Component Cooling Through Seals

Abstract

The present application thus provides for a turbine component cooling system. The component cooling system may include a turbine component, an airflow passing adjacent to the turbine component, and a seal positioned adjacent to the turbine component. The seal may include a number of apertures so as to allow the airflow to pass therethrough.

Description

TECHNICAL FIELD

[0001] The present application relates generally to gas turbine engines and more particularly relates through cooling hot engine components via airflow through seal components.

BACKGROUND OF THE INVENTION

[0002] A portion of the airflow through a turbine engine may be diverted and used for cooling purposes. The overall efficiency of the turbine engine, however, is decreased by the amount of air that is diverted for cooling purposes as opposed to being used for combustion. The less airflow diverted for cooling purposes or other types of parasitic airflows, the better the efficiency and operation of the gas turbine engine as a whole.

[0003] By way of example, a retainer ring positioned about a stage one nozzle outer base and a stage one shroud may be cooled by a compressor discharge airflow. The retainer ring may include a number of circumferential grooves and radial slots therein. The retainer ring may be cooled with a cooling flow from the core airflow. Circumferential cooling of the retaining ring, however, may cause a circumferential temperature gradient therein. Moreover, machining the grooves in the retaining ring requires machining time and labor costs.

[0004] There is thus a desire for improved systems and methods for component cooling that involves less airflow while increasing overall system efficiency. The systems and methods preferably also allow the use of less complicated and costly components

SUMMARY OF THE INVENTION

[0005] The present application thus provides for a turbine component cooling system. The component cooling system may include a turbine component, an airflow passing adjacent to the turbine component, and a seal positioned adjacent to the turbine component. The seal may include a number of apertures so as to allow the airflow to pass therethrough.

[0006] The present application further provides for a turbine component cooling system. The component cooling system may include a turbine component, an airflow passing adjacent to the turbine component, and a W-seal positioned adjacent to the turbine component. The W-seal may include a number of apertures so as to allow the airflow to pass therethrough.

[0007] The present application further provides for a turbine component cooling system. The component cooling system may include a stage one nozzle outer base, a retaining ring positioned adjacent to the a stage one nozzle outer base, a stage one shroud positioned adjacent to the retaining ring, an airflow passing adjacent to the retaining ring along an air sealing plate, and a seal positioned between the air sealing plate and the stage one shroud. The seal may include a number of apertures so as to allow the airflow to pass therethrough.

[0008] These and other features of the present application will become apparent to one of ordinary skill in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

[0009] FIG. 1 is a side cross-sectional view of a gas turbine engine.

[0010] FIG. 2 is a side cross-sectional view of the inner section of a stage one nozzle outer base and a stage one shroud showing a retaining ring.

[0011] FIG. 3 is a side cross-sectional view of the inner section of a stage one nozzle outer base and a stage one shroud showing a retaining ring and a W-seal as is described herein.

[0012] FIG. 4 is a side plan view of the W-seal as may be used herein.

[0013] FIG. 5 is a cross-sectional view of the W-seal.

DETAILED DESCRIPTION

[0014] Referring now to the drawings, in which like numbers refer to like elements throughout the several views, FIG. 1 shows a cross-sectional view of a gas turbine engine 100. As is known, the gas turbine engine 100 may include a compressor 110 to compress an incoming flow of air. The compressor 110 delivers the compressed flow of air to a combustor 120. The combustor 120 mixes the compressed flow of air with a compressed flow of fuel and ignites the mixture. The hot combustion gases are in turn delivered to a turbine 130. The hot combustion gases drive a number of turbine buckets 140 so as to produce mechanical work. The mechanical work produced in the turbine 130 drives the compressor 110 and also an external load such as an electrical generator and the like. The gas turbine engine 100 may use natural gas, various types of syngas, and other types of fuels. Other types of gas turbine engines also may be used herein. The gas turbine engine 10 may have other configurations and may use other types of components. Multiple gas turbine engines 100, other types of turbines, and other types of power generation equipment may be used herein together.

[0015] FIG. 2 is an expanded view of a portion of the turbine 130. Specifically, the intersection of a retaining ring 145 as positioned about a stage one nozzle outer base 150 and a stage one shroud 160 as positioned within a turbine shell 170. An air sealing plate 180 may be positioned between the retaining ring 145 and the stage one shroud 160. The air sealing plate 180 faces a number of circumferential grooves 190 positioned within the retaining ring 145 on one side and one or more W-seals 200 on the other side. The retaining ring 145 with the air sealing plate 180 may be cooled by an airflow 195, in this case a compressor discharge airflow 195 passing through the circumferential groves 190. Such circumferential air cooling, however, may cause a temperature gradient along the retaining ring 145 of up to about 20 degrees Fahrenheit (about 6.7 degrees Celsius) or so.

[0016] FIG. 3 shows a component cooling system 210 as is described herein. The component cooling system 210 may include a turbine component 220. The turbine component 220 may be a retaining ring 225. In this example, the retaining ring 225 includes an air sealing plate 230 but not the circumferential grooves 190. The component cooling system 210 also may include one or more W-seals 240 positioned against the air sealing plate 210. As is shown in FIGS. 4 and 5, the W-seal 240 may include a number of apertures 250 therein so as to permit a cooling airflow 255 therethrough. The apertures 250 may be about 0.07 inches (about 1.78 millimeters) in diameter and about 300 in number. The size and number of the apertures 250 may be varied with the desired cooling flow rate.

[0017] The W-seal 240 may be made out of Inconel 718 or similar types of materials. (Inconel 718 is a nickel chromium alloy made precipitation hardenable by additions of aluminum and titanium and having creep rupture strength at high temperatures to about 1290 degrees Fahrenheit (about 700 degrees Celsius)). Inconel is a trademark of Huntington Alloys Corporation of Huntington, W.V. Other types or combinations of materials may be used herein. The Inconel material may have a thickness of about 0.01 inches (about 0.254 millimeters).

[0018] The airflow 255 through the W-seal 240 may be about 0.215% W25. The use of the W-seals 240 may provide more uniform cooling of the retaining ring 225. Material and labor costs in producing circumferential grooves 190 also are eliminated. Likewise, eliminating the circumferential grooves 190 makes the retaining ring 225 structurally stronger. The W-seal 240 thus both meters the cooling airflow 255 therethrough while acting as the seal between the retaining ring 145 and the stage one shroud 160. The W-seals 240 described herein further may be used anywhere a cooling flow may be required. Other types of seals and other types of turbine components 220 may be used with the component cooling system 210 described herein.

[0019] It should be apparent that the foregoing relates only to certain embodiments of the present application and that numerous changes and modifications may be made herein by one of ordinary skill in the art without departing from the general spirit and scope of the invention as defined by the following claims and the equivalents thereof.

Claims

1. A turbine component cooling system, comprising: a turbine component; an airflow passing adjacent to the turbine component; a seal positioned adjacent to the turbine component; and the seal comprising a plurality of apertures so as to allow the airflow to pass therethrough.

2. The turbine component cooling system of claim 1, wherein the turbine component comprises a retaining ring.

3. The turbine component cooling system of claim 2, further comprising an air sealing plate positioned adjacent to the retaining ring.

4. The turbine component cooling system of claim 3, wherein the seal is positioned about the air sealing plate and a stage one shroud.

5. The turbine component cooling system of claim 1, wherein the seal comprises a W-seal.

6. The turbine component cooling system of claim 1, wherein the plurality of apertures comprises about 300 apertures.

7. The turbine component cooling system of claim 1, wherein the seal comprises a nickel chromium alloy.

8. The turbine component cooling system of claim 1, wherein the airflow comprises about 0.215% W25.

9. A turbine component cooling system, comprising: a turbine component; an airflow passing adjacent to the turbine component; a W-seal positioned adjacent to the turbine component; and the W-seal comprising a plurality of apertures so as to allow the airflow to pass therethrough.

10. The turbine component cooling system of claim 9, wherein the turbine component comprises a retaining ring with an air sealing plate.

11. The turbine component cooling system of claim 9, wherein the seal comprises a nickel chromium alloy.

12. The turbine component cooling system of claim 9, wherein the airflow comprises about 0.215% W25.

13. A turbine component cooling system, comprising: a stage one nozzle outer base; a retaining ring positioned adjacent to the stage one nozzle outer base; a stage one shroud positioned adjacent to the retaining ring; an airflow passing adjacent to the retaining ring along an air sealing plate; a seal positioned between the air sealing plate and the stage one shroud, and the seal comprising a plurality of apertures so as to allow the airflow to pass therethrough.

14. The turbine component cooling system of claim 13, wherein the seal comprises a W-seal.

15. The turbine component cooling system of claim 13, wherein the plurality of apertures comprises about 300 apertures.

16. The turbine component cooling system of claim 13, wherein the seal comprises a nickel chromium alloy.

17. The turbine component cooling system of claim 13, wherein the airflow comprises about 0.215% W25.