System For Reducing Combustion Noise And Improving Cooling

Abstract

A novel and improved system for cooling and reducing combustion noise in a gas turbine combustor is disclosed. The system includes a flow sleeve for a gas turbine combustor comprising a tubular portion and a conical portion, with a plurality of flow straighteners extending radially inward and between the tubular and conical portions and a plurality of rows of cooling holes extending about the flow sleeve, and an aft ring secured to the outlet end of the conical portion, where the aft ring includes an overlapping piston ring that is able to expand or contract in diameter. A combustion liner extends through the flow sleeve and engages an inlet of a transition duct while the piston ring of the flow sleeve also engages the transition duct to form a seal.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] None.

TECHNICAL FIELD

[0002] The present invention generally relates to a system for reducing combustion noise and directing cooling air into a gas turbine combustor.

BACKGROUND OF THE INVENTION

[0003] In a typical gas turbine engine used in a powerplant application, a plurality of combustors are arranged in an annular array about a centerline of the engine. The combustors receive pressurized air from the engine's compressor, add fuel to create a fuel/air mixture, and ignite the mixture to produce hot combustion gases. The hot combustion gases exit the combustors and enter a turbine, where the expanding gases are utilized to drive a turbine, which is in turn coupled through a shaft to the compressor. The engine shaft is also coupled to a shaft that drives a generator for generating electricity.

[0004] The combustors typically include at least a pressurized case and a combustion liner contained within the case. The fuel, which is supplied by a plurality of fuel nozzles, mixes with air and reacts (i.e. ignites) within the combustion liner. In order to actively cool the combustion liner, the compressed air that is used for combustion is first directed through the pressurized case and along the combustion liner. The air then mixes with fuel and reacts in the combustion liner.

[0005] Prior art configurations of combustors include a flow sleeve extending through the case and used to support a combustion liner in place within the combustor. The flow sleeve often includes a series of holes through which compressed air passes. Air passing through these holes is intended to impinge on the combustion liner wall. However, in prior gas turbine combustor configurations, air streams have been known to be ineffective in maintaining active cooling through impingement, thereby leading to premature degradation and damage of the combustion liner.

SUMMARY

[0006] In accordance with the present invention, there is provided a novel and improved system for cooling and reducing combustion noise in a gas turbine combustor. An embodiment of the present invention includes a flow sleeve for a gas turbine combustor comprising a tubular portion and a conical portion, with a plurality of flow straighteners extending radially inward and between the tubular and conical portions. The flow sleeve also includes a plurality of rows of cooling holes extending about the flow sleeve and an aft ring secured to the outlet end of the conical portion, where the aft ring includes a receptacle containing a piston ring that is able to expand or contract in diameter.

[0007] In another embodiment of the present invention, a cooling system for a gas turbine combustor is disclosed comprising a flow sleeve, a combustion liner, and a transition duct. The flow sleeve includes a tubular portion, a conical portion, a plurality of rows of cooling holes, and an aft ring having a receptacle and a piston ring contained within the receptacle. A cooling passage is formed between the combustion liner and the flow sleeve and directs air received from the flow sleeve to the inlet of the combustor. The transition duct is coupled to the combustion liner at the liner outlet. A piston ring is positioned in the aft end of the flow sleeve to create a seal between the flow sleeve such that compressor air is only permitted to enter the passageway through holes in the flow sleeve.

[0008] In yet another embodiment of the present invention, a sealing device for an aft end of a gas turbine combustor, the device comprising a sleeve having a conical portion and an aft ring with a receptacle. A piston ring is located in the receptacle, where the piston ring is able to expand or contract in order to form a seal with the transition duct.

[0009] Additional advantages and features of the present invention will be set forth in part in a description which follows, and in part will become apparent to those skilled in the art upon examination of the following, or may be learned from practice of the invention. The instant invention will now be described with particular reference to the accompanying drawings.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

[0010] The present invention is described in detail below with reference to the attached drawing figures, wherein:

[0011] FIG. 1 is a cross section view of a portion of a gas turbine combustor in accordance with the prior art;

[0012] FIG. 2 is a detailed cross section view of a portion of the gas turbine combustor of FIG. 1 in accordance with the prior art;

[0013] FIG. 3 is a cross section view of a portion of a gas turbine combustor in accordance with an embodiment of the present invention;

[0014] FIG. 4 is a detailed cross section view of a portion of the gas turbine combustor of FIG. 3 in accordance with an embodiment of the present invention; and,

[0015] FIG. 5 is a cross section view of a flow sleeve depicted in FIG. 3 in accordance with an embodiment of the present invention.

DETAILED DESCRIPTION

[0016] The subject matter of the present invention is described with specificity herein to meet statutory requirements. However, the description itself is not intended to limit the scope of this patent. Rather, the inventors have contemplated that the claimed subject matter might also be embodied in other ways, to include different components, combinations of components, steps, or combinations of steps similar to the ones described in this document, in conjunction with other present or future technologies.

[0017] The present invention is directed generally towards a system for reducing combustion noise controlling the flow of cooling air to a combustion liner. Referring initially to FIGS. 1 and 2, a portion of a gas turbine combustor according to the prior art is shown in cross section. A flow sleeve 100 includes a combustion liner 102 contained therein. The combustion liner 102 engages a double wall transition duct 104 for purposes of passing the flow of hot combustion gases from the combustion liner 102 to an inlet of a turbine (not shown). In the prior art configuration, there existed a piston ring in an annular gap 106 at the aft end of the flow sleeve 100 and the inlet of transition duct 104. The piston ring serves to minimize the flow of air through the gap 106 created by the flow sleeve 100 and transition duct 104. However, the amount of cooling air passing through the annular gap 106 is not minimized due to a large opening in the piston ring that is created at operating temperatures, thereby affecting the even distribution of air flow to the head end of the combustor.

[0018] The present invention can be better understood when considering FIGS. 3-5. Referring initially to FIG. 3, a cooling system for a gas turbine combustor 200 is disclosed providing improved control of the flow of cooling air thereby also providing improved control of combustion noise due to variations in rates of air provided for mixing with fuel and resulting combustion.

[0019] Referring now to FIGS. 3-5, the gas turbine combustor 200 comprises a flow sleeve 202 having a tubular portion 204 with a tubular outlet 204A and a tubular inlet 204B. The flow sleeve 202 also includes a conical portion 206 that is connected to the tubular portion 204 at the tubular inlet 204B, where the conical portion has a conical outlet 206A and a conical inlet 206B. The flow sleeve 202 also includes a plurality of rows of cooling holes 208, where each row 208 extends about a perimeter of the flow sleeve 202. An aft ring 210 is located about the conical inlet 206B and has a receptacle 212 for receiving a piston ring 214. The piston ring 214 incorporates a split-ring design so as to be capable of expanding or contracting in diameter depending on the size of the mating hardware. As such, the piston ring 214 provides an overlap area (not shown) so as to maintain a flow blockage under all conditions. In an embodiment of the present invention, the flow sleeve 202 also includes a plurality of flow straighteners 216, as depicted in FIG. 3. The flow straighteners 216 are designed to channel the flow of cooling air from cooling holes 208 axially thereby reducing swirling characteristics of the cooling air flow.

[0020] A combustion liner 220 is located within the flow sleeve 202 and extends through the axial length of the flow sleeve 202. The combustion liner 220 is generally tubular in shape and, as a result of its location within the combustion system, creates a cooling passage 222 between the combustion liner 220 and flow sleeve 202. It is the air passing through passageway 222 that flow straighteners 216 attempt to straighten so as to achieve a more uniform flow path.

[0021] The present invention also includes a transition duct 230 which is coupled to the outlet of the combustion liner 220 for receiving the hot combustion gases from the combustion liner 220. Due to the geometric requirements of the combustion liner 220 compared to the inlet region of a turbine, the transition duct 230 transitions from a generally circular cross section at the duct inlet to a generally rectangular cross section at the duct outlet.

[0022] As discussed above, the flow sleeve 202 includes a piston ring 214. The piston ring 214 creates a seal between the flow sleeve 202 and the transition duct 230. Because the piston ring is expandable, it has the ability to adjust to various tolerance conditions and differential thermal growth between the flow sleeve 202, which is relatively cold, and the transition duct 230, which is relatively hot, due to the combustion gases contained within. By maintaining a constant contact between the piston ring 214 and the transition duct 230, a seal is formed that prevents compressor air from entering the passageway 222 from the aft end of the flow sleeve 202. Instead all of the compressed air is directed towards the conical and tubular portions of the flow sleeve 202 so that it may enter through the plurality of rows of cooling holes 208.

[0023] As discussed above, the flow sleeve 202 includes a plurality of cooling holes 208 arranged in multiple rows. As shown in FIGS. 3 and 4, at least one of the rows of cooling holes 208 is located in the tubular portion 204 and at least one of the rows is located in the conical portion 206. In an embodiment, the cooling holes 208 are oriented generally perpendicular with respect to the tubular portion 204 and conical portion 206. In an alternate embodiment, the cooling holes 208 can be oriented at a surface angle a relative to both the tubular portion 204. The surface angle a can vary depending on the flow sleeve configuration, but is preferable between approximately 10 and 20 degrees. The cooling holes 208 can also vary in diameter between 0.5 inches and 1.75 inches. The hole sizes of the flow sleeve are specifically sized based on their distance to the combustion liner. This enables a more consistent flow around the flow sleeve and liner annulus, which leads to a more even flow to the head end of the combustor. The hole size is based generally on the ratio of the hole area to the flow area from the hole to the combustion liner.

[0024] In an embodiment of the invention, the flow sleeve 202 has four rows of holes 208 with each row having 24 holes. The aft-most row of holes 208A, closest to the aft ring 210, have a diameter of approximately 0.95 inches, with the adjacent row moving forward, indicated as 208B, have a diameter of approximately 1.5 inches, while the next row 208C have a diameter of approximately 1.575 inches, and the forward-most row of holes 208D having a diameter of approximately 0.65 inches. As a result of the cooling holes 208 being placed in the flow sleeve 202, the cooling holes 208 impart a jet of cool air onto the combustion liner 220 that is located in the flow sleeve 202. The jet impinges air on the liner which cools the liner wall and then the air travels upstream towards an inlet to the combustion liner 220.

[0025] The present invention has been described in relation to particular embodiments, which are intended in all respects to be illustrative rather than restrictive. Alternative embodiments will become apparent to those of ordinary skill in the art to which the present invention pertains without departing from its scope.

[0026] From the foregoing, it will be seen that this invention is one well adapted to attain all the ends and objects set forth above, together with other advantages which are obvious and inherent to the system and method. It will be understood that certain features and sub-combinations are of utility and may be employed without reference to other features and sub-combinations. This is contemplated by and within the scope of the claims.

Claims

1. A flow sleeve for a gas turbine combustor comprising: a tubular portion having a tubular inlet and a tubular outlet; a conical portion having a conical inlet and a conical outlet, the conical inlet secured to the tubular outlet of the tubular portion; a plurality of flow straighteners extending radially inward from the tubular portion and the conical portion; a plurality of rows of cooling holes, each row of cooling holes extending about the flow sleeve; an aft ring having a receptacle, the aft ring secured to the conical outlet; and a piston ring positioned within the receptacle, the piston ring capable of expanding or contracting in diameter.

2. The flow sleeve of claim 1, wherein at least one of the plurality of rows of cooling holes is located in the tubular portion.

3. The flow sleeve of claim 2, wherein at least one of the plurality of rows of cooling holes is located in the conical portion.

4. The flow sleeve of claim 1, wherein the cooling holes impart a jet of cooling air onto a combustion liner located within the flow sleeve.

5. The flow sleeve of claim 1, wherein the piston ring is able to slide axially and radially within the receptacle.

6. The flow sleeve of claim 1, wherein the plurality of rows of holes are generally perpendicular to the tubular portion and the conical portion.

7. The flow sleeve of claim 1, wherein the cooling holes range in diameter from approximately 0.5 inches to approximately 1.75 inches.

8. The flow sleeve of claim 7, wherein the conical portion is oriented at an angle between 10 degrees and 20 degrees relative to the tubular portion.

9. A cooling system for a gas turbine comprising: a flow sleeve comprising: a tubular portion having a tubular inlet and a tubular outlet; a conical portion having a conical inlet and a conical outlet, the conical portion connected to the tubular portion at the tubular outlet; a plurality of rows of cooling holes, each row extending about the perimeter of the flow sleeve; an aft ring located about the conical outlet and having a receptacle; and a piston ring positioned within the receptacle; a combustion liner comprising a generally tubular body and extending through the flow sleeve thereby forming a cooling passage therebetween, the combustion liner having a liner inlet and a liner outlet; and a transition duct coupled to the liner outlet, the transition duct receiving hot combustion gases from the combustion liner, the transition duct having an outer wall transitioning from a generally circular duct inlet to a generally rectangular duct outlet; wherein the piston ring of the flow sleeve creates a seal between the flow sleeve and the transition duct, thereby preventing compressor air from entering a passageway formed between the flow sleeve and the combustion liner, and instead directing the compressed air to enter the passageway through the plurality of rows of cooling holes.

10. The cooling system of claim 9, wherein at least one of the plurality of rows of cooling holes is located in the tubular portion of the flow sleeve.

11. The cooling system of claim 9, wherein at least one of the plurality of rows of cooling holes is located in the conical portion of the flow sleeve.

12. The cooling system of claim 9, wherein the plurality of rows of cooling holes are oriented generally perpendicular to a surface of the tubular portion.

13. The cooling system of claim 9, wherein the plurality of rows of cooling holes are oriented at an angle relative to a surface of the tubular portion.

14. The cooling system of claim 9, wherein the conical portion is oriented at an angle between 10 degrees and 20 degrees relative to the tubular portion.

15. A sealing device for an aft end of a gas turbine combustor comprising: an axially-extending sleeve having a conical portion; an aft ring with a U-shaped receptacle at an outlet end of the conical portion; and a piston ring located in the U-shaped receptacle, the piston ring capable of expanding or contracting in diameter upon placement of an inlet end of a transition duct into the outlet end of the conical portion; wherein a seal is formed between the piston ring and the transition duct, thereby preventing air from entering a combustor through a gap between the sleeve and the transition duct.

16. The sealing device of claim 15, wherein the U-shaped receptacle has a width greater than a width of the piston ring.

17. The sealing device of claim 15, wherein the U-shaped receptacle has a diameter greater than a diameter of the piston ring.

18. The sealing device of claim 15, wherein the piston ring has a generally rectangular cross section.