wheel space portion <NUM>, <NUM> is part of wheel space <NUM>. Wheel space <NUM> is defined: circumferentially between shanks 148A-C (<FIG>) of adjacent turbine rotor blades 120A-C, axially between shank <NUM> and an adjacent nozzle <NUM>, and radially by platform <NUM> and rotor disks <NUM>. Wheel space portions <NUM>, <NUM> are that part of wheel space <NUM> that is axially beside a particular blade's shank <NUM>.

As shown best in <FIG>, coolant transfer passage <NUM> includes a first open end <NUM> in fluid communication with first wheel space portion <NUM> and a second open end <NUM> in fluid communication with second wheel space portion <NUM>. Thus, coolant transfer passage <NUM> allows a wheel space coolant <NUM> to pass between wheel space portions <NUM>, <NUM> on circumferentially opposing sides of shank <NUM>. First open end <NUM> and second open end <NUM> may face in an axial-circumferential direction relative to airfoil body <NUM>, or any direction that will allow wheel space coolant <NUM> to pass between wheel space portions <NUM>, <NUM>. Wheel space coolant <NUM> may be any now known or later developed coolant, e.g., routed from compressor <NUM> (<FIG>). As noted previously, outer walls <NUM>, <NUM> of airfoil body <NUM> define radially extending chamber <NUM> that may extend into shank <NUM>. As shown in <FIG>, coolant transfer passage <NUM> is fluidly isolated from radially extending chamber <NUM>, i.e., coolant flow <NUM> (<FIG>) from chamber <NUM> does not mix with wheel space coolant <NUM>.

Any number of angel wings <NUM> may be employed. In one example, shown in <FIG>, <FIG> and <FIG>, a first angel wing <NUM> extends laterally from first side <NUM> of shank <NUM>, and a second angel wing <NUM> extends laterally from a second, opposing side <NUM> of shank <NUM>. In another example, shown in <FIG>, a first pair of radially spaced angel wings <NUM> may extend laterally from first side <NUM> of shank <NUM>, and none extend from side <NUM> of shank <NUM>. In another embodiment, as shown in <FIG>, a first pair of radially spaced angel wings <NUM> may extend laterally from first side <NUM> of shank <NUM>, and a second pair of radially spaced angel wings <NUM> may extend laterally from a second, opposing side <NUM> of shank <NUM>. In any event, each angel wing <NUM> may include a respective coolant transfer passage <NUM>. Alternatively, while each angel wing <NUM> is shown including a coolant transfer passage <NUM>, selective angel wings may not include a coolant transfer passage.

Coolant transfer passage <NUM> allows wheel space coolant <NUM> (<FIG>) to move between wheel space portions <NUM>, <NUM>, allows cooling of angel wings <NUM>, and reduces the weight turbine rotor blade <NUM>.

Referring to <FIG>, <FIG>, <FIG>, another integral feature according to embodiments of the invention includes a hollow blade mount <NUM>. In this embodiment, root <NUM> is provided including shank <NUM> having radially extending chamber <NUM> defined therein. Blade mount <NUM> is at a radial inner end of shank <NUM>. In contrast to many conventional blade mounts, blade mount <NUM> has a hollow interior <NUM> defined therein, e.g., by inner wall surfaces <NUM> of blade mount <NUM>. Hollow interior <NUM> is in fluid communication with radially extending chamber <NUM>. Hollow interior <NUM> may have any desired interior shape, e.g., expanding radially as shown in <FIG>. Blade mount <NUM> may have any now known or later developed exterior shape configured for mounting to a rotor wheel <NUM> (<FIG>) coupled to rotor <NUM> (<FIG>), e.g., a dovetail, or fir tree shape.

Turbine rotor blade root <NUM> may further include a lattice support structure <NUM> disposed within hollow interior <NUM> of the blade mount <NUM>. Lattice support structure <NUM> may take a variety of hollow support structure forms. In one example, lattice support structure <NUM> may include a plurality of radially extending V-shaped sections <NUM>. V-shaped sections <NUM> may be integral with inner wall surfaces <NUM> of blade mount <NUM>. Root <NUM> including shank <NUM> and blade mount <NUM>, including lattice support structure <NUM>, may be made by additive manufacture. Shank <NUM> and blade mount <NUM> thus may include a plurality of integral material layers.

Root <NUM> according to this embodiment, i.e., with lattice support structure <NUM>, may also include platform <NUM>, as described herein relative to <FIG>. Platform <NUM>, as noted, is positioned radially outward of shank <NUM> and extends laterally outward relative to the shank, terminating at at least one slash face <NUM>. Platform <NUM> may include cooling circuit <NUM> defined within the platform and in fluid communication with a source of a coolant flow, e.g., radially extending chamber <NUM>. Cooling passage(s) <NUM> (<FIG>) may be defined in platform <NUM> and in fluid communication with cooling circuit <NUM>. As noted, cooling passage(s) <NUM> extending in a non-linear configuration from cooling circuit <NUM> to exit through slash face(s) <NUM> of the platform. Slash face(s) <NUM> may include extension member <NUM> through which cooling passage(s) <NUM> extend. Cooling passage(s) <NUM> may have: a helical shape (<FIG>); at least one first turn in a first direction, and at least one second turn in a second, opposite direction (<FIG>); a plurality of branches (<FIG>); or a curved shape (e.g., <FIG>, <FIG>, <FIG>).

Root <NUM> according to this embodiment, i.e., with lattice support structure <NUM>, may also include angel wing(s) <NUM> extending laterally from at least one side of shank <NUM>, as described herein relative to <FIG>. As noted, a coolant transfer passage <NUM> may be defined through angel wing(s) <NUM>. As shown in <FIG>, coolant transfer passage <NUM> fluidly couples a first wheel space portion <NUM> defined between shank 148A and a first adjacent shank 148B of a first adjacent turbine rotor blade root 144B and a second wheel space portion <NUM> defined between shank 148A and a second adjacent shank 148C of a second adjacent turbine rotor blade root 144C. Coolant transfer passage <NUM> includes a first open end <NUM> in fluid communication with first wheel space portion <NUM> and a second open end <NUM> in fluid communication with second wheel space portion <NUM>. As shown in <FIG>, first open end <NUM> and second open end <NUM> may face in a circumferential direction relative to shank <NUM>. Coolant transfer passage <NUM> may be fluidly isolated from radially extending chamber <NUM> in shank <NUM>.

Root <NUM> according to this embodiment, i.e., with lattice support structure <NUM>, may also include both platform <NUM> and angel wing(s) <NUM>, as described herein. Additive manufacture allows for formation of root <NUM> with shank <NUM>, hollow blade mount <NUM>, lattice support structure <NUM>, and platform <NUM> and/or angel wing(s) <NUM>, creating a plurality of integral material layers for whatever features are provided.

Root <NUM> including integral lattice support structure <NUM> in hollow interior <NUM> of blade mount <NUM> provides a lighter turbine rotor blade <NUM>, and additional cooling of blade mount <NUM>.

While the various embodiments have been described and illustrated herein as used together, it is understood that the various embodiments can be used alone or in a combination.

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.

Claim 1:
A turbine rotor blade (<NUM>), comprising:
an airfoil body (<NUM>) including a concave pressure side outer wall (<NUM>) and a convex suction side outer wall (<NUM>) that connect along leading and trailing edges (<NUM>, <NUM>), the outer walls (<NUM>, <NUM>) defining a radially extending chamber (<NUM>) for receiving a coolant (<NUM>) flow;
a platform (<NUM>) extending laterally outward relative to the airfoil body (<NUM>) and terminating at at least one slash face (<NUM>, 230PS, 230SS);
a cooling circuit (<NUM>) defined within the platform (<NUM>) and in fluid communication with a source of the coolant (<NUM>) flow; and
at least one cooling passage (<NUM>, <NUM>) defined in the platform (<NUM>) and in fluid communication with the cooling circuit (<NUM>), the at least one cooling passage (<NUM>, <NUM>) extending in a non-linear configuration from the cooling circuit (<NUM>) to exit (<NUM>) through the at least one slash face (<NUM>, 230PS, 230SS) of the platform (<NUM>); characterized in that the at least one cooling passage (<NUM>, <NUM>) has a helical shape.