Patent Application
20170058680
The present application relates generally to the design of the tips of gas turbine rotor blades. More specifically, but not by way of limitation, the present application relates to configurations of rotor blade tips that enhance aerodynamic and cooling performance.
In a gas turbine engine, it is well known that air is pressurized in a compressor and used to combust a fuel in a combustor to generate a flow of hot combustion gases, whereupon such gases flow downstream through one or more turbines so that energy can be extracted therefrom. In accordance with such a turbine, generally, rows of circumferentially spaced rotor blades extend radially outward from a supporting rotor disc. Each blade typically includes a root that permits assembly and disassembly of the blade in a corresponding slot formed in the rotor disc, as well as an airfoil that extends away from the root in a radially outward direction.
The airfoil has a generally concave pressure side and generally 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 turbine shroud for minimizing leakage therebetween of the combustion gases flowing downstream between the turbine blades. Maximum efficiency of the engine is obtained by minimizing the tip clearance or gap such that leakage is prevented, but this strategy is limited somewhat by the different thermal and mechanical expansion and contraction rates between the rotor blades and the turbine shroud and the motivation to avoid an undesirable scenario of having excessive tip rub against the shroud during operation.
Because turbine blades are bathed in hot combustion gases, effective cooling is required for ensuring a useful part life. Typically, the blade airfoils are hollow and disposed in flow communication with the compressor so that a portion of pressurized air bled therefrom is received for use in cooling the airfoils. Airfoil cooling in certain areas of the rotor blade is quite sophisticated and may be employed using various forms of internal cooling channels and features, as well as cooling outlets through the outer walls of the airfoil for discharging the cooling air. Nevertheless, airfoil tips are particularly difficult to cool since they are located directly adjacent to the turbine shroud and are heated by the hot combustion gases that flow through the tip gap. Accordingly, a portion of the air channeled inside the airfoil of the blade is typically discharged through the tip for the cooling thereof
It will be appreciated that conventional blade tip design includes several different geometries and configurations that are meant to prevent leakage and increase cooling effectiveness, as well as, improve aerodynamic performance and reduce mixing losses. However, conventional blade tip cooling designs, particularly those having a “squealer tip” design, have certain shortcomings, including the inefficient usage of compressor bypass air, which reduces plant efficiency. As a result, an improved turbine blade tip design that increases the overall effectiveness of the coolant directed to this region would be highly desired.
The present application thus describes a rotor blade for a turbine of a gas turbine system. The rotor blade may include: an airfoil that includes a pressure sidewall and a suction sidewall defining an outer periphery, wherein the pressure sidewall and suction sidewall of the airfoil connect along a leading edge and a trailing edge; a tip that defines an outer radial end of the airfoil, wherein the tip included a cap on which an outboard projecting rail defines a tip cavity; and a rail gap formed through an aftward section of the 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.
The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:
The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings.
Aspects and advantages of the invention 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. Thus, 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 appreciate 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 configuration, function, and/or usage of the component being referenced and described, including the manner in which the term relates to the several figures, and, of course, the precise usage of the terminology in the appended claims. Further, while the following examples are presented in relation to a certain type of gas turbine or turbine engine, the technology of the present invention also may be applicable to other types of turbine engines as would the understood by a person of ordinary skill in the relevant technological arts.
Several descriptive terms may be used throughout this application so to explain the functioning of turbine engines and/or the several sub-systems or components included therewithin, and it may prove beneficial to define these terms at the onset of this section. Accordingly, these terms and their definitions, unless stated otherwise, are as follows. The terms “forward” and “aft” or “aftward”, without further specificity, refer to directions relative to the orientation of the gas turbine. “Forward” refers to the compressor end of the engine, while “aft” or “aftward” refers to the turbine end of the engine. Each of these terms, thus, may be used to indicate movement or relative position along the longitudinal axis of the machine. The terms “downstream” and “upstream” are used to indicate position within a specified conduit relative to the general direction of flow moving through it. As will be appreciated, these terms reference a direction relative to the direction of flow expected through the specified conduit during normal operation, which should be plainly apparent to any skilled in the art. As such, the term “downstream” refers to the direction in which the fluid is flowing through the specified conduit, while “upstream” refers to the opposite of that. Thus, 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 and beyond, may be described as beginning at an upstream location toward an upstream or forward end of the compressor and terminating at an downstream location toward a downstream or aft end of the turbine.
In regard to describing the direction of flow within a common type of combustor, as discussed in more detail below, it will be appreciated that compressor discharge air typically enters the combustor through impingement ports that are concentrated toward the aft end of the combustor (relative to the combustors longitudinal axis and the aforementioned compressor/turbine positioning defining forward/aft distinctions). Once in the combustor, the compressed air is guided by a flow annulus formed about an interior chamber toward the forward end of the combustor, where the air flow enters the interior chamber and, reversing it direction of flow, travels toward the aft end of the combustor. In yet another context, the flow of coolant through cooling channels or passages may be treated in the same manner.
Additionally, given the configuration of compressor and turbine about a central common axis, as well as the cylindrical configuration common to many combustor types, terms describing position relative to an axis may be 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 this case, 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 than the second component, the first component will be described herein as being either “radially outward” or “outboard” of the second component. Additionally, as will be appreciated, the term “axial” refers to movement or position parallel to an axis. Finally, the term “circumferential” refers to movement or position around an axis. As mentioned, while these terms may be applied in relation to the common central axis that extends through the compressor and turbine sections of the engine, these terms also may be used in relation to other components or sub-systems of the engine.
Finally, the term “rotor blade”, without further specificity, is a reference to the rotating blades of either the compressor or the turbine, which 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, which include both compressor stator blades and turbine stator blades. The term “blades” will be used herein to 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, and turbine stator blades.
By way of background, referring now to the figures,
In one example of operation, the rotation of the rotor blades 14 within the axial compressor 11 compresses a flow of air. In the combustor 13, energy is released when the compressed air flow is mixed with a fuel and ignited. The resulting flow of hot gases from the combustor 13, which may be referred to as the working fluid, is then directed over the turbine rotor blades 16, with the flow of working fluid thereby inducing the rotation of the rotor blades 16 about the shaft. In this manner, the energy of the flow of working fluid may be transformed into the mechanical energy of the rotating blades and, given the connection between the rotor blades and the shaft via the rotor disc, the rotating shaft. The mechanical energy of the shaft then may be used to drive the rotation of the compressor rotor blades, such that the necessary supply of compressed air is produced, and also, for example, a generator for the production of electricity.
For background purposes,
With reference also to
In general, rotor blades may include blade tips that are shrouded or, as illustrated, unshrouded. In the case of unshrouded tips, the blade tip 50, as illustrated, may include a tip cap or cap 51 that is disposed atop the radially outer edges of the pressure 26 and suction sidewalls 27. The cap 51 typically bounds interior cooling passages (which, as provided below, are referenced herein as a “cooling passages 36”) that may be defined between the pressure 26 and suction sidewalls 27 of the airfoil 25. As is typical, coolant, such as compressed air bled from the compressor, may be circulated through the cooling passages 36 during operation. The cooling passages 36 may be supplied the coolant via connector channels (not shown) that are formed through the root 21 of the rotor blade 16. The cap 51 typically includes a plurality of outlet ports 35 that connect to the cooling passages 36. During operation, the outlet ports 35 release coolant after it has circulated through the interior of the airfoil 25 and may be positioned to promote film cooling over the surface of the blade tip 50. The cap 51 may be integral to the rotor blade 16 or a portion may be welded/brazed into place after the rotor blade is cast.
Due to certain performance advantages, such as those related to greater aerodynamic and cooling efficiencies, the tips 50 of rotor blades frequently include a tip cavity 52 formed via one or more rails 53 that project radially from the cap 51. This type of blade tip is commonly referred to as a “squealer tip” or, alternatively, as a tip having a “squealer pocket” or “squealer cavity.” The positioning of the one or more rails 53 may coincide approximately with the profile of the pressure 26 and suction sidewalls 27 of the airfoil 25. As such, the one or more rails 53 may be referred to as including, respectively, a pressure rail 54 and a suction rail 55. The pressure rail 54, as illustrated, may extend radially outward from the cap 51 and, in relation thereto, may form an angle of approximately 90°. The pressure rail 54 may extend from a position near the leading edge 28 of the airfoil 25 to a position near the trailing edge 29. As illustrated, the path of the pressure rail 54 may coincide approximately with the profile of the pressure sidewall 26. Similarly, as illustrated, the suction rail 55 may project radially outward from the cap 51 and be approximately perpendicular therewith. The suction rail 55 may extend from a position near the leading edge 28 of the airfoil 25 to a position near the trailing edge 29. As illustrated, the path of the suction rail 55 may coincide approximately with the profile of the suction sidewall 27. Further, as indicated in
Those of ordinary skill in the art will appreciate that squealer tips of rotor blade in which the present invention is employed might vary somewhat from the characteristics described above. For example, the rail 53 may not necessarily follow precisely the profile of the outer radial edge of the pressure and/or suction sidewalls 26, 27. That is, in alternative types of blades tips, the tip rails 53 may be moved away from the outer periphery of the cap 51. In addition, as provided herein, the tip rails 53 may not surround the tip cavity completely and, in certain cases, as provided herein, large gaps may be formed within the tip rails 53, therein, particularly in the portion of the rail 53 positioned toward the trailing edge 29 of the airfoil. In some cases, sections of the rail 53 might be removed from either the pressure side or the suction side of the tip 50. Alternatively, one or more internal rails or ribs may be positioned between the pressure rail 54 and suction rails 55.
The tip rail 53, as shown, generally, is configured to circumscribe the cap 51 such that the tip cavity 52 is defined thereon. The height and width of the pressure rail 54 and/or the suction rail 55 (and, thus, the depth of the tip cavity 52) may be varied depending on best performance and the size of the overall turbine assembly. It will be appreciated that the cap 51 forms the floor of the tip cavity 52 (i.e., the inner radial boundary of the tip cavity), the tip rail 53 forms the sidewalls of the tip cavity 52, and that the tip cavity 52 remains open through an outer radial face, which, once installed within a turbine engine, is bordered closely by a stationary shroud 20 (as shown in
A plurality of outlet ports 35 may be disposed on the blade tip 50 as well as through the other outer surfaces of the airfoil 25. Typically, outlet ports 35 are provided through the pressure sidewall 26 of the airfoil 25 as well as through the cap 51. Some designs positions many such outlet ports 35 in the limited space available at the tip 35 in an effort to flood the pressures side tip region with coolant. In regard to the outlet ports 35 disposed on the pressure sidewall 26, it may be desired that, after its released, the coolant carries over the pressure rail 54 and into the tip cavity 52 to provide cooling therein and, then, over the suction rail 55 to provide cooling to this region. Toward this objective, the outlet ports 35 may be oriented so to direct coolant being expelled therefrom in the radially outward direction. As illustrated in
The winglet 80 may further be described as extending between a leading end 86 and a trailing end 87. As illustrated, the leading end 86 of the winglet may be positioned toward the leading edge 28 of the airfoil 25, while the trailing end 87 may be positioned near the trailing edge 29 of the airfoil 25. According to an exemplary embodiment, the leading end 86 of the winglet 80 may be positioned approximately 0% to 50% of a chordwise distance from the leading edge 28 of the airfoil 25, and the trailing end 87 of the winglet 80 may be positioned approximately 50% to 100% of the chordwise distance from the leading edge 28 of the airfoil 25. More preferably, the leading end of the winglet may be positioned approximately 20% to 40% of a chordwise distance from the leading edge, and the trailing end may be positioned approximately 70% to 90% of the chordwise distance from the leading edge.
Additionally, as illustrated, a plurality of cooling outlets 35 may be formed through the cap 51 so to release coolant within the tip cavity 52. The cooling outlets 35 may be connected to one or more cooling passages formed through the airfoil 25. The cooling outlets 35 may have rectangular cross-sectional shape, as illustrated, or may be circular, oval or other shape as required. Though other configurations are also possible, the radial section 82 of the winglet 80 may approximately coincide with the tip rail 53. In such cases, as will be appreciated, the inboard boundary 83 of the winglet 80 may be approximately coplanar with the cap 51, and the outboard boundary 83 of the winglet 80 may be approximately coplanar with the outboard rail surface 58 of the rails 53. According to another embodiment, the winglet 80 may be configured to begin at a radial location that is inboard or outboard of the cap 51.
Because the winglet 80 is inclined outwardly from the radial direction, it should be appreciated that its inclusion may result in widening the tip cavity 52 as the tip cavity 52 extends from the cap 51 in the outboard direction. The result of this is that, in operation, gases leaking over certain areas of the pressure rail 54 encounter a region of the tip cavity 52 that has an increased width. As will be appreciated, this increased width of the tip cavity 52 may enable the overflowing gases more of an opportunity to reattach within the tip cavity, and so remain in contact with the tip 50 longer and, thereby, improve cooling and aerodynamic performance. As discussed below, the winglet 80 feature may be used with the certain other aspects of the present invention discussed below so to achieve greater performance benefits.
Turning now to
According to alternative embodiments, as illustrated via the progressively larger width dimensions of the rail gaps 90 of
According to alternative embodiments, the rail gap 90 is combined with a suction side winglet 80, as depicted in the
Accordingly, the present invention provides a blade tip configuration in which a full or approximately full rail is provided on the pressure rail of the squealer tip, while an optimally truncated rail is provided on the suction rail of the squealer tip. As will be appreciated, the truncated rail forms a rail gap which may be configured to advantageously control the exit of coolant from the blade tip as well as affect over the blade tip leakage flowing toward the aft portions of the suction side of the airfoil. It will be appreciated that by combining the rail gap feature with the winglet, the cross-sectional flow area from the enclosed region of the tip cavity is made larger due to the of the manner in which the outward flaring of the winglet widens the tip cavity in the radial direction (as the tip cavity extends between the cap and the outer rail surface). By increasing this flow area, the rail gap, according to preferred embodiments, may be optimized at around 15%. At this design point, the flow area from the enclosed portion of the tip cavity is still larger enough to handle all of the extra flow captured by the winglet and/or supplied through outlet ports and induce flow along the desired flowpath, which is basically afterward within the tip cavity for eventual exit through the rail gap. As will be appreciated, without the winglet feature, the same flow area could only be achieved by increasing the size of the rail gap, which would negatively impact performance. Accordingly, rail gaps conforming to the embodiments provided herein, particularly when combined with the winglet feature, have been shown through experimental data to lower mixing losses, increase cooling efficiency, improve rotor blade torque, and provide lower pressures within the tip cavity of the squealer tip, as it establishes aerodynamic communication between the tip cavity and the low pressure region associate with the suction sidewall of the airfoil. As will be appreciated, the reduction in tip cavity pressure by this feature may generally assist in maintaining a relatively lower supply pressure for the blade tip cooling circuit so to improved back flow margin, which, in turn, may result in overall turbine efficiency improvement and reduced risk for rotor blade ingestion and overheating. For example, experimental data suggests that rail gaps such as those disclosed herein may be employed to provide 5-10 psi benefit in tip cavity pressure. Additionally, efficiency benefits associated with such rail gaps have been shown to improve overall system efficiency by approximately 0.1-0.2 points, while also having the flexibility to realize such benefits across a variety of the blade tip configurations, including varied squealer cavity depths and winglet features.
While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.
