The invention relates to cooling of a turbine airfoil's trailing edge that can be cast using a ceramic core. Specifically, the invention discloses using angled impingement cooling jets to target cast-in surface features.
Conventional turbine airfoils used in blades and vanes of gas turbine engines have a trailing edge that is thin for aerodynamic efficiency. However, a lack of cooling surface area on the interior makes it difficult to cool the thin trailing edge. The trailing edge is typically cast integrally with the entire blade by using a ceramic core. The features and size of the ceramic core are reflected in the trailing edge. However, core design considerations must be weighed against trailing edge design considerations. For example, larger core features that create impingement channels in the trailing edge are better for core strength, but larger impingement channels mean reduced flow metering. Hence, a well designed core that balances core considerations with trailing edge cooling requirements is a key aspect of a well designed trailing edge design.
Impingement cooling along the mean camber line in a turbine airfoil trailing edge is known. In this arrangement orifices are cast as part of the trailing edge and are oriented with the mean camber line create high-speed impingement jets of cooling fluid. These impingement jets may impinge a surface between adjacent downstream impingement orifices, and this results in an increased heat transfer rate. Single, double, or triple impingement may occur before the spent cooling fluid is exhausted from the trailing edge into the combustion gas path. The series of impingement orifices also act to meter the flow and this provides a more efficient use of the cooling fluid.
By virtue of their location on the mean camber line the impingement orifices are located between the concave interior surface on the suction side and the convex interior surface on the pressure side of the airfoil. Prior cooling schemes have improved heat transfer by angling the impingement orifices such that they produce impingement jets that impinge the concave and convex interior surfaces. This, in turn, cools the respective exterior surfaces of the trailing edge. Other prior cooling schemes place various surface features on the interior surfaces coincident with the impingement jets. However, operating temperatures of gas turbine engines continue to increase. This leaves room in the art for improvements to cooling of the trailing edge.
The invention is explained in the following description in view of the drawings that show:
The present inventors have developed a cooling arrangement for a turbine airfoil's trailing edge, where the trailing edge and all elements of the cooling arrangement are integrally cast with the airfoil using a ceramic casting core. The invention capitalizes on advances in casting core technology to form an arrangement where the elements harmonize to form an unexpectedly extremely efficient cooling arrangement. Specifically, the airfoil and trailing edge are cast around a ceramic casting core configured to form impingement orifices and chevrons within the trailing edge. Some of the impingement orifices direct impingement jets toward chevrons disposed on an interior surface on the pressure side of the airfoil. Other impingement orifices direct impingement jets toward chevrons disposed on an interior surface on the suction side. There may be one or several rows of impingement orifices. Spent impingement air exhausts from the trailing edge into the flow of combustion gases. Compared to impingement jets all pointing in the same direction, alternating the target of the impingement jest from suction side to pressure side not only helps to increase t he surface area being cooled, but it also serves to strengthen the trailing edge section of the ceramic core, thereby increasing production yield while allowing the diameter of the impingement jets to be smaller. Moreover, the use of chevrons not only increases surface area, but it also serves to spread the cooling air in order to more evenly cool the surface and to increase the area being cooled effectively when compared to traditional turbulators or parallel grooves.
Spent cooling fluid 64 from the impingement jet 48 of the first row 22 becomes fresh cooling fluid 40 for the second row 26. The fresh cooling fluid 40 enters the impingement orifice inlet 44 of an impingement orifice 24 of the second row 26, and travels through and exhausts from the impingement orifice 24 via an impingement orifice outlet 46 in the form of an impingement jet 48. The impingement orifice inlet 44 of the impingement orifice 24 of the second row 26 may be at a different elevation than the impingement orifice outlet 46 of the impingement orifice 24 of the first row 22, and hence the impingement orifice 24 of the second row 26 is represented using dotted lines. A center of the impingement orifice outlet 46 of the impingement orifice 24 of the second row 26 is disposed on the pressure side 12 of the mean camber line 20, although it need not necessarily be so long as the respective impingement jet 48 is directed at an angle to the mean camber line 20. In this cross section the impingement jet 48 of the second row 26 is directed toward a target area 60 on a convex second-row interior surface 66 on the pressure side 12. A surface feature (not shown) or plural surfaces features are positioned such that at least a portion of the surface feature is within the target area 60.
Spent cooling fluid from the impingement jet 48 of the second row 26 becomes fresh cooling fluid 40 for the third row 28. The fresh cooling fluid 40 enters the impingement orifice inlet 44 of an impingement orifice 24 of the third row 28, and travels through and exhausts from the impingement orifice 24 via an impingement orifice outlet 46 in the form of an impingement jet 48. The impingement orifice inlet 44 of the impingement orifice 24 of the third row 28 may be at a same elevation as the impingement orifice outlet 46 of the impingement orifice 24 of the first row 22, and hence the impingement orifice 24 of the third row 28 is represented using solid lines. A center of the impingement orifice outlet 46 of the impingement orifice 24 of the third row 28 is disposed on the suction side 14 of the mean camber line 20, although it need not necessarily be so long as the respective impingement jet 48 is directed at an angle to the mean camber line 20. In this cross section the impingement jet 48 of the third row 28 is directed toward a target area 60 on a concave third-row interior surface 68 on the suction side 14. A surface feature (not shown) or plural surfaces features are positioned such that at least a portion of the surface feature is within the target area 60. Spent cooling fluid 64 from the impingement jet 48 of the third row 28 exhausts from the trailing edge portion 30 via the exhaust orifices 32.
In this exemplary embodiment, the rows 22, 26, 28 within this cross section alternate from suction side 14 to pressure side 12 to suction side. It is possible that all three rows 22, 26, 28 in a single cross section may be directed to the same side, or they may point to different sides but not necessarily in an alternating pattern as shown. For example, in an alternate exemplary embodiment the first row 22 and the second row may point to the pressure side 12 while the third row may point to the suction side 14. Any combination may be envisioned. Likewise, the arrangement seen may vary as the location of the cross section is varied from base to tip of the airfoil 10.
Also visible in
The impingement orifices 24 may be circular in cross section, but because they are angled toward the first-row interior surface 62 impingement orifices 24 with a circular cross section form an oval-shaped target area 60. The target area 60 may range in size, and may include smaller 130, mid-range 132, and larger 134 target areas, where the size is relative to how much of the chevron arrangement 120 lies within the target area 60. A shape of a perimeter 136 of the target area 60 may be varied by varying a shape of the cross section of the impingement orifice 24 itself. For example, if the cross section of the impingement orifice were oval with a longer axis oriented in and out of the page, the ovality of the perimeter 136 would be increased from that produced by the impingement orifice having the circular cross section. Conversely, if the ovality of the cross section were oriented such that the longer axis was more parallel to the first-row interior surface 62, then the shape of the perimeter 136 would be more circular. Likewise, by changing the angle of impingement 52 the ovality of the perimeter 136 can be changed. The shape of the cross section of the impingement orifice 24 and the angle of impingement 52 can be manipulated as necessary to achieve whatever shape is desired for the perimeter 136 of the target area 60. In addition, the shape of the perimeter 136 may be the same for all target areas 60, or some or all of the target areas 60 may have their own, unique perimeter shape. These perimeter shapes may be selected to accommodate local cooling requirements and local geometries etc.
Each chevron 122 includes a tip 140 and two wings 142. Adjacent chevrons 122 form a groove 144 there between that may be used to guide the cooling fluid. The tip 140 may be a closed tip 146 or an open tip 148. The wings 142 may be continuous 150 or discontinuous 152. The configuration of chevron arrangements 120 may vary from one chevron arrangement 120 to the next and may be selected to accommodate local cooling requirements and local geometries etc. The chevrons 122 may span an entirety of its target area 60, or the target area 60 may be larger than the span of the chevron 122. The tip 140 of one or all chevrons 122 in chevron arrangement 120 may be disposed within the target area 60.
Spent cooling fluid 64 may flow in the grooves 144 formed by the wings 142 of the chevrons 122. These grooves 144 may be oriented so that they guide the spent cooling fluid 64 along a same path the spent cooling fluid 64 would have taken if the chevron arrangement 120 were not present. In other words, streamlines 160 present in the spent cooling fluid 64 would naturally follow a course if the chevron arrangement 120 were not present. The chevron arrangement 120 can be configured so that the wings 142 and/or the grooves 144 follow the same streamlines as shown in chevron arrangement 162. The result is that the spent cooling fluid 64 will lose little or no energy as a result of the presence of the chevron arrangement 120, but will benefit from the increased surface area created by the chevron arrangement.
Alternately, the wings 142 and/or the grooves 144 can be disposed at an angle to the streamlines 160 that the spent cooling fluid 64 would naturally form, as shown in chevron arrangement 164. This configuration forces the spent cooling fluid 64 to flow over the wings 142 and this creates turbulence, thereby increasing a cooling effect. A length of the wings 142 and a wing angle 166 of the wing 142 to the natural streamline 160 need to be designed to strike a balance between a desire to increase turbulence, and hence increase a cooling efficiency, and a desire to reduce a boundary layer that may form on a downstream side of the wing 142, which forms at longer wing 142 lengths and greater wing angles 166. The wing angle 166 will also determine how well the created turbulence follows the wings 142 and/or grooves, which also affects heat transfer. Similarly, if the wing 142 is discontinuous, a length between gaps 168 needs to be selected to maximize cooling effectiveness by balancing turbulence creation with boundary layer formation.
In an exemplary embodiment the wings 142 and/or the grooves 144 can be configured to guide spend cooling fluid 64 toward the impingement orifice inlet 44 of a subsequent impingement orifice 24. For example, chevron arrangement 180 guides spent cooling fluid 64 from second row impingement orifice 182 toward impingement orifice inlets 44 of third row impingement orifices 184, 186. This may be done to improve flow efficiency through the trailing edge portion 30. Alternately, the wings 142 and/or the grooves 144 can be configured to guide spent cooling fluid 64 toward interstitial structure 188 between impingement orifice 24 should greater fluidic chaos be desired at that location.
In an exemplary embodiment the cooling arrangement may be configured such that a stagnation point 190 within a target area 60 is arranged upstream of the tip 140 of one or all chevrons 122 in the chevron arrangement 120. Doing so ensures spent cooling fluid 64 flows along the wings 142 and/or the grooves 144 away from the tips 140 as opposed to flowing upstream toward the tips 140, which ensures a more uniform flow.
Various cooling arrangements with other surface features and flow paths were considered but this combination of multiple rows of angled impingement on respective chevron ribs provided the greatest heat transfer rate, resulting at least from the increased surface area and increased turbulence, while allowing formation of the trailing edge portion 30 integral to the airfoil 10 using a casting core 82 (which may be made of a ceramic material). The improved heat transfer test results using the arrangement disclosed herein can be seen in
While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.
Development for this invention was supported in part by Contract No. SC0001359 awarded by the United States Department of Energy. Accordingly, the United States Government may have certain rights in this invention.
Number | Name | Date | Kind |
---|---|---|---|
3240468 | Watts et al. | Mar 1966 | A |
4118146 | Dierberger | Oct 1978 | A |
4183716 | Takahara et al. | Jan 1980 | A |
4236870 | Hucul, Jr. et al. | Dec 1980 | A |
4297077 | Durgin et al. | Oct 1981 | A |
4770608 | Anderson et al. | Sep 1988 | A |
5246340 | Winstanley et al. | Sep 1993 | A |
5263820 | Tubbs | Nov 1993 | A |
5356265 | Kercher | Oct 1994 | A |
5702232 | Moore | Dec 1997 | A |
5931638 | Krause et al. | Aug 1999 | A |
6174134 | Lee et al. | Jan 2001 | B1 |
6206638 | Glynn et al. | Mar 2001 | B1 |
6234754 | Zelesky et al. | May 2001 | B1 |
6254346 | Fukuno et al. | Jul 2001 | B1 |
6273682 | Lee | Aug 2001 | B1 |
6439847 | Taeck | Aug 2002 | B2 |
6607356 | Manning et al. | Aug 2003 | B2 |
6779597 | DeMarche et al. | Aug 2004 | B2 |
7056083 | Gray | Jun 2006 | B2 |
7128533 | Liang | Oct 2006 | B2 |
7334992 | Downs et al. | Feb 2008 | B2 |
7670112 | Boury et al. | Mar 2010 | B2 |
7690892 | Liang | Apr 2010 | B1 |
7862299 | Liang | Jan 2011 | B1 |
8052378 | Draper | Nov 2011 | B2 |
8096767 | Liang | Jan 2012 | B1 |
8398370 | Liang | Mar 2013 | B1 |
20030133797 | Dailey | Jul 2003 | A1 |
20040096313 | Harvey et al. | May 2004 | A1 |
20060042255 | Bunker et al. | Mar 2006 | A1 |
20080089787 | Abdel-Messeh et al. | Apr 2008 | A1 |
20090087312 | Bunker et al. | Apr 2009 | A1 |
20120014808 | Lee | Jan 2012 | A1 |
20130142666 | Lee et al. | Jun 2013 | A1 |
Number | Date | Country |
---|---|---|
1079071 | Feb 2001 | EP |
1327747 | Jul 2003 | EP |
1793083 | Jun 2007 | EP |
Entry |
---|
R.Roy, et al, “Designing a Turbine Blade Cooling System Using a Generalised Regression Genetic Algorithm”, CIRP Annals, 2003, vol. 52/1, p. 415-418. |
Number | Date | Country | |
---|---|---|---|
20150118034 A1 | Apr 2015 | US |