The inventive subject matter generally relates to turbine assemblies, and more particularly relates to turbine blades for turbine assemblies.
Gas turbine engines, such as turbofan gas turbine engines, may be used to power various types of vehicles and systems, such as, for example, aircraft. Typically, these engines include turbine blades that are impinged on by high-temperature compressed air that causes a turbine of the engine to rotate at a high speed. Consequently, the blades are subjected to high heat and stress loadings which, over time, may reduce their structural integrity.
To improve blade structural integrity, a blade cooling scheme is typically incorporated into the turbines. The blade cooling scheme directs cooling air through an internal cooling circuit formed in the blade to maintain blade temperatures within acceptable limits. The internal cooling circuit may include a simple channel extending through a length of the blade or may consist of a series of connected, serpentine cooling passages, which incorporate raised or depressed structures therein. The serpentine cooling passages increase the cooling effectiveness by extending the length of the air flow path. In this regard, the blade may have multiple internal walls that form the intricate cooling passages through which the cooling air flows. The cooling passages then direct the cooling air to openings on a tip and a trailing edge of the blade.
As the desire for increased engine efficiency continues to rise, engine components are increasingly being subjected to higher and higher operating temperatures. For example, newer engine designs may employ operating temperatures that are over 1100° C. However, current engine components, such as the blades, may not be adequately designed to withstand such temperatures. In particular, the blade tip may abrade against an inner surface of a surrounding shroud, and as a result, the openings providing an outlet for the cooling air in the blade tip may become filled with the blade material. Hence, the blade may not be cooled as desired.
Accordingly, it is desirable to have a blade with an improved manner for cooling the blade tip. Additionally, it is desirable for the blade to maintain coolant flow during engine operation. Furthermore, other desirable features and characteristics of the inventive subject matter will become apparent from the subsequent detailed description of the inventive subject matter and the appended claims, taken in conjunction with the accompanying drawings and this background of the inventive subject matter.
In an embodiment, by way of example only, a turbine blade includes a first side wall including a first tip edge, a second side wall opposite the first side wall and including a second tip edge, a tip wall extending between the first side wall and the second side wall, the tip wall being recessed from the first tip edge of the first side wall and the second tip edge of the second side wall to form a coolant cavity, a tip recess cavity, a first parapet wall on the first side wall, and a second parapet wall on the second side wall, the coolant cavity defined in part by an interior surface of the tip wall, and the tip recess cavity defined, in part, by a surface of the tip wall, the first parapet wall, and the second parapet wall, a step formed between the first tip edge and the tip wall, a cooling hole extending through the first parapet wall, the step, and the tip wall, the cooling hole including an open channel section and a closed channel section, the open channel section extending from the first tip edge of the parapet wall to the step, and the closed channel section extending through the step and the tip wall.
In another embodiment, by way of example only, a turbine rotor includes a rotor and a plurality of blades extending radially outwardly from the rotor. Each blade comprises a first side wall including a first tip edge, a second side wall opposite the first side wall and including a second tip edge, a tip wall extending between the first side wall and the second side wall, the tip wall being recessed from the first tip edge of the first side wall and the second tip edge of the second side wall to form a coolant cavity, a tip recess cavity, a first parapet wall on the first side wall, and a second parapet wall on the second side wall, the coolant cavity defined in part by an interior surface of the tip wall, and the tip recess cavity defined, in part, by a surface of the tip wall, the first parapet wall, and the second parapet wall, a step formed between the first tip edge and the tip wall, and a cooling hole extending through the first parapet wall, the step, and the tip wall, the cooling hole including an open channel section and a closed channel section, the open channel section extending from the first tip edge of the parapet wall to the step, and the closed channel section extending through the step and the tip wall.
In another embodiment, by way of example only, a method of manufacturing a blade includes casting a blade including a first side wall including a first tip edge, a second side wall opposite the first side wall and including a second tip edge, and a tip wall extending between the first side wall and the second side wall, the tip wall being recessed from the first tip edge of the first side wall and the second tip edge of the second side wall to form a coolant cavity, a tip recess cavity, a first parapet wall on the first side wall, and a second parapet wall on the second side wall, the coolant cavity defined in part by an interior surface of the tip wall, and the tip recess cavity defined, in part, by a surface of the tip wall, the first parapet wall, and the second parapet wall, forming a step between the first tip edge and the tip wall, and machining a cooling hole into the blade, wherein the cooling hole extends through the first parapet wall, the step, and the tip wall, the cooling hole includes an open channel section and a closed channel section, the open channel section extends from the first tip edge of the parapet wall to the step, and the closed channel section extends through the step and the tip wall.
The inventive subject matter will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and
The following detailed description is merely exemplary in nature and is not intended to limit the inventive subject matter or the application and uses of the inventive subject matter. Furthermore, there is no intention to be bound by any theory presented in the preceding background or the following detailed description.
An improved turbine blade is provided that is capable of withstanding temperature environments that are higher than those for which conventional turbine blades are designed. Generally, the improved turbine blade includes a first side wall including a first tip edge, a second side wall opposite the first side wall and including a second tip edge, and a tip wall extending between the first side wall and the second side wall, the tip wall being recessed from the first tip edge of the first side wall and the second tip edge of the second side wall to form a coolant cavity, a tip recess cavity, a first parapet wall on the first side wall, and a second parapet wall on the second side wall, the coolant cavity defined in part by an interior surface of the tip wall, and the tip recess cavity defined, in part, by an exterior surface of the tip wall, the first parapet wall, and the second parapet wall. To provide improved cooling, the turbine blade further includes a step formed between the first tip edge and the tip wall, the step extending along a majority of a length of the first tip edge of the first side wall, and a cooling hole having a centerline extending from the first parapet wall, through the step, and through the tip wall, the continuous cooling hole including an open channel section and a closed channel section, the open channel section extending from the first tip edge of the parapet wall to the step, and the closed channel section extending through the step and the tip wall. The improved turbine blade may be implemented into turbine assemblies for turbine engines or for other turbine applications.
The rotor 50 includes a blade ring 52, a disk 54, and a plurality of blades 56. The blade ring 52 and the disk 54 are bonded together and may be made of similar or different materials, in an embodiment. In another embodiment, the blade ring 52 is inserted into the disk 54 through a disk attachment mechanism. Suitable materials that may be used for manufacturing the blade ring 52 and/or the disk 54 include, but are not limited to superalloys, such as nickel-based superalloys, that are equi-axed, uni-directional, or single crystal. The uni-directional and single crystal materials may each have a preferential crystal orientation.
The blade ring 52 has a plurality of inlets 66 for receiving air for cooling the blades 56. The inlets 66 communicate with coolant cavities (e.g., coolant cavity 316 of
The airfoil 104 is generally made up of a concave, pressure side wall 110, a convex, suction side wall 112 opposite the concave, pressure side wall 110, and a tip wall 114 extending between and coupling the pressure sidewall 110 and the suction side wall 112 together. The walls 110, 112, 114 may each have varying thicknesses along their lengths. In an embodiment, the walls 110, 112, 114 may have thicknesses that range between about 0.20 mm and 1.80 mm. In still other embodiments, the walls 110, 112, 114 may each have equal thicknesses, while in other embodiments the walls 110, 112, 114 may each have substantially equal thickness. In any case, the walls 110, 112, 114 have outer surfaces that together define an airfoil shape. The airfoil shape is made up of a leading edge 116, a trailing edge 118, a pressure side 120 along the concave, pressure side wall 110, a suction side 122 along the convex, suction side wall 112, one or more trailing edge slots 124, an airfoil platform fillet 126, and a tip recess cavity 128.
In any case, the tip wall 308 extends between the first side wall 304 and the second side wall 306 and is recessed a distance from the first and second tip edges 312, 314 to define first and second parapet walls 320, 322 on the first and second side walls 304, 306, respectively. An exposed surface 328 of the recessed tip wall 308, a first parapet wall 320 on the first side wall 304, and a second parapet wall 322 on the second side wall 306 together form a tip recess cavity 318. The parapet walls 320, 322 are substantially equal in height (as measured from the exposed surface 328 of the tip wall 308 to the first and second tip edges 312, 314, respectively), as depicted in
During operation, as noted above, when the rotor (e.g., rotor 50) rotates, air from an airflow is ingested and directed to a corresponding blade, such as blade 300. Because the radial gap between the rotor and the shroud (e.g., shroud 20 in
The step 330 is formed between the first tip edge 312 and the exposed surface 328 of the tip wall 308. Although the step 330 is depicted as being formed on the first parapet wall 320, other embodiments alternatively may include the step 330 on the second parapet wall 322. By including the step 330, the parapet wall 320 is divided into an outer radial section 332 and inner radial section (e.g., the step 330). The outer radial section 332 is defined by the tip edge 312 and an outer axial surface 340. The step 330 is defined by a radial surface 342 and an inner axial surface 344. Although illustrated in
The outer radial section 312 is configured to contact or have a small radial gap to the shroud (e.g., shroud 20 in
The cooling hole 334 has a centerline 338 and extends continuously from the parapet wall 320 (e.g., through the outer radial section 332 and the step 330) and the tip wall 308. The cooling hole 334 has an open channel section 364 and a closed channel section 366 (shown in
As illustrated in the exemplary embodiment of
In another embodiment, the centerline is substantially orthogonal relative to the first tip edge.
In still another embodiment, the shape and dimensions of the open and closed channel sections are not constant.
The closed channel section 766 extends along the centerline 738 and has a shape, where the shape of the diffuser angle section is not a portion of the shape of the closed channel section 766. For example, the closed channel section 766 has a cylindrical shape, as illustrated in
Turning now to
A plurality of cooling holes 832 are formed through the first parapet wall 820 including the step 830. Each hole 832 has a largest diameter in a range of about 0.20 mm to about 0.70 mm. In other embodiments, the holes 832 are larger or smaller than the aforementioned range. The cooling holes 832 are substantially evenly spaced apart, in an embodiment. In other embodiments, the cooling holes 832 unevenly spaced apart. In any case, each cooling hole 832 has a largest diameter, and each cooling hole 832 spaced apart a distance from an adjacent cooling hole 832 that is equal to between about three to about eight largest cooling hole diameters. Depending on a total length of the first parapet wall 820, a total number of holes 832 on the first parapet wall 820 can fall within in a range of about 15 to about 25 holes.
To manufacture a blade including the features described above, the blade including a tip portion with parapet walls (e.g., walls 320, 322, 520, 720, 820, 822) and a recessed tip wall (e.g., 308, 508, 708, 808) may be formed by a conventional lost wax casting process. A step (e.g., step 330, 530, 730, 830) is also included in the blade that is formed by the conventional lost wax casting process, in an embodiment. In another embodiment, the step and the holes are electro-discharge machined into the desired parapet wall. In still another embodiment the step and the holes are formed into the desired parapet wall by employing a different fabrication process, such as by laser sintering. Cooling holes (e.g., holes 334, 532, 732, 832) are electrodischarge machined through the parapet wall, in accordance to a desired configuration similar to one described above. In still another embodiment, the hole or the step or both can be machined by laser machining.
A blade has now been provided that has an improved manner for cooling a tip section of the blade. In particular, by including a step in a parapet wall of the blade, and by extending a cooling hole through the parapet wall (including the step) and a tip wall of the blade, cool air from a coolant cavity of the blade can be directed to the tip wall and the parapet wall. Moreover, because a portion of the cooling hole is configured as an open channel section (e.g., groove), air can still be supplied to the parapet wall in an event in which a portion of the parapet wall abrades against the shroud and causes blockage in the hole openings on the outer radial section (eg. 312, 512, and 712).
While at least one exemplary embodiment has been presented in the foregoing detailed description of the inventive subject matter, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the inventive subject matter in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the inventive subject matter. It being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the inventive subject matter as set forth in the appended claims.
This inventive subject matter was made with Government support under USAF F33615-03-D-2355 awarded by the United States Air Force. The Government has certain rights in this inventive subject matter.
Number | Name | Date | Kind |
---|---|---|---|
4390320 | Eiswerth | Jun 1983 | A |
4529357 | Holland | Jul 1985 | A |
4589823 | Koffel | May 1986 | A |
5039562 | Liang | Aug 1991 | A |
5192192 | Ourhaan | Mar 1993 | A |
5282721 | Kildea | Feb 1994 | A |
5688107 | Downs et al. | Nov 1997 | A |
5733102 | Lee et al. | Mar 1998 | A |
6179556 | Bunker | Jan 2001 | B1 |
6190129 | Mayer et al. | Feb 2001 | B1 |
6231307 | Correia | May 2001 | B1 |
6422821 | Lee et al. | Jul 2002 | B1 |
6478535 | Chung et al. | Nov 2002 | B1 |
6494678 | Bunker | Dec 2002 | B1 |
6527514 | Roeloffs | Mar 2003 | B2 |
6602052 | Liang | Aug 2003 | B2 |
6634860 | Lee et al. | Oct 2003 | B2 |
6672829 | Cherry et al. | Jan 2004 | B1 |
6790005 | Lee et al. | Sep 2004 | B2 |
6981846 | Liang | Jan 2006 | B2 |
6994514 | Soechting et al. | Feb 2006 | B2 |
7192250 | Boury et al. | Mar 2007 | B2 |
7351035 | Deschamps et al. | Apr 2008 | B2 |
7473073 | Liang | Jan 2009 | B1 |
7494319 | Liang | Feb 2009 | B1 |
7510376 | Lee et al. | Mar 2009 | B2 |
7530788 | Boury et al. | May 2009 | B2 |
7591070 | Lee | Sep 2009 | B2 |
7695248 | Mons et al. | Apr 2010 | B2 |
7857587 | Correia et al. | Dec 2010 | B2 |
7922451 | Liang | Apr 2011 | B1 |
7972115 | Potier | Jul 2011 | B2 |
8061987 | Liang | Nov 2011 | B1 |
8075268 | Liang | Dec 2011 | B1 |
8092178 | Marini et al. | Jan 2012 | B2 |
8113779 | Liang | Feb 2012 | B1 |
8182221 | Liang | May 2012 | B1 |
8246307 | Cheong et al. | Aug 2012 | B2 |
8366394 | Liang | Feb 2013 | B1 |
8414265 | Willett, Jr. | Apr 2013 | B2 |
8435004 | Liang | May 2013 | B1 |
20020197159 | Roeloffs | Dec 2002 | A1 |
20030021684 | Downs et al. | Jan 2003 | A1 |
20050232771 | Harvey et al. | Oct 2005 | A1 |
20060120869 | Wilson et al. | Jun 2006 | A1 |
20070237637 | Lee et al. | Oct 2007 | A1 |
20080118363 | Lee et al. | May 2008 | A1 |
20090148305 | Riahi et al. | Jun 2009 | A1 |
20100135822 | Marini et al. | Jun 2010 | A1 |
20100221122 | Klasing et al. | Sep 2010 | A1 |
20120201695 | Little | Aug 2012 | A1 |
Number | Date | Country |
---|---|---|
1281837 | Feb 2003 | EP |
1422383 | May 2004 | EP |
1726783 | Nov 2006 | EP |
1736636 | Dec 2006 | EP |
2434097 | Mar 2012 | EP |
Entry |
---|
EP Search Report, EP 11174595.6-2321 dated May 10, 2011. |
Kwak, JS, et al.; Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade; Turbine Heat Transfer Laboratory, Department of Mechanical Engineering, Texas A&M University, vol. 125, Oct. 2003, Transactions of the ASME, [Retrieved from Internet Jul. 10, 2013] [http://turbomachinery.asmedigitalcollection.asme.org]. |
Ahn, J, et al.; Film-Cooling Effectiveness on a Gas Turbine Blade Tip Using Pressure-Sensitive Paint; Turbine Heat Transfer Laboratory, Department of Mechanical Engineering, Texas A&M University, Journal of Heat Transfer, vol. 127, May 2005, [Retrieved from Internet Jul. 10, 2013] [http://heattransfer.asmedigitalcollection.asme.org]. |
Number | Date | Country | |
---|---|---|---|
20120070307 A1 | Mar 2012 | US |