Patent Grant
9169733
The invention described herein relates generally to a turbine airfoil assembly. More specifically, the invention relates to a turbine airfoil assembly configured for improved cooling performance.
Turbine airfoil assemblies direct gaseous flow passing through rotor assemblies within a gas turbine. For example, a stator vane assembly may include one or more stator vane airfoils extending radially between an inner and an outer platform. The temperature of core gas flow passing the stator vane airfoil typically requires cooling within the stator vane, and this cooling helps to increase stator vane life.
In many gas turbines, some components must be cooled to extend operating life. Cooling air at a lower temperature and higher pressure than the core gas is typically introduced into an internal cavity of a stator vane, where it absorbs thermal energy. The cooling air subsequently exits the vane via apertures in the vane walls, transporting the thermal energy away from the vane. The pressure difference across the vane walls and the flow rate at which the cooling air exits the vane is important, particularly along the leading edge where temperatures may be elevated. In the past, internal vane structures have been defined by first establishing the minimum acceptable pressure difference at any point along the leading edge (internal versus external pressure), and subsequently manipulating the internal vane structure along the entire leading edge such that the minimal allowable pressure difference is present along the entire leading edge. The problem with this approach is that core gas flow pressure gradients along the leading edge of a vane may have one or more small regions (i.e., “spikes”) at a pressure considerably higher than the rest of the gradient along the leading edge. This is particularly true for those stator vanes disposed aft of rotor assemblies, where relative motion between rotor blades and stator vanes can significantly influence the core gas flow profile. Increasing the minimum allowable pressure to accommodate the spikes consumes an excessive amount of cooling air.
Prior approaches have modified the internal vane structure, but this approach does not permit customization. Turbines may be installed in a wide variety of locations (e.g., hot, cold, dry, humid, etc.) and the same turbine in a very cold and humid environment may experience a very different core gas flow pressure gradient than a turbine installed in a hot and dry environment.
In an aspect of the present invention, a turbine airfoil assembly has an airfoil with an inner wall, an outer wall, a leading edge and a trailing edge. The airfoil has one or more chambers extending in a substantially chordwise direction of the airfoil. An insert has a plurality of impingement holes, and the insert is configured to be inserted within one of the chambers. The insert is configured to cool the airfoil via the plurality of impingement holes. A chambering element is attached only to the insert, the chambering element is configured to provide an increased cooling gas pressure inside a boundary area defined by the chambering element relative to an area outside the boundary area. A gap exists between the inner wall of the airfoil and the chambering element, and the gap allows cooling gas to exit the boundary area and enter the area outside the boundary area.
In another aspect of the present invention, a turbine airfoil assembly has an airfoil with an inner wall. The airfoil has one or more chambers extending in a substantially chordwise direction of the airfoil. An insert includes a plurality of impingement holes, and the insert is configured to be inserted within one of the chambers. The insert is configured to cool the airfoil via the plurality of impingement holes. A chambering element is attached only to the insert or only to the airfoil. The chambering element is configured to provide an increased cooling gas pressure inside a boundary area defined by the chambering element relative to an area outside the boundary area. A gap exists between the chambering element and the inner wall of the airfoil or the insert. The gap allows cooling gas to exit the boundary area and enter the area outside the boundary area.
One or more specific aspects/embodiments of the present invention will be described below. In an effort to provide a concise description of these aspects/embodiments, all features of an actual implementation may not be described in the specification. It should be appreciated that in the development of any such actual implementation, as in any engineering or design project, numerous implementation-specific decisions must be made to achieve the developers' specific goals, such as compliance with machine-related, system-related and business-related constraints, which may vary from one implementation to another. Moreover, it should be appreciated that such a development effort might be complex and time consuming, but would nevertheless be a routine undertaking of design, fabrication, and manufacture for those of ordinary skill having the benefit of this disclosure.
When introducing elements of various embodiments of the present invention, the articles “a,” “an,” “the,” and “said” are intended to mean that there are one or more of the elements. The terms “comprising,” “including,” and “having” are intended to be inclusive and mean that there may be additional elements other than the listed elements. Any examples of operating parameters and/or environmental conditions are not exclusive of other parameters/conditions of the disclosed embodiments. Additionally, it should be understood that references to “one embodiment”, “one aspect” or “an embodiment” or “an aspect” of the present invention are not intended to be interpreted as excluding the existence of additional embodiments or aspects that also incorporate the recited features.
The chambers 111, 113 may be configured to accept an insert (not shown in
To counteract regions of high core gas pressure, a chambering element 240 is attached to the insert 221 and is configured to provide an increased cooling gas pressure inside the boundary area 250 defined by the chambering element 240 relative to an area 260 outside the boundary area 250. The boundary area 250 is the region of space inside the chambering element border, and the area 260 is the region of space external to the boundary area 250. The increased internal pressure in boundary area 250 may also help if a crack occurred in the airfoil wall, in the location of high external pressures, because the hot core gas will not be ingested through the crack (due to the increased internal pressure) which may cause a structural failure of the airfoil. The chambering element 240 may be comprised of a wire, or physical member that partially isolates the inner region 250 from the outer region 260. The chambering element 240 may be attached to the insert 221 by welding, brazing, a mechanical connection or by adhesive.
A gap 275 exists between the inner wall 231 and the insert 221. Post impingement cooling gas travels along this gap and then exits the airfoil 210. A plurality of standoffs 270 may be configured to maintain this gap. The standoffs are attached to the insert 221 (e.g., by welding) or cast into the inner wall 231 and have a predetermined height and/or spacing. For example, the desired gap may be 2 mm, so the height of one or more standoffs 221 may be about 2 mm.
The turbine airfoil assembly 100, according to an aspect of the present invention, could be configured for use as a bucket, blade, nozzle, a shroud or vane in a gas turbine, steam turbine, or any other turbomachinery component that requires cooling. As mentioned previously, gas turbines and steam turbines (or any other turbomachine or turbo-engine) operate in widely varying environmental conditions and the fuel used may also vary greatly. It would be highly beneficial to be able to “customize” each turbine to its individual operating and environmental conditions, and this was not possible in the past. The present invention now enables the turbomachine to be quickly customized or repaired so that any problem areas (e.g., hot spots on airfoils) can be configured so that additional cooling gas can be directed and maintained in the areas that need it most.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
| Number | Name | Date | Kind |
|---|---|---|---|
| 3963368 | Emmerson | Jun 1976 | A |
| 4403917 | Laffitte | Sep 1983 | A |
| 4461612 | Dodd | Jul 1984 | A |
| 5741117 | Clevenger | Apr 1998 | A |
| 6468031 | Yu et al. | Oct 2002 | B1 |
| 6733229 | Burdgick et al. | May 2004 | B2 |
| 7497655 | Liang | Mar 2009 | B1 |
| 7556476 | Liang | Jul 2009 | B1 |
| 20060056968 | Jacala et al. | Mar 2006 | A1 |
| 20090087312 | Bunker et al. | Apr 2009 | A1 |
| 20100221123 | Pal et al. | Sep 2010 | A1 |
| 20110027102 | Nakamata | Feb 2011 | A1 |
| Number | Date | Country | |
|---|---|---|---|
| 20140286762 A1 | Sep 2014 | US |
