Patent Application
20170183971
The present application relates generally to apparatus, methods and/or systems concerning the design, manufacture, and use of rotor blades in combustion or gas turbine engines. More specifically, but not by way of limitation, the present application relates to apparatus and assemblies pertaining to turbine rotor blades having tip shrouds.
In combustion or gas turbine engines (hereinafter “gas turbines”), it is well known that air pressurized in a compressor is used to combust fuel in a combustor to generate a flow of hot combustion gases, whereupon the gases flow downstream through one or more turbines so that energy can be extracted therefrom. In accordance with such engines, generally, rows of circumferentially spaced rotor blades extend radially outwardly from a supporting rotor disc. Each rotor blade typically includes a dovetail that permits assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disc, as well as an airfoil that extends radially outwardly from the dovetail and interacts with the flow of the working fluid through the engine. The airfoil has a concave pressure side and convex suction side extending axially between corresponding leading and trailing edges, and radially between a root and a tip. It will be understood that the blade tip is spaced closely to a radially outer stationary surface for minimizing leakage therebetween of the combustion gases flowing downstream between the turbine blades.
Shrouds at the tip of the airfoil or “tip shrouds” often are implemented on aftward stages or rotor blades to provide a point of contact at the tip, manage bucket vibration frequencies, enable a damping source, and to reduce the over-tip leakage of the working fluid. Given the length of the rotor blades in the aftward stages, the damping function of the tip shrouds provides a significant benefit to durability. However, taking full advantage of the benefits is difficult considering the weight that the tip shroud adds to the assembly and the other design criteria, which include enduring thousands of hours of operation exposed to high temperatures and extreme mechanical loads. Thus, while large tip shrouds are desirable because of the effective manner in which they seal the gas path and the stable connections or interfaces they form between neighboring rotor blades, it will be appreciated that such shrouds are troublesome because of the increased pull loads on the rotor blade, particularly at the base of the airfoil because it must support the entire load of blade. That is to say, to the extent weight may be reduced while still fulfilling structural requirements, the life of the rotor blade may be extended.
As will be appreciated, according to these and other criteria, the design of tip shrouded rotor blades includes many complex, often competing considerations. Novel designs that balance these in a manner that optimizes or enhances one or more desired performance criteria—while still adequately promoting structural robustness, part-life longevity, component manufacturability, and/or cost-effective engine operation—represent economically valuable technology.
The present application thus describes a rotor blade for a gas turbine that includes an airfoil and a tip shroud having a cavitied configuration. The tip shroud may have a seal rail that projects radially from an outboard surface and extends circumferentially. The tip shroud may further include: a rotationally leading circumferential face; a rotationally trailing circumferential face; and an outboard face of the seal rail. The tip shroud may be circumferentially divided into three parallel reference zones that include: a rotationally leading edge zone, a rotationally trailing edge zone and, formed between and separating the rotationally leading edge zone and the rotationally trailing edge zone, a middle zone. The seal rail may include a hollow cavity wholly contained within at least one of the rotationally leading edge zone and the rotationally trailing edge zone. The cavity may include a mouth formed through at least one of the rotationally leading circumferential face, the rotationally trailing circumferential face, and the outboard face of the seal rail.
These and other features of the present application will become apparent upon review of the following detailed description of the preferred embodiments when taken in conjunction with the drawings and the appended claims.
These and other features of this invention will be more completely understood and appreciated by careful study of the following more detailed description of exemplary embodiments of the invention taken in conjunction with the accompanying drawings, in which:
Aspects and advantages of the present application are set forth below in the following description, or may be obvious from the description, or may be learned through 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 designations to refer to features in the drawings. Like or similar designations in the drawings and description may be used to refer to like or similar parts of embodiments of the invention. As will be appreciated, each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present invention without departing from the scope or spirit thereof. For instance, features illustrated or described as part of one embodiment may be used on another embodiment to yield a still further embodiment. It is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. It is to be understood that the ranges and limits mentioned herein include all sub-ranges located within the prescribed limits, inclusive of the limits themselves unless otherwise stated. Additionally, certain terms have been selected to describe the present invention and its component subsystems and parts. To the extent possible, these terms have been chosen based on the terminology common to the technology field. Still, it will be appreciated that such terms often are subject to differing interpretations. For example, what may be referred to herein as a single component, may be referenced elsewhere as consisting of multiple components, or, what may be referenced herein as including multiple components, may be referred to elsewhere as being a single component. In understanding the scope of the present invention, attention should not only be paid to the particular terminology used, but also to the accompanying description and context, as well as the structure, configuration, function, and/or usage of the component being referenced and described, including the manner in which the term relates to the several figures, as well as, of course, the precise usage of the terminology in the appended claims. Further, while the following examples are presented in relation to certain types of gas turbines or turbine engines, the technology of the present application also may be applicable to other categories of turbine engines, without limitation, as would the understood by a person of ordinary skill in the relevant technological arts. Accordingly, it should be understood that, unless otherwise stated, the usage herein of the term “gas turbine” is intended broadly and with limitation as the applicability of the present invention to the various types of turbine engines.
Given the nature of how gas turbines operate, several terms prove particularly useful in describing certain aspects of their function. These terms and their definitions, unless specifically stated otherwise, are as follows. The terms “forward” and “aft” or “aftward” refer to directions relative to the orientation of the gas turbine and, more specifically, the relative positioning of the compressor and turbine sections of the engine. Thus, as used therein, the term “forward” refers to the compressor end while “aft” or “aftward” refers to the turbine end. It will be appreciated that each of these terms may be used to indicate movement or relative position within the engine. The terms “downstream” and “upstream” are used herein to indicate position within a specified conduit relative to the general direction of fluid flowing through it. Thus, the term “downstream” refers to the direction in which the fluid is flowing through the specified conduit, while “upstream” refers to the direction opposite that. These terms may be construed as relating to what would be understood by one skilled in the art as the expected direction of flow through the conduit assuming normal or anticipated operation. Accordingly, for example, the primary flow of working fluid through a gas turbine, which begins as air moving through the compressor and then becomes combustion gases within the combustor for subsequent expansion through the turbine, may be described herein as beginning at a forward or upstream location toward a forward or upstream end of the gas turbine and terminating at an aft or downstream location toward an aft or downstream end of the gas turbine. Finally, as many components of gas turbines rotate during operation, such as compressor and turbine rotor blades, the terms rotationally lead and rotationally trail may be used to delineate included subcomponents or subregions. As will be appreciated, these terms differentiate position relative to a direction of rotation, which may be understood as being an expected direction of rotation given normal operation of the gas turbine.
In addition, given the configuration of the gas turbines, particularly the arrangement of the compressor and turbine sections about a common shaft or rotor, as well as the cylindrical configuration common to many combustor types, terms describing position relative to an axis may be regularly used herein. In this regard, it will be appreciated that the term “radial” refers to movement or position perpendicular to an axis. Related to this, it may be required to describe relative distance from the central axis. In such cases, for example, if a first component resides closer to the central axis than a second component, the first component will be described as being either “radially inward” or “inboard” of the second component. If, on the other hand, the first component resides further from the central axis, the first component will be described as being either “radially outward” or “outboard” 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 around an axis. Unless otherwise stated or contextually apparent, these terms describing position relative to an axis should be construed as relating to the central axis of the compressor and turbine sections of the engine as defined by the rotor extending through each. However, the terms also may be used relative to the longitudinal axis of certain components or subsystems within the gas turbine, such as, for example, the longitudinal axis around which conventional cylindrical or “can” combustors are typically arranged.
Finally, the term “rotor blade”, without further specificity, is a reference to the rotating blades of either the compressor or the turbine, and so may include both compressor rotor blades and turbine rotor blades. The term “stator blade”, without further specificity, is a reference to the stationary blades of either the compressor or the turbine and so may include both compressor stator blades and turbine stator blades. The term “blades” may be used to generally refer to either type of blade. Thus, without further specificity, the term “blades” is inclusive to all type of turbine engine blades, including compressor rotor blades, compressor stator blades, turbine rotor blades, turbine stator blades and the like.
By way of background, referring now to the figures,
In one example of operation for the gas turbine 10, the rotation of compressor rotor blades 14 within the axial compressor 11 may compress a flow of air. In the combustor 13, energy may be released when the compressed air is mixed with a fuel and ignited. The resulting flow of hot gases or working fluid from the combustor 13 is then directed over the turbine rotor blades 16, which induces the rotation of the turbine rotor blades 16 about the shaft. In this way, the energy of the flow of working fluid is transformed into the mechanical energy of the rotating blades and, given the connection between the rotor blades and the shaft, the rotating shaft. The mechanical energy of the shaft may then be used to drive the rotation of the compressor rotor blades 14, such that the necessary supply of compressed air is produced, and also, for example, a generator to produce electricity.
For background purposes,
The rotor blade 16, as illustrated, may include a root 21 that is used for attaching to a rotor disc. The root 21, for example, may include a dovetail 22 configured for mounting in a corresponding dovetail slot in the perimeter of a rotor disc. The root 21 may further include a shank 23 that extends between the dovetail 22 and a platform 24. The platform 24, as shown, forms the junction of the root 21 and an airfoil 25, which is the active component of the rotor blade 16 that intercepts the flow of working fluid through the turbine 12 and induces rotation. The platform 24 may define the inboard end of the airfoil 25 and a section of the inboard boundary of the working fluid flowpath through the turbine 12.
The airfoil 25 of the rotor blade may include a concave pressure face 26 and a circumferentially or laterally opposite convex suction face 27. The pressure face 26 and suction face 27 may extend axially between opposite leading and trailing edges 28, 29, respectively. The pressure face 26 and suction face 27 also may extend in the radial direction from an inboard end, i.e., the platform 24, to an outboard tip 31 of the airfoil 25. The airfoil 25 may include a curved or contoured shape extending between the platform 24 and the outboard tip 31. As illustrated in
For descriptive purposes, as provided in
The rotor blade 16 may further include an internal cooling configuration 36 having one or more cooling channels 37 through which a coolant is circulated during operation. The 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 outlet or surface ports through which coolant is exhausted from the rotor blade 16 and into the working fluid flowpath.
As shown, the tip shroud 41 may be positioned near or at the outboard end of the airfoil 25. The tip shroud 41 may include an axially and circumferentially extending flat plate or planar component, which is supported towards its center by the airfoil 25. For descriptive purposes, the tip shroud 41 may include an inboard surface 45, outboard surface 44, and edge 46. As illustrated, the inboard surface 45 opposes the outboard surface 44 across the narrow radial thickness of the tip shroud 41, while the edge 46 connects the inboard surface 45 to the outboard surface 44 and, as used herein, defines a peripheral profile or shape of the tip shroud 41.
A seal rail 42 may be positioned along the outboard surface 44 of the tip shroud 41. Generally, as illustrated, the seal rail 42 is a fin-like projection that extends radially outward from the outboard surface 44 of the tip shroud 41. The seal rail 42 may extend circumferentially between opposite ends of the tip shroud 41 in the direction of rotation or “rotation direction” of the rotor blade 16. As will be appreciated, the seal rail 42 may be used to deter leakage of working fluid through the radial gap that exists between the tip shroud 41 and the surrounding stationary components that define the outboard boundary of the working fluid flowpath through the turbine. In some conventional designs, the seal rail 42 may extend radially into an abradable stationary honeycomb shroud that opposes it across that gap. The seal rail 42 may extend across substantially the entire circumferential length of the outboard 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 rotation direction 50. For descriptive purposes, as indicated in
A cutter tooth 43 may be disposed on the seal rail 42. As will be appreciated, the cutter tooth 43 may be provided for cutting a groove in the abradable coating or honeycomb of the stationary shroud that is slightly wider than the width of the seal rail 42. As will be appreciated, the honeycomb may be provided to enhance seal stability, and the use of the cutter tooth 43 may reduce spillover and rubbing between stationary and rotating parts by clearing this wider path.
The tip shroud 41 may include fillet regions 48, 49 that are configured to provide smooth surficial transitions between the divergent surfaces of the tip shroud 41 and the airfoil 25, as well as those between the tip shroud 41 and the seal rail 42. As such, configurations of the tip shroud 41 may include an inboard fillet region 49 that is formed between the inboard surface 45 of the tip shroud 41 and the pressure and suction faces 26, 27 of the airfoil 25. The tip shroud 41 also may include an outboard fillet region 48 that is formed between the outboard surface 44 of the tip shroud 41 and the rail forward face 56 and aftward face 57 of the seal rail 42. As will be appreciated, the inboard fillet region 49 may further be described as including: a pressure inboard fillet region between the pressure face 26 of the airfoil 25 and the inboard surface 45 of the tip shroud 41; and a suction inboard fillet region between the suction face 26 of the airfoil 25 and the inboard surface 45 of the tip shroud 41. Similarly, the outboard fillet region 48 may be described as including: a pressure outboard fillet region between the rail forward face 56 and the outboard surface 44 of the tip shroud 41; and a suction outboard fillet region between the rail aftward face 57 and the outboard surface 44 of the tip shroud 41. As depicted, each of these fillet regions 48, 49 may be configured to provide smoothly curving transitions between the several planar surfaces that form abrupt or steeply angle transitions. As will be appreciated, such fillet regions may improve aerodynamic performance as well as spread stress concentrations that would otherwise occur in those areas. Even so, these areas remain highly stressed due to the overhanging or cantilevered load of the tip shroud 41 and the rotational speed of the engine. As will be appreciated, without adequate cooling, the stresses in these areas are a significant limit on the useful life of the component.
With particular reference now to
With particular reference now to
The tip shroud 41 may be described as including circumferential faces that, relative the rotation direction, may be designated as a rotationally leading circumferential face 72 and a rotationally trailing circumferential face 73. As used herein, the rotationally leading circumferential face 72 includes the rotationally leading edge 52 of tip shroud 41 and the rotationally leading edge 62 of seal rail 42. The rotationally trailing circumferential face 73 includes the rotationally trailing edge 53 of tip shroud 41 and the rotationally trailing edge 63 of seal rail 42. Further, an outboard face 59 of the seal rail 42 may be defined along the outer radial edge or face of the seal rail 42 that faces in the outboard direction. (Note that this component was previously referenced herein as the outboard edge 59 of the seal rail 42. Either term may be used interchangeably.) As illustrated in
With reference now to
The present invention may include tip shrouds having a configuration in which hollow cavities, pockets, chambers, voids and the like (which collectively will referred to herein as cavities) are formed to reduce tip shroud mass while also maintaining structural performance and robustness. These cavities may be enclosed via preformed coverplates that are brazed or welded into place. Alternatively, the coverplates may be applied by laser cladding, laser deposition, or other additive manufacturing processes. According to exemplary embodiments, such cavities may be strategically positioned so to reduce stresses applied to the tip shroud fillet regions and/or contact faces without also reducing the overall stiffness and structural performance in the affected regions. As described below, such cavities may be formed via conventional machining processes, including electro-chemical, chemical or mechanical processes. In alternative embodiments, the cavities may be formed during conventional blade casting processes for additive manufacturing processes. According to certain preferred embodiments, the cavities may be formed through one of several identified tip shroud surfaces, which are described below, and the cavities may be substantially or wholly contained within certain prescribed internal target regions associated with the seal rail. In this manner, the present invention may enable the removal of dead mass from particular internal regions of the tip shroud and/or seal rail so to reduce weight while maintaining overall structural resilience. As will be shown, present configurations may reduce the overall weight of the rotor blade without reducing or compromising other areas that are more structurally critical, such as those within the fillet regions or structurally active internal areas of the airfoil. The hollowed or cavitied portions may be optimally limited to target areas, which are readily identifiable based on the relative positioning and configuration of the tip shroud and seal rail. The present invention may optimize the location of the cavitied portions by delineating those internal regions that bear minimal bending load. In this manner, bending stiffness and overall structural robustness may be maintained, while mass is removed and, thus, operational stresses reduced.
As will be appreciated, such mass reduction may enable significant performance benefits. The weight reduction, for example, may simply reduce overall pull forces acting on the rotor blade during operation, and, thereby, extend creep life at life-limiting locations on the airfoil. Analysis of present configurations show creep life improvements to critical areas, such as fillet regions, by 5% to 20%. Alternatively, the weight reduction enabled by the present invention may be used to increase the overall size of the tip shroud without increasing overall weight. This, for example, may enable increasing the size of the contact faces of the tip shroud, which may reduce stress concentrations that occur when the tip shrouds of neighboring rotor blades engage during engine operation. Other examples include the possible reduction of fillet sizes or increase in tip shroud coverage, which may boost aerodynamic performance without increasing stress levels. Additionally, as provided below, the present invention includes efficient methods by which such enhanced tip shrouds may be constructed. That is to say, many of the present configurations may be cost-effectively constructed per the processes described herein. Additionally, the post-cast manufacturability of the exemplary methods allow for the efficient retrofitting of existing rotor blades, which may be used to extend component life.
Referring specifically now to
As will be appreciated, the edge zones 82, 83 may be used herein to define a range in which the cavities 90 of the present invention may be located. As stated, the cavities 90 may be defined has being wholly or substantially contained within one of the edge zones 82, 83, which, as used herein, means that the cavity 90 does not extend beyond or substantially beyond the edge zone and into the middle zone 84. As shown in
Further, according to preferred embodiments, as shown in
Additionally, the cavities 90 may be radially or circumferentially aligned. More specifically,
With specific reference now to
As one of ordinary skill in the art will appreciate, the many varying features and configurations described above in relation to the several exemplary embodiments may be further selectively applied to form the other possible embodiments of the present invention. For the sake of brevity and taking into account the abilities of one of ordinary skill in the art, all of the possible iterations is not provided or discussed in detail, though all combinations and possible embodiments embraced by the several claims below or otherwise are intended to be part of the instant application. In addition, from the above description of several exemplary embodiments of the invention, those skilled in the art will perceive improvements, changes and modifications. Such improvements, changes and modifications within the skill of the art are also intended to be covered by the appended claims. Further, it should be apparent that the foregoing relates only to the described embodiments of the present application and that numerous changes and modifications may be made herein without departing from the spirit and scope of the application as defined by the following claims and the equivalents thereof.
