Patent Grant
8157504
The present application relates generally to apparatus, methods and/or systems concerning the design of turbine rotor blade tips. More specifically, but not by way of limitation, the present application relates to apparatus, methods and/or systems related to turbine blade tips that include a trailing edge trench cavity that, among other advantages, improves the cooling of the blade tip.
In a gas turbine engine, it is well known that air pressurized in a compressor is 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 turbine rotor blades extend radially outwardly from a supporting rotor disk. Each blade typically includes a dovetail that permits assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disk, 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 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. Improved 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 the tip rub against the shroud during operation.
In addition, 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 is quite sophisticated and may be employed using various forms of internal cooling channels and features, as well as cooling holes 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 prevent leakage and increase cooling effectiveness. Exemplary patents include: U.S. Pat. No. 5,261,789 to Butts et al.; U.S. Pat. No. 6,179,556 to Bunker; U.S. Pat. No. 6,190,129 to Mayer et al.; and, U.S. Pat. No. 6,059,530 to Lee. Conventional blade tip designs, however, all have certain shortcomings, including a general failure to adequately reduce leakage and/or allow for efficient tip cooling that minimizes the use of efficiency-robbing compressor bypass air. Improvement in the pressure distribution near the tip region is still sought to further reduce the overall tip leakage flow and thereby increase turbine efficiency. As a result, a turbine blade tip design that alters the pressure distribution near the tip region and otherwise reduces the overall tip leakage flow, thereby increasing the overall efficiency of the turbine engine, would be in great demand. Further, it is also desirable for such a blade tip to enhance the cooling characteristics of the cooling air that is released at the blade tip, as well as, enhancing the overall aerodynamic performance of the turbine blade. Particularly, it would be desirable for an improved tip design that better allowed the flow of cooling air to move toward the trailing edge of the tip blade, which, generally, is a difficult area to cool.
The present application thus describes a blade tip of a turbine rotor blade for a gas turbine engine, the turbine rotor blade including an airfoil and a root portion for mounting the airfoil along a radial axis to a rotor disk inboard of a turbine shroud, a pressure sidewall and a suction sidewall that join together at a leading edge and a trailing edge, the pressure sidewall and suction sidewall extending from the root portion to the blade tip, and a squealer tip cavity formed at the blade tip, the blade tip comprising: a trailing edge trench originating at the squealer tip cavity, wherein the trailing edge trench generally extends toward the trailing edge of the blade tip.
In some embodiments, the blade tip comprises a tip plate that extends between the outer radial edge of the pressure sidewall to the outer radial edge of the suction sidewall; the squealer tip cavity is formed on a first side by a pressure tip wall that extends radially outwardly from the tip plate, traversing from the leading edge to the trailing edge such that the pressure tip wall resides approximately adjacent to the termination of the pressure sidewall; and the squealer tip cavity is formed on a second side by a suction tip wall that extends radially outwardly from the tip plate, traversing from the leading edge to the trailing edge such that the suction tip wall resides approximately adjacent to the termination of the suction sidewall.
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:
Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,
Each blade 18 generally includes a dovetail 22 which may have any conventional form, such as an axial dovetail configured for being mounted in a corresponding dovetail slot in the perimeter of the rotor disk 16. A hollow airfoil 24 is integrally joined to dovetail 22 and extends radially or longitudinally outwardly therefrom. The blade 18 also includes an integral platform 26 disposed at the junction of the airfoil 24 and the dovetail 22 for defining a portion of the radially inner flowpath for combustion gases 12. It will be appreciated that the blade 18 may be formed in any conventional manner, and is typically a one-piece casting.
It will be seen that the airfoil 24 preferably includes a generally concave pressure sidewall 28 and a circumferentially or laterally opposite, generally convex suction sidewall 30 extending axially between opposite leading and trailing edges 32 and 34, respectively. The sidewalls 28 and 30 also extend in the radial direction between a radially inner root 36 at the platform 26 and a radially outer tip or blade tip 38, which will be described in more detail in the discussion related to
The inside of the airfoil 24 may have any configuration including, for example, serpentine flow channels with various turbulators therein for enhancing cooling air effectiveness, with cooling air being discharged through various holes through airfoil 24 such as conventional film cooling holes 44 and trailing edge discharge holes 46.
As better appreciated in
Similarly, the suction tip wall 52 generally extends radially outwardly from the tip plate 48 and extends axially from the leading edge 32 to the trailing edge 34. The path of suction tip wall 52 is adjacent to or near the termination of the suction sidewall 30 (i.e., at or near the periphery of the tip plate 48 along the suction sidewall 30). The height and width of the pressure tip wall 50 and/or the suction tip wall 52 may be varied depending on best performance and the size of the overall turbine assembly. As shown, the pressure tip wall 50 and/or the suction tip wall 52 may be approximately rectangular in shape; other shapes are also possible. A tip mid-chord line 60 also is depicted as a dashed line on
The pressure tip wall 50 and the suction tip wall 52 generally form what is referred to herein as a squealer tip cavity 62. In generally terms, the squealer tip cavity 62 may include any radially inward extending depression or cavity formed on the blade tip 38. Generally, the squealer tip cavity 62 has a similar shape or form as the airfoil 24, though other shapes are possible, and be bound by: 1) a radially outward extending wall aligned with the pressure sidewall 28, which herein has been described as the pressure tip wall 50; 2) a radially outward extending wall aligned with the suction sidewall 30, which herein has been described as the suction tip wall 52; 3) and an inner radial floor, which herein has been described as the tip plate 48. The squealer tip cavity 62 may be open through the plane that defines the outer radial limits of the cavity 62. As a result, generally, upon installation, the squealer tip cavity 62 is substantially enclosed by the surrounding stationary shroud 20, though the outer surface of pressure tip wall 50 and the suction tip wall 52 are offset from the shroud 20 by a desired clearance.
As one of ordinary skill in the art will appreciate, one or more cooling apertures (not shown in
The trailing edge trench 72 of the present invention may comprise several different shapes, sizes, alignments, and configurations, as discussed in detail below. For example, as shown in
As better appreciated in
The depth of the trailing edge trench 72 may be substantially constant as it extends toward the trailing edge 34. Note that as used herein, the depth of the trailing edge trench 72 is meant to refer to the maximum radial height of the trench 72 at a given location on its path. Thus, in the case of a semi-elliptical profile, the depth of the trailing edge trench 72 occurs at the inward apex of the elliptical shape. In some preferred embodiments, the depth of the trailing edge trench 72 may be between approximately 110% and 40% of the depth of the aft end of the squealer tip cavity 62 (i.e., the approximate position in the squealer tip cavity 62 where the trailing edge trench 72 originates). More preferably, the depth of the trailing edge trench 72 may be between approximately 100% and 75% of the depth of the aft end of the squealer tip cavity 62 (i.e., the approximate position in the squealer tip cavity 62 where the trailing edge trench 72 originates).
In other embodiments, as shown in
In some embodiments, the trailing edge trench 72 may have a substantially constant width as it extends from the squealer tip cavity 62 to the trailing edge 34. Note that as used herein, the width of the trench 72 is meant to comprise the distance across the trench 72 at its mouth. In preferred embodiments, the width of the squealer tip cavity 62 generally may be between 95% and 40% of the width of the aft end of the squealer tip cavity 62 (i.e., the approximate position in the squealer tip cavity 62 where the trailing edge trench 72 originates). More preferably, the width of the squealer tip cavity 62 may be between 80% and 50% of the width of the aft end of the squealer tip cavity 62.
In other preferred embodiments, the width of the trailing edge trench 72 may gradually decrease as the trench 72 extends from the aft end of the squealer tip cavity 62 toward the trailing edge 34 of the airfoil. In such cases, the width of the trench 72 generally narrows in proportion to the narrowing shape of the aft end of the blade tip 38. The width of trench 72, in such embodiments, generally may be between approximately 30%-80% of the width of the blade tip 38 through aft end of the airfoil. More preferably, the width of trench may be between approximately 40%-70% of the width of the blade tip 38 through aft end of the airfoil.
Note that the transition between the squealer tip cavity 62 and the trailing edge trench 72 may be made in several different ways. For example, the transition between the squealer tip cavity 62 and the narrower width of the squealer tip cavity 62 may be “stepped” in nature (i.e., a sharp corner) or have a blended edge (i.e., a smooth or rounded corner). As one of ordinary skill in the art will appreciate, in some applications, the blended edge may promote smoother flow into the trailing edge trench 72, which, generally, may allow more of the cooling air to remain in the trailing edge trench 72 as it moves toward the trailing edge 34 of the blade tip 38, which may enhance the cooling effects of the air.
The trailing edge trench 72 may have one or more trench cooling apertures 74, which similar to the previously discussed cooling apertures. The trench cooling apertures 74 are openings within the trench 72 that connect to cooling cavities within the airfoil. Per conventional means, a coolant may be directed through the trench cooling apertures 74 and, along with the flow of coolant from the squealer tip cavity 62, keep the surrounding surface area of the blade tip 38 cool by convecting away heat and insulating the part from the extreme temperatures of the working fluid. More particularly, the coolant may better cool the trailing edge portion of the blade tip 38. As shown, the trench cooling apertures may be regularly spaced through the trailing edge trench 72 and positioned on the floor of the trench 72, i.e., near the deepest portion of the trench 74.
As shown in
In use, the trailing edge trench 72 generally improves the cooling of the trailing edge 34 of the blade tip 38 without an increase in the amount of coolant flow. The trench 72 generally takes coolant flow of the squealer tip cavity 62 that would otherwise be washed over the suction tip wall 52 and directs it toward the trailing edge 34 of the blade tip 38. Particularly, the trailing edge trench 72 generally provides a downstream oriented path that allows the coolant within the squealer tip cavity 62 to more effectively reach the lower pressure gradients that generally exist during operation at the trailing edge 34 of the blade tip 38. The coolant thereby reaches the trailing edge region without: 1) being washed away by the pressure side hot gases; or 2) without creating disturbances on the suction side flow. Further, as one of ordinary skill in the art will appreciate, the resulting decrease in trailing edge temperatures generally reduces the amount of oxidation that occurs during operation along the trailing edge 34 of the blade tip 38. The reduction of oxidation improves the aerodynamic performance of the airfoil and, ultimately, reduces repair costs. In addition, the flow patterns that results from the geometry of the trailing edge trench 72 act as a seal across that portion of the blade tip 38 as they prevent flow from slipping over the blade tip 38 from the pressure side to the suction side, which, as one of ordinary skill in the art will appreciate, improves engine performance. As such, in sum, the trailing edge trench of the present application generally decreases the metal temperatures at the trailing edge of the blade tip, thereby increasing the part life, improving the performance of the engine by preventing oxidation, and reducing the costs of maintenance, while also improving engine efficiency with its better sealing characteristics.
From the above description of preferred 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 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.
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| Number | Date | Country | |
|---|---|---|---|
| 20100266410 A1 | Oct 2010 | US |
