Internal cooling configurations in turbine rotor blades

Abstract

Rotor blade (16) for use in a turbine (12) of a gas turbine (10), comprising: a blade (25) formed between a concave pressure surface (26) and a laterally opposing convex suction surface (27), wherein the pressure surface (26) and the suction surface (27) axially between a front and a rear edge (28, 29) which are opposite each other, extend and extend radially between an outer tip (31) and an inner end connected to a root (21) which is configured to connect the rotor blade (16) to a rotor disk; a tip cover band (41) connected to the outer tip (31) of the airfoil (25), wherein the tip cover band (41) has an axially and circumferentially extending substantially planar component which is supported by the outer tip (31) of the airfoil (25); a cooling configuration (36) comprising multiple cooling channels (37) for receiving and directing a coolant through an interior of the rotor blade (16), wherein the cooling channels (37) each have fluidically connected successive segments, wherein: a feed segment (72) extends radially through the blade (25); an outlet segment (74) discharges coolant from the rotor blade (16) at a shallow angle relative to a flow direction of a working fluid through the turbine (12), each of the outlet segments (74) extending through an inner throat section (49) formed between the pressure surface (26) of the blade (25) and the tip cover (41) or between the suction surface (27) of the blade (25) and the tip cover (41); and a knee segment (73) connects the inlet segment (72) with the outlet segment (74) and is positioned adjacent to the outer tip (31) of the blade (25), wherein the knee segment (73) is configured to include a change of direction between the inlet segment (72) and the outlet segment (74) of between 80 and 100 degrees.

Description

BACKGROUND OF THE INVENTION

[0001] The present application relates to internal cooling channels and internal cooling configurations of blades in gas turbines. More specifically, but not by way of limitation, the present application relates to internal cooling channels and structural configurations formed near the outer radial tip of turbine rotor blades.

[0002] It will be understood that internal combustion or gas turbine engines ("gas turbines") comprise a compressor section and a turbine section in which rows of blades are arranged axially in stages. Each stage includes a row of circumferentially spaced stator blades that are stationary and a row of rotor blades that rotate about a central turbine axis or turbine shaft. In operation, the compressor rotor blades rotate generally about the shaft and, acting in concert with the stator blades, compress a flow of air. The supply of compressed air is then used in a combustion chamber to combust a supply of fuel. The resulting flow of hot, expanding gases from combustion, i.e., the working fluid, is expanded by the turbine section of the engine.The flow of working fluid through the turbine is diverted by the stator blades onto the rotor blades to cause rotation. The rotor blades are connected to a central shaft so that the rotation of the rotor blades rotates the shaft. In this way, the energy contained in the fuel is converted into the mechanical energy of the rotating shaft, which can be used, for example, to rotate the rotor blades of the compressor to create the supply of compressed air needed for combustion, and the coils of a generator to produce electrical power. During operation, due to the extreme temperatures of the hot gas path, the velocity of the working fluid, and the rotation speed of the drive, the blades within the turbine are subjected to extreme mechanical and thermal stresses.

[0003] Developing efficient and economical gas turbines is an ongoing and significant goal. While some strategies for increasing gas turbine efficiency are known, it remains a challenging goal because such alternatives—which include, for example, increasing the size of the engine, raising temperatures throughout the hot gas path, and increasing rotor blade rotation speeds—generally impose additional loads on the blades and other hot gas path component parts that are already subject to high stress. As a result, improved devices, methods, and systems that reduce the operating loads imposed on the turbine blades or that enable the turbine blades to better withstand these loads, allowing the engines to operate more efficiently, remain a significant area for technical improvement.

[0004] One strategy to mitigate the stresses on the blades is to actively cool them during operation. Such cooling can allow the blades to better withstand higher firing temperatures and mechanical stresses, which can extend blade life and generally make the engine more economical and efficient to operate. One way to cool the blades during operation is through the use of internal cooling ducts or circuits. Generally, this involves passing a relatively cool supply of compressed air, exhausted from the compressor, through internal cooling ducts. For a number of reasons that will become apparent, great care is required in the design and manufacture of these internal cooling ducts.

[0005] First, the use of cooling air reduces the efficiency of the engine. In particular, air from the compressor diverted for cooling purposes is air that could otherwise be used in the combustion process. As a result, the use of such air necessarily reduces the air availability for combustion and thereby reduces overall efficiency. This makes it imperative that the cooling channels be highly efficient so that the use of air for cooling is minimized. Second, newer turbine blade designs call for aggressively shaped aerodynamic configurations that are thinner and more curved or serpentine. These new blade configurations impose an additional requirement for compact and efficient channels. The new designs also create spatial constraints that hinder or limit the manufacture of traditional cooling channel configurations using conventional approaches.Third, internal cooling channels must be designed to favor lightweight rotor blades while still providing sufficiently robust structures to withstand extreme loads. This means that while cooling channel design is an effective way to reduce overall blade weight—promoting efficiency and reducing mechanical loads—the blades must still remain highly resilient. Cooling channels must therefore be engineered to remove both material and weight while still favoring structural resilience. Internal arrangements must also avoid stress concentrations or insufficiently cooled areas (or "hot spots"), which can negatively impact component life. Fourth, cooling configurations must also be designed so that discharged coolant enables surface cooling and efficient, aerodynamic operation.In particular, because cooling channels typically discharge coolant into the working fluid flowpath after circulating through the internal cooling channels, another design consideration concerns the use of discharged coolant for surface cooling and to minimize the associated aerodynamic losses. Discharged coolant is often relied upon to provide cooling to external surfaces or regions of the blades after its release, and this must be coordinated with internal cooling strategies and take aerodynamic performance into account.

[0006] US 2,687,278 A describes a turbine rotor blade having an airfoil extending axially between a leading edge and a trailing edge and radially between a root portion and an outer tip, and a cooling configuration including a plurality of cooling channels for receiving and conducting a coolant through the interior of the rotor blade. The cooling channels each extend from the root portion through the airfoil, with some of the cooling channels terminating at the outer tip of the airfoil and other cooling channels terminating at the trailing edge of the airfoil to form curved passages.

[0007] DE 698 11 624 T2 describes a rotor blade of a turbine having an airfoil, a tip shroud at the outer tip of the airfoil and an internal cooling configuration having internal cooling channels extending through the airfoil and cooling holes drilled through the tip shroud and connected to the internal cooling channels to receive coolant therefrom and to conduct it through the tip shroud to cool the tip shroud and to discharge the coolant at opposite suction and pressure side edges of the tip shroud.

[0008] US 2007 / 0 065 283 A1 describes a rotor blade of a turbine with internal cooling channels which run through the blade and exit radially outwards through a tip shroud or branch off into cooling through-holes which exit outwards through a surface of the blade below an inner throat section between the blade and the tip shroud.

[0009] As will be understood, the design of internal cooling configurations within turbine blades according to these and other criteria presents many complex, often competing requirements. New designs that balance these in a manner that optimizes or improves one or more desired performance criteria—while still sufficiently promoting structural robustness, component longevity, economical propulsion, and efficient use of coolant—represent significant technological advantages.

[0010] Based on this, it is an object of the invention to provide a rotor blade for use in a turbine of a gas turbine and a gas turbine having a plurality of such turbine rotor blades with internal cooling configurations within the rotor blades that reduce one or more of the above-mentioned deficiencies and improve performance criteria, in particular enabling an improvement in the overall efficiency and performance of the cooling of the rotor blades.

[0011] To achieve this object, the invention provides a rotor blade for use in a turbine of a gas turbine having the features of patent claim 1 and a gas turbine having the features of patent claim 10. Particularly preferred embodiments of the invention are defined in the dependent claims. BRIEF DESCRIPTION OF THE INVENTION

[0012] According to a first aspect of the invention, a rotor blade for use in a gas turbine turbine is provided, comprising an airfoil formed between a concave pressure surface and a laterally opposed convex suction surface, a tip shroud, and a cooling configuration including a plurality of cooling channels for receiving and directing a coolant through an interior of the rotor blade. The pressure surface and the suction surface extend axially between opposite leading and trailing edges and radially between an outer tip and an inner end connected to a root configured to connect the rotor blade to a rotor disk. The tip shroud is connected to the outer tip of the airfoil and includes an axially and circumferentially extending substantially planar component supported by the outer tip of the airfoil.The cooling channels each comprise fluidly connected successive segments, wherein: a supply segment extends radially through the airfoil; an outlet segment that discharges coolant from the rotor blades at a shallow angle relative to the flow direction of a working fluid through the turbine, each of the outlet segments extending through an inner throat section formed between the pressure surface of the airfoil and the tip shroud or between the suction surface of the airfoil and the tip shroud; and an elbow segment connecting the supply segment and the outlet segment and located near the outer tip of the airfoil. The elbow segment is configured to include a change in direction between the supply segment and the outlet segment of between 80 and 100 degrees.

[0013] In any embodiment of the turbine rotor blade, it may be advantageous for the outlet segment to extend between an upstream end connected to the knee segment and a downstream end at which an outlet port is formed on an outer surface of the rotor blade. It is also advantageous for the outlet port of each of the cooling channels to have a location adjacent to and inboard of the tip shroud. It is also advantageous for a rear half of the airfoil to be formed as an airfoil section that is rearward of an axial center of the airfoil.

[0014] In any embodiment of the turbine rotor blade, it may be advantageous that the outlet port of each of the cooling channels is formed by a target surface area of ​​the rotor blade. It is also advantageous that the knee segment of each of the cooling channels is arranged to extend through a target interior area of ​​the rotor blade. It is also advantageous that the interior fillet area is configured to form a smooth transition between a surface of the airfoil and an interior surface of the tip shroud, wherein the interior fillet area comprises a pressure surface interior fillet area and a suction surface interior fillet portion, corresponding respectively to the pressure surface and the suction surface of the airfoil.

[0015] In any embodiment of the turbine rotor blade, it may be advantageous for the target surface area of ​​the outlet segment to comprise the pressure surface of the airfoil adjacent to the outer tip of the airfoil. It is also advantageous for the target interior area to comprise an interior area of ​​the airfoil adjacent to the suction surface of the airfoil.

[0016] In any embodiment of the turbine rotor blade, it may be advantageous for the target surface area of ​​the outlet segment to comprise a rear half of the suction surface of the airfoil adjacent to the outer tip of the airfoil. It is also advantageous for the target interior area to comprise an interior area of ​​the airfoil adjacent to the rear half of the suction surface of the airfoil.

[0017] In any embodiment of the turbine rotor blade, it may be advantageous for the target surface area of ​​the outlet segment to have a surface area corresponding to the suction surface inner fillet area. It is also advantageous for the target inner area to have the suction surface inner fillet section.

[0018] In any embodiment of the turbine rotor blade, it may be advantageous for the target surface area of ​​the outlet segment to have a rear half of a surface area corresponding to the suction surface inner throat area. It is also advantageous for the target inner area to have a rear half of the suction surface inner throat section.

[0019] In any embodiment of the turbine rotor blade, it may be advantageous for the target surface area of ​​the outlet segment to comprise: a rear half of the suction surface of the airfoil adjacent to the outer tip of the airfoil and / or a rear half of the surface area corresponding to the suction surface inner fillet area. It is also advantageous for the target interior area to comprise: an interior area of ​​the airfoil adjacent to the rear half of the suction surface of the airfoil adjacent to the outer tip of the airfoil and / or a rear half of the suction surface inner fillet section.

[0020] In any embodiment of the turbine rotor blade, it may be advantageous for the supply segment of each of the cooling channels to have an upstream end formed through the root of the rotor blade, at which the supply segment is fluidly connected to a coolant source, and a downstream end fluidly connected to the knee section; and wherein the supply segment has a radially directed linear passage formed through the airfoil.

[0021] In any embodiment of the turbine rotor blade, it may be advantageous that, relative to an expected flow direction of the coolant within each of the cooling channels, the supply segment is upstream relative to the knee segment and the knee segment is upstream relative to the outlet segment.

[0022] In any embodiment of the turbine rotor blade, it may be advantageous for the change of direction through the knee segment to be 90 degrees.

[0023] In any embodiment of the turbine rotor blade, it may be advantageous for a radius of curvature along a length of the knee segment to be at least greater than twice a diameter of a cross-sectional flow area through the knee segment.

[0024] In any embodiment of the turbine rotor blade, it may be advantageous for the supply segments of the cooling channels to extend approximately parallel to each other in a radial direction through the airfoil and to be grouped within the rear half of the airfoil. It is also advantageous for the outlet ports to comprise a row formed along the rear half of the airfoil's suction surface adjacent to the outer tip of the airfoil and / or the rear half of the surface area corresponding to the suction surface inner fillet area.

[0025] In any embodiment of the turbine rotor blade, it may be advantageous for the supply segments of the cooling channels to be regularly spaced along the camber line of the blade. It is also advantageous for the row of outlet ports to be approximately parallel to the tip shroud.

[0026] In any embodiment of the turbine rotor blade, it may be advantageous that a discharge direction for each of the outlet segments comprises a direction in which a fluid discharged therefrom flows. It is also advantageous that an axial reference line comprises a reference line that is parallel to a central axis of the turbine and points in the direction of flow of the working fluid through the turbine. It is also advantageous that a discharge angle is the reference angle formed between the discharge direction of each of the outlet segments and the axial reference line, wherein the discharge angle comprises a tangential discharge angle describing an axial inclination between the discharge direction and the axial reference line, and a radial discharge angle describing a radial inclination between the discharge direction and the axial reference line.It is also advantageous that for each of the outlet segments: the tangential discharge angle has a range of between + / -30 degrees; and / or the radial discharge angle has a range of between + / -30 degrees.

[0027] In any embodiment of the turbine rotor blade, it may be advantageous that for each of the outlet segments: the tangential discharge angle has a range of between + / -10 degrees; and / or the radial discharge angle has a range of between + / -10 degrees.

[0028] In any embodiment of the turbine rotor blade, it may be advantageous for a discharge direction for each of the outlet segments to comprise a direction in which a fluid discharged therefrom flows. It is also advantageous for an axial reference line to comprise a reference line that is parallel to a central axis of the turbine and points to the flow direction of the working fluid through the turbine. It is also advantageous for the discharge direction for each of the outlet segments to be substantially parallel to the axial reference line.

[0029] Another aspect of the invention relates to a gas turbine having a row of rotor blades in a turbine, the row of rotor blades comprising a plurality of rotor blades, each comprising: an airfoil formed between a concave pressure surface and a laterally opposite convex suction surface, the pressure surface and the suction surface extending axially between oppositely extending leading and trailing edges and radially between an outer tip and an inner end connected to a root configured to connect the rotor blade to a rotor disk; a tip shroud connected to the outer tip of the airfoil, the tip shroud comprising an axially and circumferentially extending substantially planar component carried by the outer tip of the airfoil;a cooling configuration comprising a plurality of cooling channels for receiving and directing coolant through an interior of the rotor blade, the cooling channels each comprising fluidly connected, successive segments wherein: a supply segment extends radially through the airfoil; an outlet segment delivering coolant from the rotor blade substantially parallel to a flow direction of a working fluid through the turbine, each of the outlet segments extending through an inner fillet region formed between the pressure surface of the airfoil and the tip shroud or between the suction surface of the airfoil and the tip shroud;and an elbow segment connecting the feed segment and the outlet segment and disposed adjacent the outer tip of the airfoil, the elbow segment configured to include a change of direction between the feed segment and the outlet segment of between 80 and 100 degrees;

[0030] These and other features of the present application will become apparent upon studying the following detailed description of the preferred embodiments when taken in conjunction with the drawings and the appended claims. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] These and other features of the invention will be more fully understood and appreciated by carefully studying the following more detailed description of exemplary embodiments of the invention taken in conjunction with the accompanying drawings in which: Fig. 1 is a schematic representation of an exemplary gas turbine that may include turbine blades according to aspects and embodiments of the present application; Fig. 2 a sectional view of the compressor section of the gas turbine Fig. 1 is; Fig. 3 a sectional view of the turbine section of the gas turbine from Fig. 1 is; Fig. 4 is a side view of an exemplary turbine rotor blade that may have an internal cooling configuration and structural arrangement according to aspects and embodiments of the present application; Fig. 5 a sectional view along the line of sight 5-5 Fig. 4 is; Fig. 6 a sectional view along the line of sight 6-6 from Fig. 4 is; Fig. 7 a sectional view along the line of sight 7-7 Fig. 4 is; Fig. 8 is a perspective view of an exemplary turbine rotor blade that may include a tip shroud and configuration according to aspects and embodiments of the present application; Fig. 9 is an enlarged perspective view of the tip cover band of Fig. 8 is; Fig. 10 is an external view of an installed assembly of turbine rotor blades that may include tip shrouds and configurations according to aspects and embodiments of the present application; Fig. 11 is an external profile of a tip cover band that may have a configuration according to aspects and embodiments of the present application; Fig. 12 is a perspective view with partial transparency of a tip shroud and an airfoil having an internal cooling configuration according to embodiments of the present application; Fig. 13 a perspective of the outlet connections of the cooling configuration Fig. 12 is; Fig. 14 is an external view with partial transparency of a tip shroud and an airfoil having an internal cooling configuration according to aspects and embodiments of the present application; and Fig. 15 is a side view with partial transparency of a tip shroud and an airfoil having an internal cooling configuration according to embodiments of the present application. DETAILED DESCRIPTION OF THE INVENTION

[0032] Aspects and advantages of the present application will be set forth below in the following description, or may be obvious from the description, or may be learned by practice of the invention. Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical reference characters to refer to features in the drawings. Like or similar reference characters in the drawings and in the description may be used to refer to like or similar parts of the embodiments of the invention. As will be understood, each example is provided for the purpose of illustrating the invention and not for limiting the invention.Indeed, it will be apparent to those skilled in the art that modifications and variations can be made to the present invention without departing from the scope or spirit thereof. For example, features illustrated or described as part of one embodiment may be used in another embodiment to result in yet another embodiment. It is intended that the present invention encompass such modifications and variations that are within the scope of the appended claims and their equivalents. It is to be understood that ranges and limits mentioned herein include all sub-ranges that fall within the given limits, including the limits themselves, unless otherwise indicated. In addition, certain terms have been chosen to explain the present invention and its component subsystems and parts.Where possible, these terms were chosen based on terminology commonly used in the technical field. It is understood that such terms are often still subject to different interpretations. For example, what is referred to herein as a single component may be referred to elsewhere as consisting of multiple components, or what is referred to herein as having multiple components may be referred to elsewhere as a single component.In understanding the scope of the present invention, attention should be paid not only to the specific terminology used, but also to the accompanying description and context, as well as to the structure, configuration, function, and / or use of the identified and described component, including how the term refers to the various figures, and, of course, to the precise use of terminology in the appended claims. Furthermore, while the following examples are presented with respect to particular types of gas turbines or turbine engines, the technology of the present application may be applicable to other categories of turbine engines without limitation, as understood by a person of ordinary skill in the relevant technology field.Accordingly, it should be understood that unless otherwise indicated, the use of the term "gas turbine" herein is intended to be broad and limited to the applicability of the present invention to the various types of turbine engines.

[0033] Given the nature of gas turbines' operation, some terms have proven particularly useful in describing certain aspects of their function. These terms and their definitions, unless expressly stated otherwise, are as follows. As will be understood, such terms can be used to describe or claim the gas turbine or any of its primary subsystems—i.e., the compressor, combustor, or turbine—as well as to describe or claim components or subcomponents for use therein. In the latter case, the terminology should be understood as describing these components as being correctly installed and / or functioning within the gas turbine engine or more primary subsystems.

[0034] Accordingly, the terms "front" and "rear" refer to directions relative to the orientation of the gas turbine, and more specifically, the relative positioning of the compressor section and turbine section of the engine. Therefore, as used herein, the term "front" refers to the compressor end, while the term "rear" refers to the turbine end. It will be understood that each of these terms may be used to indicate the direction of movement or relative position along the central axis of the engine. As stated above, these terms may be used to describe attributes of the gas turbine or any of its primary subsystems, as well as for components or subcomponents arranged therein. Therefore, if, for example,When a component, such as a rotor blade, is described or claimed as having a "leading surface," it may be understood as referring to a surface oriented in the leading direction as defined by the orientation of the gas turbine (i.e., the positioning of the combustor and turbine subsystems). Unless otherwise specified, this assumption applies to the following descriptive terms as well.

[0035] The terms "downstream" and "upstream" are used herein to indicate the position within a specified conduit or flow path relative to the direction of flow (hereinafter "flow direction") flowing therethrough. Therefore, the term "downstream" refers to the direction in which the fluid flows through the specified conduit, while "upstream" refers to the direction opposite thereto. These terms may be interpreted as referring to what is understood by those skilled in the art to be the direction of flow through the conduit, assuming normal or expected operation. As will be understood, the working fluid within the compressor section and the turbine section of the gas turbine is directed downstream through an annularly shaped working fluid flow path typically formed about the central and common axis of the gas turbine.As such, the term "flow direction" as used herein within the compressor section and the turbine section of the engine refers to a reference direction that represents an idealized or generalized direction of flow of the working fluid through the working fluid flow path of the engine during an expected or normal operating condition. Therefore, the "flow direction" terminology within the compressor and turbine sections refers to flow that is parallel to the central axis of the gas turbine and oriented in a downstream or aft direction.

[0036] Accordingly, the flow of working fluid through the working fluid flow path of the gas turbine can be described as beginning as pressurized air through the compressor for the flow direction, becoming combustion gases in the combustor after being burned with a fuel, and finally expanding in the flow direction as they pass through the turbine. Similarly, the flow of working fluid can be described as beginning at a forward or upstream location toward a forward or upstream end of the gas turbine, moving generally in a downstream or aft direction, and finally ending at an aft or downstream location toward a rear or downstream end of the gas turbine.

[0037] Since many gas turbine components rotate during operation, such as compressor and turbine rotor blades, the terms "rotationally forward" and "rotationally aft" may be used to describe relative positioning of subcomponents or subregions in the expected rotational direction within the engine. Therefore, as will be understood, these terms may distinguish positions with respect to the direction of rotation (hereinafter "rotational direction") within the compressor or turbine. As used herein, such a rotational direction may be understood as the expected direction of rotation of a component, assuming normal or expected operation of the gas turbine.

[0038] Furthermore, given the configuration of gas turbines, particularly the arrangement of the compressor and turbine sections around a common shaft or rotor, as well as the cylindrical configuration common to many combustor types, terms describing position relative to an axis may regularly be used herein. In this context, it will be understood that the term "radial" refers to movement or position perpendicular to an axis. Associated with this, it may be necessary to describe a relative distance from the central axis. In such cases, if a first component is arranged closer to the central axis than a second component, the first component will be described as being "radially inward" or "inboard" of the second component.On the other hand, if the first component is located farther from the central axis, the first component will be described as being either "radially outward" or "outward" of the second component. As used herein, the term "axial" refers to movement or position parallel to an axis, while the term "circumferential" refers to movement or position about the axis. Unless otherwise stated or clearly obvious from the context, these terms should be construed as referring to the central axis of the compressor section and / or the turbine section of the gas turbine, as defined by the rotor extending through each of them, even if the terms describe or claim attributes of non-integral components—such as rotor or stator blades—operating therein.Unless otherwise specified, these terms may be used with reference to the longitudinal axis of specific components or subsystems within the gas turbine, such as the longitudinal axis around which a conventional cylindrical or "tube" combustor is typically arranged.

[0039] The term "rotor blade" refers, without further specification, to the rotating blades of either the compressor or the turbine, and can thus include both compressor rotor blades and turbine rotor blades. The term "stator blade" refers, without further specification, to the stationary blades of either the compressor or the turbine, and can thus include both compressor stator blades and turbine stator blades. The term "airfoil" can be used to refer generally to any type of airfoil. Therefore, without further specification, the term "airfoil" is inclusive of all types of turbine blades, including compressor rotor blades, compressor stator blades, turbine rotor blades, turbine stator blades, and the like.

[0040] As background, now referring specifically to the figures, the Fig. 1-3 illustrate an exemplary gas turbine in accordance with the present invention or within which the present invention may be used. It will be understood by those skilled in the art that the present invention is not limited to this type of use. As stated, the present invention may be used in gas turbines, such as engines used in power generation and in aircraft, in steam turbine engines, as well as other types of rotating engines, as understood by one of ordinary skill in the art. The examples provided are therefore not intended to be limiting unless otherwise stated. Fig. Figure 1 is a schematic diagram of a gas turbine 10. Generally, gas turbines operate by extracting energy from a pressurized flow of hot gas produced by the combustion of a fuel in a flow of compressed air. As described in Fig. 1, the gas turbine 10 may be configured with an axial compressor 11 mechanically connected by a common shaft or rotor to a downstream turbine section or turbine 12, and a combustion chamber 13 disposed between the compressor 11 and the turbine 12. As shown in Fig. 1, the gas turbine may be arranged around a common central axis 19.

[0041] Fig. Figure 2 illustrates a view of an exemplary multi-stage axial compressor 11 used in the gas turbine of Fig. 1. As illustrated, the compressor 11 may include a plurality of stages, each including a row of compressor rotor blades 14 and a row of compressor stator blades 15. Thus, a first stage may include a row of compressor rotor blades 14 rotating about a central shaft, followed by a row of compressor stator blades 15 that remain stationary during operation. Fig. 3 illustrates a partial view of an exemplary turbine section or turbine 12 used in the gas turbine of Fig. 1 can be used. The turbine 12 may also have a plurality of stages. Three example stages are illustrated, but more or fewer stages may be present. Each stage may include a plurality of turbine nozzles or stator vanes 17 that remain stationary during operation, followed by a plurality of turbine blades or rotor vanes 16 that rotate about the shaft during operation. The turbine stator vanes 17 are generally circumferentially spaced from one another and are attached to an outer casing about the axis of rotation. The turbine rotor vanes 16 may be mounted to a turbine wheel or rotor disk (not shown) for rotation about a central axis. It will be understood that the turbine stator vanes 17 and the turbine rotor vanes 16 lie in the hot gas path or working fluid flow path through the turbine 12.The direction of flow of combustion gases or working fluid within the working fluid flow path is indicated by the arrow.

[0042] In one example of the operation of the gas turbine 10, the rotation of the compressor rotor blades 14 within the axial compressor 11 can compress a flow of air. In the combustion chamber 13, energy can be released when the compressed air is mixed with a fuel and ignited. The resulting flow of hot gases or working fluid from the combustion chamber 13 is then directed over the turbine rotor blades 16, which triggers the rotation of the turbine rotor blades 16 about the shaft. In this way, the energy of the flow of the working fluid is converted into mechanical energy of the rotating blades and, provided the connection between the rotor blades and the shaft, the rotating shaft. The mechanical energy of the shaft can then be used to drive the rotation of the compressor rotor blades 14, thus generating the necessary supply of compressed air, and also, for example,to drive a generator to produce electricity.

[0043] For background information purposes, the Fig. 4-7 are views of a turbine rotor blade 16 in accordance with, or within which, aspects of the present invention may be practiced. As will be understood, these figures are provided to illustrate conventional configurations of rotor blades and describe spatial relationships between components and regions within such airfoils for future reference, while also describing the geometric constraints and other criteria affecting the internal and external design thereof. While the airfoil of this example is a rotor blade, it will be understood that, unless otherwise indicated, the present invention may be applied to other types of airfoils within the gas turbine.As stated above, the description of such components may include terms that derive their meaning based on the orientation and function of the gas turbine, and more specifically the working fluid flow path, and therefore should be understood in that context, i.e., such description assumes that the rotor blade is correctly installed and operating under expected or normal conditions within the engine.

[0044] The rotor blade 16, as illustrated, may include a root 21 configured for attachment to a rotor disk. The root 21 may, for example, include a dovetail 22 configured for mounting in a corresponding dovetail slot in the periphery of a rotor disk. The root 21 may also include a shank 23 extending between the dovetail 22 and a platform 24. The platform 24 generally, as illustrated, forms the connection between the root 21 and the airfoil 25, which is the active component of the rotor blade 16 that interrupts the flow of working fluid through the turbine 12 and has the desired rotation. The platform 24 may form the inner end of the airfoil 25. The platform may also form a portion of the inner boundary of the working fluid flow path through the turbine 12.

[0045] The airfoil 25 of the rotor blade may typically have a concave pressure surface 26 and a circumferentially or laterally opposed convex suction surface 27. The pressure surface 26 and the suction surface 27 may extend axially between respective opposite leading and trailing edges 28, 29 and in the radial direction between an inner end, which may be formed at the junction with the platform 24, and an outer tip 31. The airfoil 25 may have a curved or contoured shape designed to promote a desired aerodynamic performance. As described in the Fig. 4 and Fig. 5, the shape of the airfoil 25 may taper successively as it extends between the platform 24 and the outer tip 31. The taper may include an axial taper that narrows the distance between the leading edge 28 and the trailing edge 29 of the airfoil 25, as shown in Fig. 4, as well as a circumferential taper that reduces the thickness of the blade 25 as defined between the suction surface 27 and the pressure surface 26, as shown in Fig. 5. As shown in the Fig. 6 and Fig. 7, the contoured shape of the airfoil 25 may also include a twist about the longitudinal axis of the airfoil 25 as it extends away from the platform 24. It will be understood that the twist may be present to successively vary a stagger angle for the airfoil 25 between the inner end and the outer tip 31.

[0046] For descriptive purposes, as in Fig. 4, the airfoil 25 of the rotor blade 16 may also be described as having a leading edge region or leading half and a trailing edge region or trailing half formed on either side by an axial centerline 32. The axial centerline 32, as used herein, may be formed by connecting the centers 34 of the camber lines 35 of the airfoil 25 between the platform 24 and the outer tip 31. Furthermore, the airfoil 25 may be described as having two radially stepped sections formed inboard and outboard of a radial centerline 33 of the airfoil 25.Therefore, as used herein, an inner portion or half of the airfoil 25 extends between the platform 24 and the radial centerline 33, while an outer portion or half extends between the radial centerline 33 and the outer tip 31. Finally, the airfoil 25 may be described as having a pressure surface portion or half and a suction surface portion or half, which, as will be understood, are formed on either side of the camber line 35 of the airfoil 25 and the respective sides 26, 27 of the airfoil 25.

[0047] The rotor blade 16 may also include an internal cooling configuration 36 with one or more cooling channels 37 through which a coolant circulates during operation. Such cooling channels 37 may extend radially outward from a connection to a supply source formed through the root 21 of the rotor blade 16. The cooling channels 37 may be linear, curved, or a combination thereof and may include one or more outlets or surface ports through which the coolant is discharged from the rotor blade 16 and into the working fluid flow path.

[0048] The Fig. 8-11 illustrate a turbine rotor blade 16 with a tip shroud 41 in accordance with the present invention or within which the present invention may be used. As will be understood, Fig. 8 is a perspective view of an exemplary turbine rotor blade 16 having a tip shroud 41, while Fig. 9 is an enlarged view of the tip cover band portion 41. Fig. Figure 10 provides a profile from an external perspective of an exemplary installed arrangement of rotor blades 16 provided with a tip cover band. Finally, Fig. 11 provides an enlarged external view of a tip cover band 41, which can be used to indicate the various regions within tip cover bands in the discussion below.

[0049] As shown, the tip shroud 41 may be positioned adjacent or at the outer end of the airfoil 25. The tip shroud 41 may comprise an axially and circumferentially extending flat plate or flat component supported by the airfoil 25 toward its center. For descriptive purposes, the tip shroud 41 may include an inner surface 45, an outer surface 44, and an edge 46. As illustrated, the inner surface 45 is opposite the outer surface 44 with respect to the narrow radial thickness of the tip shroud 41, while the edge 46 connects the inner surface 45 to the outer surface 44 and forms, as used herein, a circumferential or outer profile of the tip shroud 41.

[0050] A sealing rail 42 may be disposed along the outer surface 44 of the tip shroud 41. Generally, as illustrated, the sealing rail 42 is a fin-like projection extending radially outward from the outer surface 44 of the tip shroud 41. The sealing rail 42 may extend circumferentially between opposite ends of the tip shroud 41 in a direction of rotation, or "rotational direction," of the rotor blade 16. As will be understood, the sealing rail 42 may be used to prevent leakage of the working fluid through the radial gap typically present between the tip shroud 41 and the surrounding stationary components that form the outer boundary of the working fluid flow path through the turbine. According to conventional designs, the sealing rail 42 may extend radially into an abradable stationary honeycomb seal opposed to it through this gap.The sealing rail 42 may extend substantially the entire circumferential length of the outer surface 44 of the tip shroud 41. As used herein, the circumferential length of the tip shroud 41 is the length of the tip shroud 41 in the rotational direction 50. For descriptive purposes, the sealing rail 42 may have opposing rail surfaces, with a forward surface 56 corresponding to the forward direction of the gas turbine and an aft surface 57 corresponding to the aft direction. As will be understood, the forward surface 56 therefore faces toward or in the flow direction of the working fluid, while the aft surface 57 faces away therefrom. Both the forward surface 56 and the aft surface 57 of the sealing rail 42 may be arranged to form a steep angle with respect to the outer surface 44 of the tip shroud 41.

[0051] Although other configurations are possible, the sealing rail 42 may have a roughly rectangular profile. The front side 56 and the rear side 57 of the sealing rail 42 may be joined along narrow peripheral edges, which, as used herein, include opposite and substantially parallel outer and inner edges, and opposite and substantially parallel rotationally leading and rotationally trailing edges. In particular, the inner edge of the sealing rail 42 may be formed as an interface between the sealing rail 42 and the outer surface 44 of the tip cover band 41. As will be understood, the inner edge is somewhat obscured due to fillet areas formed for structural purposes between the sealing rail 42 and the tip cover band 41 and, therefore, is not specifically designated by a numerical identification.The outer edge 59 of the seal rail 42 is radially offset from the outer surface 44 of the tip shroud 41. This radial offset generally represents the radial height of the seal rail 42, as will be understood. As indicated, a rotationally leading edge 62 of the seal rail 42 projects radially from the edge 46 of the tip shroud 41 that overhangs the suction surface 27 of the airfoil 25. Therefore, the rotationally leading edge 62 is the component that "guides" the seal rail 42 as the rotor blade 16 rotates during operation. At the opposite end of the seal rail 42, a rotationally trailing edge 63 projects from the edge 46 of the tip shroud 41 that overhangs the pressure surface 26 of the airfoil 25. Given this arrangement, the trailing edge 63 in the direction of rotation is the component that “follows” the sealing rail 42 when the rotor blade 16 rotates during operation.

[0052] A cutting tooth 43 may be provided on the sealing rail 42. As will be understood, the cutting tooth 43 may be provided to cut a groove in the abradable coating or honeycomb of the stationary sealing band that is slightly wider than the width of the sealing rail 42. As will be understood, the honeycomb may be provided to improve sealing stability, and the use of the cutting tooth 43 may reduce spillage and rubbing between stationary and rotating parts by clearing the wider path. The cutting tooth 43 is generally an area of ​​increased width along the circumferential length of the sealing rail 42. More specifically, the cutting tooth 43 may comprise a circumferential portion of the sealing rail 42 that is axially flared. This axially flared area may extend radially between the outer surface 44 of the tip cover band 41 and the outer edge of the sealing rail 42.The cutting tooth 43 may be disposed adjacent the center or central portion of the sealing rail 42. As provided below, the cutting tooth 43 may be disposed within the airfoil portion of the outer surface 44 of the tip shroud 41. The cutting tooth 43 may have a substantially rectangular profile, although other profiles are also possible.

[0053] The tip cover band 41 may include fillet portions 48, 49 configured to provide smooth surface transitions between the diverging surfaces of the tip cover band 41 and the airfoil 25, as well as between the tip cover band 41 and the seal rail 42. As such, configurations of the tip cover band 41 may include an outer fillet portion 48 formed between the outer surface 44 of the tip cover band 41 and the leading surface 56 and the trailing surface 57 of the seal rail 42. The tip cover band 41 may also include an inner fillet portion 49 formed between the inner surface 45 of the tip cover band 41 and the pressure and suction surfaces 26, 27 of the airfoil 25.It will be understood that the inner fillet section 49 may be more accurately described as including: a pressure inner fillet section, which is the section formed between the pressure surface 26 of the airfoil 25 and the inner surface 45 of the tip shroud 41; and a suction inner fillet section, which is the section formed between the suction surface 27 of the airfoil 25 and the inner surface 45 of the tip shroud 41. The outer fillet region 48 may be more accurately described as including: a forward outer fillet section, which is the section formed between the forward side 56 of the seal rail 42 and the outer surface 44 of the tip shroud 41; and an aft outer fillet section, which is the section formed between the aft surface 57 of the sight rail 42 and the outer surface 44 of the tip shroud 41.As illustrated, each of these fillet regions 49, 48 can be configured to provide smooth curved transitions between the multiple planar surfaces forming abrupt or steeply angled transitions. It will be appreciated that such fillet sections can improve aerodynamic performance as well as distributed stress concentrations that would otherwise occur in these regions. Even so, these regions remain highly stressed due to the overhanging or cantilevered loading of the tip shroud 41 and the speed of the drive. As will be understood, without adequate cooling, the stresses in these regions are a significant limitation to the useful life of the component.

[0054] With particular reference now to the Fig. 10, tip shrouds may be configured to have a contact interface in which contact surfaces or contact edges, during operation, engage corresponding surfaces or edges formed on tip shrouds 41 of adjacent rotor blades. As will be understood, this may be done, for example, to reduce leakage or harmful vibrations. Fig. 10 provides an external view of tip shrouds 41 on turbine rotor blades as they may appear in an assembled condition. As indicated, for descriptive purposes, the edge 46 of the tip shroud 41 may include a rotationally leading contact edge 52 and a rotationally trailing contact edge 53 relative to the rotational direction 50. Therefore, as shown, the tip shroud 41 in a rotationally leading position may be configured to be formed with a rotationally trailing contact edge 53 that contacts or comes near the rotationally leading contact edge 52 of the tip shroud 41 located in a rotationally trailing position relative thereto.While this area of ​​contact between adjacent tip cover bands 41 may generally be referred to as a contact interface, given the profile of the exemplary configuration, it may also be referred to as a "Z-groove" interface. Other configurations are also possible. By forming the contact interface, the edge 46 of the tip cover band may be configured with a grooved portion designed to cooperatively contact or bias an adjacent and similarly configured tip cover band 41 in a predetermined manner.

[0055] Now with particular reference to the Fig. 11, the outer profile of the tip shroud 41 may have a scalloped shape. Although other configurations are possible, the exemplary scalloped shape is one that performs well in reducing leakage while also minimizing weight. Whatever the profile, it will be understood that regions or sections forming the tip shroud 41 may be described in terms of their position relative to the seal rail 42 and / or the profile of the underlying airfoil 25. Therefore, as used herein, an airfoil section 65 of the tip shroud 41 is the portion formed within a profile of the airfoil 25 projected radially through the tip shroud 41.A compression portion 66 of the tip shroud 41 is the portion that projects away from and overhangs the compression surface 26 of the airfoil 25, while a suction portion 67 of the tip shroud 41 is the portion that projects away from and overhangs the suction surface 27 of the airfoil 25. Finally, a forward portion 68 of the tip shroud 41 is the portion that extends in a forward direction from the seal rail 42 to overhang the leading edge 28 of the airfoil 25, while an aft portion 69 of the tip shroud 41 is the portion that extends in a rearward direction from the seal rail 42 to overhang the trailing edge 29 of the airfoil 25.

[0056] Referring now to the Fig. 12-15 illustrate several internal cooling configurations consistent with aspects and exemplary embodiments of the present invention. As will be understood, these examples are described with reference to and in light of systems, components, and related concepts previously provided herein, particularly those discussed with reference to the preceding figures.

[0057] As illustrated, the present invention may include internal cooling configurations for gas turbine rotor blades. As will be appreciated, such cooling configurations may include one or more internal cooling passages for receiving and directing a coolant, such as compressed air bled from the compressor, through the interior of the rotor blades. According to the present configurations, an internal cooling configuration 36 is provided having one or more of the cooling passages formed within the rotor blade, each of which may include fluidly connected and consecutive sections or segments extending through the root, airfoil, and / or portions of the tip shroud. As provided below and with reference to the various figures, these segments may include a supply segment 72, an elbow segment 73, and an outlet segment 74.

[0058] As illustrated, the supply segment 72 may form the upstream end of the cooling passage and may therefore extend through the root 21 of the rotor blade 16 to connect the cooling configuration 36 to a coolant source. The cooling configuration 36 may include a plurality of cooling passages, each of which may include a discrete and separate supply segment 72 extending radially outward through the airfoil 25. According to exemplary configurations, as shown in FIGS. Fig. 12-15, each of the supply segments 72 may be configured as a radially directed coolant passage formed through the interior of the airfoil 25. The supply segment 72 may be linear, although other configurations are also possible. At a downstream end, each of the supply segments 72 may be fluidly connected to one of the elbow segments 73.

[0059] The elbow segment 73, according to the present configurations, is the portion of the cooling channel through which an abrupt change in direction is achieved. The elbow segment 73, as illustrated, may connect the supply segment 72 to the outlet segment 74 while containing the change in direction that occurs therebetween. As illustrated, this change in direction may be abrupt or sharp. According to exemplary embodiments, the change in direction that occurs through the elbow segment 73 may include an angle of approximately 90 degrees. More generally, according to other embodiments, the change in direction through the elbow segment 73 may have an angle within a range of values. In particular, the elbow segment is configured to effect a change in direction of between 80 and 100 degrees. By implementing this change in direction, the elbow segment 73 may be configured such that the curvature is sharp rather than gradual.As will be understood, the sharp curvature of the knee segment 73 can be described by relating the radius of curvature along the length of the knee segment 73 to the diameter of the cross-sectional flow area of ​​the knee segment 73. According to preferred embodiments, the radius of curvature is at least greater than twice the diameter of the cross-sectional flow area.

[0060] As will be explained, the knee segment 73 may be configured to be located within a target interior region of the rotor blade 16. The target interior region may be one that is to be convectively cooled by the coolant flowing through the knee segment 73. As will be understood, the change in direction by the knee segment 73 may induce turbulent flow of the coolant flow, which may increase efficiency. According to certain preferred embodiments, the target interior region within which the knee segment 73 is located comprises either the interior region of the airfoil 25 adjacent to the tip shroud 41 or the interior fillet section 49 between the airfoil 25 and the tip shroud 41. According to other preferred embodiments, the target interior region is the suction surface interior fillet section 49.In certain preferred embodiments, the target interior region is limited to the rear half of the airfoil 25. Even more preferably, the target interior region comprises the rear half of the suction surface inner throat section 49.

[0061] The outlet segment 74, as shown in the Fig. 12-15 is the portion of the cooling channel located between the elbow segment 73 and the outlet port 75 formed on the outer surface of the rotor blade 16. Therefore, the outlet segment 74 is the portion through which coolant flowing within the cooling channels is discharged into the working fluid flow path. The outlet segment 74 may have an upstream end connected to the elbow segment 73 and a downstream end formed as the exit port 75. As illustrated, the outlet segment 74 may extend between the elbow segment 73 and the outlet port 75 along a linear path, although other configurations are also possible. Within any one of the cooling channels, the cross-sectional flow area of ​​the outlet segment 74 and the elbow segment 73, as well as the supply segment 72, may be approximately the same, although other configurations are also possible.As will be understood, the outlet segment 74 of the cooling passage of the present invention enables fluidic communication with the working fluid flow path through the turbine 12 because the outlet port 75 is formed on an outer surface of the rotor blade 16.

[0062] According to preferred embodiments, the outlet ports 75 of the outlet segments 74 may be formed at predetermined target surface areas. The target surface areas may generally be any of several surface areas of the airfoil 25 and / or the tip shroud 41 previously discussed herein. Therefore, in accordance with exemplary embodiments, the target surface areas for the cooling configuration 36 may include one or more of the following: the pressure surface 26 of the airfoil 25; the suction surface 27 of the airfoil 25; the surface area corresponding to the inner fillet region 49; the surface area corresponding to the outer fillet region 48; the inner surface 45 of the tip shroud 41; and the edge 46 of the tip shroud 41.According to preferred embodiments, the outlet ports 75 are formed just inboard of the tip shroud 41 and on the suction surface 27 of the airfoil 25. As illustrated, the preferred location may also include an outer surface region of the inboard throat section 49 and / or an outer region of the suction surface 27 of the airfoil 25 (i.e., the portion of the suction surface 27 adjacent to the tip shroud 41). As illustrated, according to preferred embodiments, the outlet ports 75 may be arrayed in a row along the aft portion of the inboard throat section 49 of the suction surface 27 of the airfoil 25. The row of outlet ports 75 may be immediately inboard of and approximately parallel to the tip shroud 41.

[0063] According to other preferred embodiments, the knee segment 73 may be shaped to bisect or extend through preselected target inner sections. These target inner sections may be the outer section of the airfoil 25 adjacent to or in contact with the tip shroud 41 and the inner fillet section 49. That is, according to preferred embodiments, the knee segment is formed immediately inward of the tip shroud 41. As will be understood from the positioning at this point, the target inner section may include the suction fillet inner section 49 and / or the outer inner section of the suction surface 27 of the airfoil 25.

[0064] The cooling configurations 36 of the present invention may include a plurality of cooling channels extending through the rotor blade 16. These may be centered within the airfoil 25 or may be offset from the suction surface 27 of the airfoil 25. The cooling channels may also be grouped in the rear axial half of the airfoil 25, which, as used herein, is rearward of the axial centerline 32 of the airfoil 25. According to preferred embodiments, a plurality of feed segments 72 may be present, and as illustrated, these may extend approximately parallel relative to one another in a radial direction through the airfoil 25. As shown, these feed segments 72 may be positioned within the rear half of the airfoil 25. According to a preferred embodiment, the feed segments 72 may be regularly spaced along the bend line 35 of the airfoil 25.In addition, the cross-sectional flow areas of the supply segments 72 as well as those of the connected knee and outlet segments 73, 74 may decrease with closer positioning of the cooling channels to the trailing edge 29 of the airfoil 25. It will be understood that the narrowing among the cooling channels may proportionally match the narrowing of the airfoil 25 as it extends to the trailing edge 29.

[0065] As indicated, the supply segments 72 may extend radially from each of the cooling passages to a position in the airfoil 25 that is adjacent or in close proximity to the tip shroud 41 or the outer radial end of the airfoil 25. Each of the supply segments 72 may be connected to a corresponding one of the knee segments 73, which bends the flow direction through the cooling passage by approximately 90 degrees, as indicated. As illustrated, the knee segment 73 may be configured to cradle the cooling passage so that the flow direction through the outlet segments 74 approximately aligns with the flow direction of the working fluid through the turbine 12. It will be understood that the flow direction through the turbine 12 is generally perpendicular relative to the longitudinal axis of the seal rail 42 and oriented such that the pressure surface 26 is the upstream surface of the airfoil 25.Each cooling channel of the elbow segment 73 may have an outlet segment 74 connecting it to one of the outlet ports 75. According to preferred embodiments, the outlet port 75 may be formed on the suction surface 27 of the airfoil 25 or the suction surface inner fillet portion 49 between the airfoil 25 and the tip shroud 41. As best shown in FIG. Fig. 13, the cooling configuration 36 may include the outlet ports 75 arranged as a spaced row along the rear half of the suction surface 27.

[0066] The outlet ports 75, as indicated, may serve to discharge coolant flowing through each of the cooling passages. According to conventional designs, radially oriented cooling passages are typically configured to extend through the tip shroud to connect to outlet ports located on the outer surface of the tip shroud. With such an arrangement, coolant is discharged from the outlet ports in a direction substantially perpendicular to the direction of flow through the turbine. As will be understood, discharging coolant in this manner causes significant aerodynamic mixing losses. That is, the discharged coolant generally interferes with and disrupts the flow of working fluid through the turbine flow path, adversely impacting aerodynamic performance and ultimately engine efficiency.

[0067] According to embodiments of the present invention, the included elbow segments 73 lie or curve each of the cooling passages as described, such that the coolant discharged therefrom is discharged at a shallow or small angle relative to the direction of flow of the working fluid through the turbine 12, or, in accordance with certain preferred embodiments, parallel thereto. That is, the elbow segment 73 may be configured according to the preferred embodiments such that the outlet segment 24 extending therefrom does so in a direction closely aligned with, or within a preferred range of angles, the direction of flow through the turbine 12.

[0068] To describe this angular relationship more precisely, it may be useful to introduce several terms (including "discharge direction," "axial reference line," and "discharge angle") and define how these terms are used herein. As stated in the Fig. 14 and Fig. 15, the discharge direction 79 represents the direction in which the outlet segment 74 is aimed to discharge the coolant exiting therefrom. The discharge direction 79 is therefore the direction in which a fluid discharged from the outlet segment 74 would flow if not otherwise acted upon. The discharge direction 79 may also be understood as being approximated by a continuation of a reference line, this reference line being defined by the central longitudinal axis of the outlet segment 74. The axial reference line 80, as used herein, is a reference line oriented according to and in the general direction of flow of the working fluid through the turbine 13. The axial reference line 80 may therefore be assumed to be substantially parallel to the central axis 19 through the turbine 12 and oriented in the downstream direction.Finally, the discharge angle 81, 82, as used herein, is the reference angle formed between the discharge direction 79 of one of the outlet segments 74 and the axial reference line 80. As will be understood, the discharge angle 81, 82 can be defined both tangentially and radially. That is, the discharge angle 81, 82 can describe the degree to which the discharge direction 79 is inclined tangentially or radially with respect to the axial reference line 80. Therefore, as shown in FIG. Fig. 14, a tangential discharge angle 81 describes a tangential inclination or angle between the discharge direction 79 and the axial reference line 80, while, as shown in Fig. 15, a radial discharge angle 82 describes a radial inclination or angle that occurs between the discharge direction 79 and the axial reference line 80.

[0069] The present cooling configurations 36 may be configured, as indicated, so that the discharge angles 81, 82 are relatively flat. As will be understood, the particular size of the discharge angles 81, 82 may be varied, as the optimal angle value or range may depend on various factors, such as the configuration of the tip shroud or airfoil, or other design criteria, such as the rotor blade cooling requirement or strategy. Furthermore, it should be understood that while the cooling performance of the present configurations may be optimized or improved at specific values ​​or within certain preferred ranges, the performance improvements may be achieved over a wide range of values ​​for the discharge angles 81, 82.Having said that, applicants have identified some preferred embodiments that are particularly advantageous when used in conjunction with known rotor blade configurations, and these embodiments are now disclosed. With respect to the tangential discharge angle θ1 as shown in FIG. Fig. 14, exemplary configurations of the outlet segment 74 have a tangential discharge angle 81 that is within a range of between about + / -30 degrees relative to the axial reference line 80. According to certain preferred embodiments, the outlet segments 74 are configured such that the tangential discharge angle 81 is within a range of about + / -10 degrees relative to the axial reference line 80. Still other preferred embodiments include outlet segments 74 having a tangential discharge angle 81 that approaches 0 degrees, i.e., the discharge direction 79 is approximately parallel to the axial reference line 80. With respect to the radial discharge angle 82 - as shown in Fig.15, exemplary configurations of the outlet segments 74 include a radial discharge angle 82 that is within a range of approximately + / - 30 degrees relative to the axial reference line 80. According to certain preferred embodiments, the outlet segments 74 are configured such that the radial discharge angle 82 is within a range of approximately + / - 10 degrees relative to the axial reference line 80. Still other preferred embodiments include outlet segments 74 with radial discharge angles 82 that approach 0 degrees, i.e., the discharge direction 79 is approximately parallel to the axial reference line 80.

[0070] As will be understood, dispensing coolant at such shallow angles relative to the downstream axial direction, as described herein, can reduce aerodynamic mixing losses. This results from the fact that the flow disruption of the working fluid is minimized. Furthermore, dispensing coolant in the manners described herein can reduce the strength of secondary vortices by more gently mixing the dispensed coolant with the flow of the working fluid. This can improve overall efficiency and performance. In addition, the dispensing angles 81, 82 and the proposed location of the cooling channels and the outlet ports 75 associated therewith (i.e.through and at airfoil regions, with respect to aft portions of the suction surface 27 and near the tip shroud 41 and the suction surface inner fillet region 49) accommodate cooling requirements in a region that is difficult to cool, while also improving or reducing the overall torsional forces acting on that region of the airfoil 25.

[0071] The invention is defined by the following claims.

[0072] A turbine rotor blade having an airfoil formed between a concave pressure surface and a laterally opposed convex suction surface, and a cooling configuration including a cooling passage for receiving and directing coolant through an interior of the rotor blade. The cooling passage may include fluidly connected segments wherein: a supply segment extends radially through the airfoil; an outlet segment that discharges coolant from the rotor blade at a shallow angle relative to the direction of flow of a flowing fluid through the turbine; and an elbow segment connecting the supply segment and the outlet segment and positioned adjacent the outer tip of the airfoil. The elbow segment may be configured to include a change in direction between the supply segment and the outlet segment.

Claims

[1] Rotor blade (16) for use in a turbine (12) of a gas turbine (10), comprising: a blade (25) formed between a concave pressure surface (26) and a laterally opposite convex suction surface (27), wherein the pressure surface (26) and the suction surface (27) axially between a front and a rear edge (28, 29) which are opposite to each other, extend and extend radially between an outer tip (31) and an inner end connected to a root (21) adapted to connect the rotor blade (16) to a rotor disk; a tip cover band (41) connected to the outer tip (31) of the airfoil (25), the tip cover band (41) having an axially and circumferentially extending substantially planar component carried by the outer tip (31) of the airfoil (25); a cooling configuration (36) comprising a plurality of cooling channels (37) for receiving and directing a coolant through an interior of the rotor blade (16), the cooling channels (37) each comprising fluidly connected successive segments in which: a feed segment (72) extends radially through the blade (25); an outlet segment (74) discharges coolant from the rotor blade (16) at a shallow angle relative to a flow direction of a working fluid through the turbine (12), each of the outlet segments (74) extending through an inner throat portion (49) formed between the pressure surface (26) of the airfoil (25) and the tip shroud (41) or between the suction surface (27) of the airfoil (25) and the tip shroud (41); and a knee segment (73) connects the feed segment (72) to the outlet segment (74) and is positioned adjacent to the outer tip (31) of the airfoil (25), the knee segment (73) being configured to include a change in direction between the feed segment (72) and the outlet segment (74) of between 80 and 100 degrees. [2] The rotor blade (16) of claim 1, wherein the outlet segment (74) extends between an upstream end connected to the knee segment (73) and a downstream end at which an outlet port (75) is formed on an outer surface of the rotor blade (16); wherein the outlet port (75) of each of the cooling channels (37) has a location adjacent to and inward of the tip cover band (41); and wherein a rear half of the airfoil (25) is defined as an airfoil section that is rearward of an axial center (32) of the airfoil (25). [3] Rotor blade (16) according to claim 2, wherein: the outlet port (75) of each of the cooling channels (37) is formed by a target surface area of ​​the rotor blade (16); and the knee segment (73) of each of the cooling channels (37) is arranged to extend through a target interior region of the rotor blade (16); wherein the inner fillet portion (49) is configured to smoothly transition between a surface of the airfoil (25) and an inner surface of the tip shroud (41), wherein the inner fillet portion (49) has both a pressure surface inner fillet portion and a suction surface inner fillet portion corresponding to the pressure surface (26) and the suction surface (27) of the airfoil (25). [4] Rotor blade (16) according to claim 3, wherein - the target surface area of ​​the outlet segments (74) comprises the suction surface (27) of the airfoil (25) adjacent to the outer tip (31) of the airfoil (25), and wherein the target interior area comprises an interior area of ​​the airfoil (25) adjacent to the suction surface (27) of the airfoil (25); or - wherein the target surface area of ​​the outlet segments (74) comprises a rear half of the suction surface (27) of the airfoil (25) adjacent to the outer tip (31) of the airfoil (25), and wherein the target interior area comprises an interior area of ​​the airfoil (25) adjacent to the rear half of the suction surface (27) of the airfoil (25); or - wherein the target surface area of ​​the outlet segments (74) has a surface area corresponding to the suction surface inner throat section, and wherein the target inner area has the suction surface inner throat section; or - wherein the target surface area of ​​the outlet segments (74) has a rear half of a surface area corresponding to the suction surface inner throat section, and wherein the target inner area has a rear half of the suction surface inner throat section. [5] The rotor blade (16) of claim 3 or 4, wherein the target surface area of ​​the outlet segments (74) comprises: a rear half of the suction surface (27) of the airfoil (25) adjacent to the outer tip (31) of the airfoil (25); and / or a rear half of a surface area corresponding to the suction surface inner fillet portion; and wherein the target interior area comprises: an interior area of ​​the airfoil (25) adjacent to the rear half of the suction surface (27) of the airfoil (25) adjacent to the outer tip (31) of the airfoil (25); and / or a rear half of the suction surface inner fillet portion. [6] Rotor blade (16) according to claim 5, wherein the supply segment (72) of each of the cooling channels (37) has an upstream end passing through the root (21) of the rotor blade (16) is formed, at which the supply segment (72) is fluidly connected to a coolant source, and has a downstream end fluidly connected to the elbow segment (73); and wherein the feed segment (72) has a radially aligned linear passage formed through the airfoil (25). [7] Rotor blade (16) according to claim 6, wherein, relative to an expected flow direction of the coolant within each of the cooling channels (37), the supply segment (72) is arranged upstream relative to the knee segment (73) and the knee segment (73) is arranged upstream relative to the outlet segment (74). [8] Rotor blade (16) according to one of claims 3 to 7, wherein the supply segments (72) of the cooling channels (37) extend approximately parallel relative to one another in a radial direction through the blade (25) and are grouped within the rear half of the blade (25); and wherein the outlet ports (75) comprise a row formed along at least one of the following: the rear half of the suction surface (27) of the airfoil (25) adjacent to the outer tip (31) of the airfoil (25); and the rear half of the surface area corresponding to the suction surface inner throat section. [9] Rotor blade (16) according to any one of the preceding claims, wherein: a discharge direction for each of the outlet segments (74) comprises a direction in which a fluid discharged therefrom flows; an axial reference line (80) comprises a reference line that is parallel to a central axis (19) of the turbine (12) and points in the direction of flow of the working fluid through the turbine (12); a discharge angle (81, 82) is a reference angle formed between each discharge direction of the outlet segments (74) and the axial reference line (80), the discharge angle (81, 82) having a tangential discharge angle (81) indicating a tangential inclination between the discharge direction and the axial reference line (80) and a radial discharge angle (82) indicating a radial inclination between the discharge direction and the axial reference line (80); and where for each of the outlet segments (74): the tangential discharge angle (81) has a range of between + / -30 degrees; and the radial discharge angle (82) has a range between + / -30 degrees. [10] Gas turbine (10) comprising a row of rotor blades (16) in a turbine (12), wherein the row of rotor blades (16) comprises a plurality of rotor blades (16), each comprising: a blade (25) formed between a concave pressure surface (26) and a laterally opposite convex suction surface (27), wherein the pressure surface (26) and the suction surface (27) extends axially between a front and rear edge (28, 29) which are opposite to each other and extends radially between an outer tip (31) and an inner end which is connected to a root (21) which is adapted to connect the rotor blade (16) to a rotor disk; a tip cover band (41) connected to the outer tip (31) of the airfoil (25), the tip cover band (41) having an axially and circumferentially extending substantially planar component carried by the outer tip (31) of the airfoil (25); a cooling configuration (36) comprising a plurality of cooling channels (37) for receiving and directing a coolant through an interior of the rotor blade (16), the cooling channels (37) each comprising fluidly connected, successive segments in which: a feed segment (72) extends radially through the blade (25); an outlet segment (74) discharges the coolant from the rotor blade (16) approximately parallel to a flow direction of a working fluid through the turbine (12), each of the outlet segments (74) extending through an inner throat section (49) formed between the pressure surface (26) of the airfoil (25) and the tip shroud (41) or between the suction surface (27) of the airfoil (25) and the tip shroud (41); and a knee segment (73) connects the feed segment (72) and the outlet segment (74) and is arranged adjacent to the outer tip (31) of the airfoil (25), the knee segment (73) being adapted to include a change of direction between the feed segment (72) and the outlet segment (74) of between 80 and 100 degrees.

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