None.
Not Applicable.
This invention relates to the cooling of an airfoil comprising a portion of a stator vane or nozzle of the first stage of a gas turbine engine; and more particularly, to the hole pattern formation in the airfoil for thin film cooling of a trailing edge of the airfoil.
In the construction of gas turbine engines, an annular array of turbine segments is provided to form a turbine stage. Generally, the turbine stage is defined by outer and inner annular bands spaced apart from each other with a plurality of vanes or airfoils extending between the bands and circumferentially spaced from one other. This construction, in turn, defines a path for a working fluid flowing through the turbine. In a gas turbine engine, this is a hot gas. As will be appreciated by those skilled in the art, the most extreme adverse operating conditions are generally encountered at the first stage of the turbine. That is because this stage is immediately downstream of the engine's combustion chamber and components comprising this stage must therefore withstand high thermal loads. As is known in the art, cooling systems for this engine stage utilize thin film cooling techniques to insure so adequate cooling is provided. Thin film cooling is accomplished by discharging air through orifices formed in portions of the nozzle. The discharged air then forms a protective thin film boundary layer between the hot stream of gases flowing through the first stage of the turbine and the surface of the nozzle.
Various problems with thin film cooling systems have been encountered and solutions to these problems have been addressed in U.S. Pat. Nos. 6,583,526, 6,561,757, 6,553,665, 6,527,274, 6,517,312, 6,506,013, 6,435,814, 6,402,466, 6,398,486, and 5,591,002, all of which are assigned to the same assignee as the present application.
The present invention is directed to an advanced film-cooling configuration for cooling the trailing edge of a nozzle used in the first stage of an advanced design gas turbine engine. The nozzle is a steam cooled component which operates at firing temperatures which require cooling of the airfoil to extend the low cycle fatigue (LCF), oxidation, and creep life of the component. While steam adequately cools the majority of the nozzle, it is not feasible for use in cooling the trailing edge of the nozzle. Rather, this requires a novel and advanced thin film cooling configuration in order for the trailing edge to not rapidly deteriorate once the turbine is in service which would require costly servicing or replacement of the nozzle and unacceptable down-time when the turbine is out of service.
Briefly stated, the present invention is directed to thin film cooling of the trailing edge of a nozzle for the first stage of a gas turbine engine. Cooling is affected by use of a plurality of rows of film cooling holes located adjacent the trailing edge of the nozzle, on both the concave side and convex side of the nozzle. In particular, three rows of film cooling holes are formed in the sidewalls of the nozzle on the respective concave and convex sides thereof. A first and forward row of holes extends generally longitudinally of the nozzle and comprises holes of varying sizes and angles formed at predetermined locations on the nozzle. Second and third rows of holes also extend generally longitudinally of the nozzle and also comprise holes of varying sizes and angles formed at predetermined locations on the nozzle. The second row of holes comprises a middle row of holes and the third row an aft row. Holes comprising the second row are spaced a substantial distance from those comprising the first row. However, the second and third row of holes are formed relatively close together with the holes comprising the second row being staggered in location with respect to those comprising the third row. By placing the middle and aft rows of holes closer together, and staggering the hole arrangement in these two rows, an effective film flow is achieved which cools the trailing edge of the nozzle thereby to minimize cooling flow, optimize performance of the turbine engine, reduce NOx produced by the engine, prolong the service life of the nozzle and reduce service and repair costs.
Two embodiments of the invention are shown with the thin film cooling arrangement of the first embodiment including substantially more holes in each row than occurs in the second embodiment.
The foregoing and other objects, features, and advantages of the invention will be in part apparent and in part pointed out hereinafter.
In the accompanying drawings which form part of the specification:
Corresponding reference numerals indicate corresponding parts throughout the several figures of the drawings.
The following detailed description illustrates the invention by way of example and not by way of limitation. The description clearly enables one skilled in the art to make and use the invention, describes several embodiments, adaptations, variations, alternatives, and uses of the invention, including what is presently believed to be the best mode of carrying out the invention.
Referring to the drawings, the present invention is directed to thin film cooling for a first stage nozzle assembly, indicated generally 10 in
With respect to
The hole pattern or arrangement of the present invention comprises three rows of openings which extend longitudinally of the airfoil, on both the concave and convex sides of the nozzle assembly, and spaced inwardly of the trailing edge. As particularly shown in
Referring again to
Table 1 is a listing of all the holes comprising rows RA–RC, RJ–RL, and the other holes formed in the bands 14 and 16 and rail 30. The table includes each hole designation, the angle of the opening with respect to the outer surface of airfoil 12, and the X, Y, Z coordinates determining the location of the hole. The distances are measured with respect to the reference point Q (0,0,0) shown in
In
The hole pattern for this embodiment again comprises three rows of openings which extend longitudinally of the airfoil, on both the concave and convex sides of the nozzle assembly, and spaced inwardly of the trailing edge. As particularly shown in
As shown in
Table 2 is a listing of all the holes comprising rows RA′–RC′, RJ′–RL′, and the other holes formed in the curved outer portion of the airfoil and rai 130. The table includes each hole designation, the angle of the opening with respect to the outer surface of airfoil 112, and the X,Y,Z coordinates of the hole locations. As with
In view of the above, it will be seen that the several objects of the invention are achieved and other advantageous results are obtained. As various changes could be made in the above constructions without departing from the scope of the invention, it is intended that all matter contained in the above description or shown in the accompanying drawings shall be interpreted as illustrative and not in a limiting sense.
Number | Name | Date | Kind |
---|---|---|---|
3560107 | Helms | Feb 1971 | A |
5591002 | Cunha et al. | Jan 1997 | A |
6325593 | Darkins et al. | Dec 2001 | B1 |
6398486 | Storey et al. | Jun 2002 | B1 |
6402466 | Burdgick et al. | Jun 2002 | B1 |
6435814 | Yu et al. | Aug 2002 | B1 |
6506013 | Burdgick et al. | Jan 2003 | B1 |
6517312 | Jones et al. | Feb 2003 | B1 |
6527274 | Heron et al. | Mar 2003 | B1 |
6553665 | Gunn et al. | Apr 2003 | B1 |
6561757 | Burdgick et al. | May 2003 | B1 |
6572335 | Kuwabara et al. | Jun 2003 | B1 |
6583526 | Griffith et al. | Jun 2003 | B1 |
Number | Date | Country | |
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20050169746 A1 | Aug 2005 | US |