The present invention generally relates to gas turbine engines, and more particularly relates to air cooled components of gas turbine engines, such as turbine and combustor components.
Gas turbine engines are generally used in a wide range of applications, such as aircraft engines and auxiliary power units. In a gas turbine engine, air is compressed in a compressor, and mixed with fuel and ignited in a combustor to generate hot combustion gases, which flow downstream into a turbine section. In a typical configuration, the turbine section includes rows of airfoils, such as stator vanes and rotor blades, disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled in turn to a main engine shaft. Hot combustion gases are delivered from the engine combustor to the annular hot gas flow path, thus resulting in rotary driving of the rotor disks to provide an engine output.
Due to the high temperatures in many gas turbine engine applications, it is desirable to regulate the operating temperature of certain engine components, particularly those within the mainstream hot gas flow path, in order to prevent overheating and potential mechanical issues attributable thereto. As such, it is desirable to cool the rotor blades and stator vanes in order to prevent damage and extend useful life. One mechanism for cooling turbine airfoils is to duct cooling air through internal passages and then vent the cooling air through holes formed in the airfoil. The holes are typically formed uniformly along a line substantially parallel to the leading edge of the airfoil and at selected distances from the leading edge to provide a film of cooling air over the convex side of the airfoil when the cooling air flows therethrough during engine operation. Other rows of cooling holes or an array of holes may be formed in the airfoil components depending upon design constraints. Film cooling attempts to maintain the airfoils at temperatures that are suitable for their material and stress level.
A typical film cooling hole is a cylindrical aperture inclined axially through one of the airfoil sides. In many conventional engines, however, disadvantageous, relatively high cooling air flows have been used to obtain satisfactory temperature control of engine components.
Accordingly, it is desirable to provide a gas turbine engine with improved film cooling. In addition, it is desirable to provide a air-cooled turbine components with improved hole configurations. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.
In accordance with an exemplary embodiment, an engine component includes a body; and a plurality of cooling holes formed in the body. At least one of the cooling holes has a cross-sectional shape with a first concave portion and a first convex portion.
In accordance with another exemplary embodiment, an engine component, comprising includes a body; and a plurality of cooling holes formed in the body. At least one of the cooling holes has a triangle cross-sectional.
The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein:
The following detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any theory presented in the preceding background or the following detailed description.
Broadly, exemplary embodiments discussed herein include gas turbine engines with turbine components having improved film cooling. The turbine components have a number of non-circular cooling holes. The cooling holes may have, for example, both convex and concave portions. For example, the cooling holes can have cross-sectional shapes such as: bean-shaped, triad-shaped, reverse B-shaped, dumbbell shaped, and/or triangle-shaped.
The stator vanes 122 project radially outwardly from a circumferential platform 126 to the annular duct wall 104. The rotor blades 124 project radially outwardly from a circumferential platform 128 that is adapted for appropriate connection to the rotor disk (not shown) at the periphery thereof. The rotor disk is generally positioned within the internal engine cavity and is coupled to a main engine shaft for rotation therewith. As shown, the rotor blade 124 and stator vane 122 may form one stage of a multistage turbine. As such, multiple rows of the stator vanes 122 and the rotor blades 124 may be provided in the gas turbine section 100, with the rotor blades 124 and associated rotor disks being rotatably driven by the hot gas flowing through the mainstream hot gas flow path 106 for power extraction. A supply of cooling air, typically obtained as a bleed flow from the compressor (not shown), may pass through cooling holes in the airfoils 122, 124 to form a surface cooling film. Although the cooling holes are discussed with reference to turbine components, the cooling holes may also be incorporated into other engine components, such as combustor components. The cooling holes are discussed in greater detail below.
The airfoil 200 generally has a body 201 with a leading edge 202 and an opposite trailing edge 204. The airfoil 200 also includes a pressure sidewall 206 that is generally concave and an opposite, suction sidewall 208 that is generally convex and is spaced-apart from the pressure sidewall 206. The pressure sidewall 206 and suction sidewall 208 extend from leading edge 202 to trailing edge 204. The airfoil 200 has a hollow interior cavity 210 such that the airfoil 200 has an inner surface 212 and an outer surface 214. Airfoils 200 used in high performance gas turbine engines, such as those used for aircraft propulsion, can be made from high heat and high stress resistant aerospace alloys, such as nickel based alloys, Rene 88, Inconel 718, single crystal materials, steels, titanium alloys or the like.
As noted above, the airfoil 200 is subject to extremely high temperatures because high velocity hot gases are ducted from the combustor (not shown) onto the airfoil 200. If unaddressed, the extreme heat may affect the useful life of an airfoil. As such, film cooling is provided for the airfoil 200 to provide a cooling film of fluid onto the surface of the airfoil 200, particularly in the area of the leading edge 202 and areas immediately aft of the leading edge 202. As noted above, cooling air is bled from the compressor (not shown) or other source and passes into the interior cavity 210 and through cooling holes 220 to the outer surface 214 of the airfoil 200. The cooling holes 220 are formed at locations on the airfoil 200, particularly the convex side 206, concave side 208, and leading edge 202, to provide optimum cooling of the engine component.
The cooling holes 220 may be formed in a selected pattern or array to provide optimum cooling. The cooling holes 220 may be disposed at any angle relative to the outer surface 206, such as about 20° to about 40°, although the cooling holes 220 may be oriented at lesser or greater angles. Computational fluid dynamic (CFD) analysis can additionally be used optimize the location and orientation of the cooling holes 220. The cooling holes 220 may be formed by casting, abrasive water jet, Electron Discharge Machining (EDM), laser drilling, or any suitable process.
In general, the cooling holes 220 may be considered to have an upstream portion 222 adjacent the inner surface 212 and a downstream portion 224 adjacent the outer surface 214. The upstream portion of each cooling hole 220, lying closer to the inner surface 212 is substantially cylindrical or circular and the downstream portion lying closer to the outer surface 214 may have a cross-sectional shape as discussed below with reference to
The cooling hole 300 may be considered to have an x-axis 380 and a y-axis 390, as shown in
The cooling hole 400 may be considered to have an x-axis 480 and a y-axis 490, as shown in
The cooling hole 500 may be considered to have an x-axis 580 and a y-axis 590, as shown in
The cooling hole 600 may be considered to have an x-axis 680 and a y-axis 690, as shown in
The cooling hole 700 may be considered to have an x-axis 780 and a y-axis 790, as shown in
In general, the cross-sectional shapes of the holes 220, 300, 400, 500, 600, 700 facilitate the distribution of the cooling air substantially completely over the outer surface of the airfoil. In particular, the cross-sectional shapes function as a diffuser to reduce the velocity and increase static pressure of the cooling airstreams exiting the holes and encourage cooling film development. The holes 220, 300, 400, 500, 600, 700 additionally increase the lateral spread distribution of the exiting airflows, decrease peak velocities, and improve adiabatic effectiveness across a number of blowing ratios. These airstreams are more inclined to cling to the surface for improved cooling rather than separate from the surface. This produces an enhanced cooling effect at the surface. Consequently, exemplary embodiments promote the service life of the airfoil (e.g., airfoils 122, 124, 200) as a result of a more uniform cooling film at the external surfaces.
For example, the cooling hole 300 of
Exemplary embodiments disclosed herein are generally applicable to air-cooled components, and particularly those that are to be protected from a thermally and chemically hostile environment. Notable examples of such components include the high and low pressure turbine nozzles and blades, shrouds, combustor liners and augmentor hardware of gas turbine engines. Additionally, the cooling holes discussed above may be incorporated into turbine components. The advantages are particularly applicable to gas turbine engine components that employ internal cooling to maintain the service temperature of the component at an acceptable level while operating in a thermally hostile environment.
While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, 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 invention 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 invention. 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 invention as set forth in the appended claims.
Number | Name | Date | Kind |
---|---|---|---|
4461612 | Dodd | Jul 1984 | A |
4526358 | Ura et al. | Jul 1985 | A |
4529358 | Papell | Jul 1985 | A |
4653983 | Vehr | Mar 1987 | A |
4664597 | Auxier et al. | May 1987 | A |
5096379 | Stroud et al. | Mar 1992 | A |
5261223 | Foltz | Nov 1993 | A |
5281084 | Noe et al. | Jan 1994 | A |
5382133 | Moore et al. | Jan 1995 | A |
5465572 | Nicoll et al. | Nov 1995 | A |
5496151 | Coudray et al. | Mar 1996 | A |
5511937 | Papageorgiou | Apr 1996 | A |
5609779 | Crow et al. | Mar 1997 | A |
5683600 | Kelley et al. | Nov 1997 | A |
5747769 | Rockstroh et al. | May 1998 | A |
6243948 | Lee et al. | Jun 2001 | B1 |
6329015 | Fehrenbach et al. | Dec 2001 | B1 |
6368060 | Fehrenbach et al. | Apr 2002 | B1 |
6420677 | Emer et al. | Jul 2002 | B1 |
6554571 | Lee et al. | Apr 2003 | B1 |
6607355 | Cunha et al. | Aug 2003 | B2 |
6616406 | Liang | Sep 2003 | B2 |
6979176 | Nakamata et al. | Dec 2005 | B2 |
6984100 | Bunker et al. | Jan 2006 | B2 |
7008186 | Heeg et al. | Mar 2006 | B2 |
7131814 | Nagler et al. | Nov 2006 | B2 |
7186085 | Lee | Mar 2007 | B2 |
7186091 | Lee et al. | Mar 2007 | B2 |
7246992 | Lee | Jul 2007 | B2 |
7249933 | Lee et al. | Jul 2007 | B2 |
7328580 | Lee et al. | Feb 2008 | B2 |
7351036 | Liang | Apr 2008 | B2 |
7374401 | Lee | May 2008 | B2 |
7540712 | Liang | Jun 2009 | B1 |
7997867 | Shih et al. | Aug 2011 | B1 |
7997868 | Liang | Aug 2011 | B1 |
8057179 | Liang | Nov 2011 | B1 |
8057180 | Liang | Nov 2011 | B1 |
8057181 | Liang | Nov 2011 | B1 |
20050023249 | Kildea | Feb 2005 | A1 |
20050135931 | Nakamata et al. | Jun 2005 | A1 |
20050232768 | Heeg et al. | Oct 2005 | A1 |
20060104807 | Lee | May 2006 | A1 |
20060171807 | Lee | Aug 2006 | A1 |
20060272335 | Schumacher et al. | Dec 2006 | A1 |
20060277921 | Patel et al. | Dec 2006 | A1 |
20070006588 | Patel et al. | Jan 2007 | A1 |
20070128029 | Liang | Jun 2007 | A1 |
20070234727 | Patel et al. | Oct 2007 | A1 |
20080003096 | Kohli et al. | Jan 2008 | A1 |
20080005903 | Trindade et al. | Jan 2008 | A1 |
20080031738 | Lee | Feb 2008 | A1 |
20080271457 | McMasters et al. | Nov 2008 | A1 |
20090246011 | Itzel | Oct 2009 | A1 |
20100040459 | Ohkita | Feb 2010 | A1 |
20100124484 | Tibbott et al. | May 2010 | A1 |
20100303635 | Townes et al. | Dec 2010 | A1 |
20110097188 | Bunker | Apr 2011 | A1 |
20110217181 | Hada et al. | Sep 2011 | A1 |
20110268584 | Mittendorf | Nov 2011 | A1 |
20110311369 | Ramachandran et al. | Dec 2011 | A1 |
Number | Date | Country |
---|---|---|
0375175 | Nov 1989 | EP |
0924384 | Jun 1999 | EP |
0992653 | Apr 2000 | EP |
1609949 | Dec 2005 | EP |
1892375 | Feb 2008 | EP |
1942251 | Jul 2008 | EP |
1970628 | Sep 2008 | EP |
07332005 | Dec 1995 | JP |
2001012204 | Jan 2001 | JP |
2006307842 | Nov 2006 | JP |
Entry |
---|
Kusterer et al., Double-Jet Film-Cooling for Highly Efficient Film-Cooling with Low Blowing Ratios, Proceedings of ASME Turbo Expo 2008: Power for Land, Sea and Air GT2008, Jun. 9-13, 2008, pp. 1-12, Berlin, Germany, GT2008-50073. |
Wayne et al., High-Resolution Film Cooling Effectiveness Comparison of Axial and Compound Angle Holes on the Suction Side of a Turbine Vane, Transactions of the ASME, pp. 202-211, Copyright 2007 by ASME. |
Lu et al., Turbine Blade Showerhead Film Cooling: Influence of Hole Angle and Shaping, International Journal of Heat and Fluid Flow 28 (2007) pp. 922-931. |
Kim et al., Influence of Shaped Injection Holes on Turbine Blade Leading Edge Film Cooling, International Journal of Heat and Mass Transfer 47 (2004) pp. 245-256. |
Ramachandran et al., Turbine Engine Components, filed with the USPTO on Jun. 24, 2009 and assigned U.S. Appl. No. 12/490,840. |
EP Search Report, EP10187079.8-2321 dated Apr. 2, 2011. |
Loh, Teck Seng; Srigrarom, Sutthiphong; Investigative Study of Heat Transfer and Blades Cooling in the Gas Turbine, The Smithsonian/NASA Astrophysics Data System; Modern Physics Letters B, vol. 19, Issue 28-29, pp. 1611-1614 (2005). |
Loh, Teck Seng; Srigrarom, Sutthiphong; Investigative Study of Heat Transfer and Blades Cooling in the Gas Turbine, Modern Physics Letters B, vol. 19, Issue 28-29, pp. 1611-1614 (2005). |
Ronald S. Bunker; A Review of Shaped Hole Turbine Film-Cooling Technology; Journal of Heat Transfer, Apr. 2005, vol. 127, Issue 4, 441 (13 pages). |
Shih, T. I.-P., Na, S.; Momentum-Preserving Shaped Holes for Film Cooling; ASME Conference Proceedings, Year 2007, ASME Turbo Expo 2007: Power for Land, Sea, and Air (GT2007), May 14-17, 2007, Montreal, Canada; vol. 4: Turbo Expo 2007, Parts A and B; Paper No. GT2007-27600, pp. 1377-1382. |
Yiping Lu; Effect of Hole Configurations on Film Cooling From Cylindrical Inclined Holes for the Application to Gas Turbine Blades, A Dissertation, Submitted to the Graduate Faculty of the Louisiana State University and Agricultural and Mechanical College, Dec. 2007. |
Colban, W., Thole, K.; Influence of Hole Shape on the Performance of a Turbine Vane Endwall Film-cooling Scheme, International Journal of Heat and Fluid Flow 28 (2007), pp. 341-356. |
Gartshore, I., Salcudean, M., Hassan, I.: Film Cooling Injection Hole Geometry : Hole Shape Comparison for Compound Cooling Orientation, American Institute of Aeronautics and Astronautics, Reston, VA, 2001, vol. 39, No. 8, pp. 1493-1499. |
Okita, Y., Nishiura, M.: Film Effectiveness Performance of an Arrowhead-Shaped Film Cooling Hole Geometry, ASME Conference Proceedings, ASME Turbo Expo 2006: Power for Land, Sea, and Air (GT2006), May 8-11, 2006 , Barcelona, Spain, vol. 3: Heat Transfer, Parts A and B, No. GT2006-90108, pp. 103-116. |
Lu, Y., Allison, D., Ekkad, S. V.: Influence of Hole Angle and Shaping on Leading Edge Showerhead Film Cooling, ASME Turbo Expo 2006: Power for Land, Sea, and Air (GT2006), May 8-11, 2006 , Barcelona, Spain, vol. 3: Heat Transfer, Parts A and B, No. GT2006-90370 pp. 375-382. |
Heidmann et al., A Novel Antivortex Turbine Film-Cooling Hole Concept, Journal of Turbomachinery, 2008 by ASME, Jul. 2008, vol. 130, pp. 031020-1-031020-9. |
Malak, F.M., et al.; Gas Turbine Engine Components With Film Cooling Holes Having Cylindrical to Multi-Lobe Configurations, U.S. Appl. No. 13/465,647, filed May 7, 2012. |
Number | Date | Country | |
---|---|---|---|
20110123312 A1 | May 2011 | US |