Combustor liner cooling thimbles and related method

Description

BACKGROUND OF THE INVENTION

This invention relates to combustors in turbo machinery and specifically, to the cooling of combustor liners in gas turbine combustors.

Conventional gas turbine combustion systems employ multiple combustor assemblies to achieve reliable and efficient turbine operation. Each combustor assembly includes a cylindrical liner, a fuel injection system, and a transition piece that guides the flow of the hot gas from the combustor to the inlet of the turbine. Generally, a portion of the compressor discharge air is used to cool the combustion liner and is then introduced into the combustor reaction zone to be mixed with the fuel and burned.

In systems incorporating impingement cooled transition pieces, a hollow sleeve surrounds the transition piece, and the sleeve wall is perforated so that compressor discharge air will flow through the cooling apertures in the sleeve wall and impinge upon (and thus cool) the transition piece. This cooling air then flows along an annulus between the sleeve surrounding the transition piece, and the transition piece itself. This so-called “cross flow” eventually flows into another annulus between the combustion liner and a surrounding flow sleeve. The flow sleeve is also formed with several rows of cooling holes around its circumference, the first row located adjacent a mounting flange where the flow sleeve joins to the outer sleeve of the transition piece. The cross flow is perpendicular to impingement cooling air flowing through the holes in the flow sleeve toward the combustor liner surface.

The presence of this crossflow has a direct influence on the cooling effectiveness in the zone near where the first row of jets in the flow sleeve would have been expected to impingement cool the combustor liner. Specifically, the crossflow impacts the first row of flow sleeve jets, bending them over and degrading their ability to impinge upon the liner. In one prior design of the flow sleeve impingement jets there are three rows of 24 jets spaced evenly around the circumference of the flow sleeve. This jet pattern in the presence of the strong crossflow from the transition piece impingement sleeve produces very low heat transfer rates on the liner surface near the first row of jets. This low heat transfer rate can lead to high liner surface temperatures and ultimately loss of strength. Several potential failure modes due to the high temperature of the liner include, but are not limited to, cracking of the aft sleeve weld line, bulging and triangulation. These mechanisms shorten the life of the liner, requiring replacement of the part prematurely.

SUMMARY OF THE INVENTION

This invention enhances the cooling of the liner in Dry Low NOx type gas turbine combustors where jet impingement is used to cool the aft portion of the combustor liner. Even though there is a strong crossflow resulting from the transition piece cooling flow, the negative impact of the crossflow is minimized by the use of collars or cooling conduits, also referred to as a thimbles, that are inserted into the cooling holes in the combustor liner flow sleeve, through which the cooling jets pass. These thimbles provide a physical blockage to the cross flow which forces the crossflow into the desired flow path while simultaneously ensuring that the cooling jets effectively impinge on the combustor liner surface to be cooled.

The thimbles or collars are preferably mounted in each hole of at least the first row of holes at the aft end of the flow sleeve, adjacent a mounting flange where the combustor liner and transition piece are joined. This arrangement decreases the gap between the jet orifice and impingement surface; blocks the cross flow that deflects the jets and forces it into the desired flowpath for the subsequent jet rows; allows the diameter of the jet to be smaller and thereby reduce cooling air; and provides consistent and accurate control over the location of jet impingement. It also stabilizes unwanted axial oscillation of the first row of jets, and prevents the formation of a thick boundary layer (and resulting reduced heat transfer) upstream of the first row of jets.

Accordingly, in its broader aspects, this invention relates to a combustor for a turbine that includes: a combustor liner; a first flow sleeve surrounding the liner, the flow sleeve having a plurality of rows of cooling holes formed about a circumference of the flow sleeve; a transition piece connected to the combustor liner and the flow sleeve and adapted to carry hot combustion gases to a stage of the turbine; the transition piece surrounded by a second flow sleeve; wherein a first of the plurality of rows of cooling in the first flow sleeve is located adjacent the transition piece; and further wherein one or more cooling conduits are mounted in the cooling holes of at least the first of the plurality of rows of cooling holes.

In another aspect, the invention relates to combustion liner flow sleeve adapted for mounting in surrounding relationship to a combustion liner, the flow sleeve comprising a tubular body formed with plural rows of cooling holes, one of the plural rows located adjacent one end of the flow sleeve; and a cooling conduit mounted in each hole of at least the first row of holes, the cooling conduits projecting radially into the flow sleeve.

In still another aspect, the invention relates to a method of cooling a combustor liner of a gas turbine combustor, the combustor liner having a substantially circular cross-section, and a flow sleeve surrounding the liner in substantially concentric relationship therewith creating an annulus therebetween; the method comprising a) providing a plurality of axially spaced rows of cooling holes in the flow sleeve, each row extending circumferentially of the flow sleeve, one of the rows adjacent a forward or downstream end of the flow sleeve; b) locating a plurality of cooling conduits in the cooling holes of at least the first of the rows, the cooling conduits extending radially towards, but not engaging, the liner; and c) supplying cooling air to the annulus such that the cooling conduits direct the cooling air against the liner.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1

is a simplified side cross section of a conventional combustor transition piece aft of the combustor liner;

FIG. 2

is a partial but more detailed perspective of a conventional combustor liner and flow sleeve joined to the transition piece;

FIG. 3

is a flow diagram illustrating impingement cooling of the combustor liner in a prior arrangement;

FIG. 4

is a side elevation of a cooling thimble in accordance with the invention;

FIG. 5

is a side section of the thimble shown in

FIG. 4

; and

FIG. 6

is a flow diagram illustrating impingement cooling of the combustor liner in accordance with this invention.

DETAILED DESCRIPTION OF THE INVENTION

With reference to

FIGS. 1 and 2

, a typical gas turbine includes a transition piece

10

by which the hot combustion gases from an upstream combustor as represented by the combustor liner

12

are passed to the first stage of a turbine represented at

14

. Flow from the gas turbine compressor exits an axial diffuser

16

and enters into a compressor discharge case

18

. About 50% of the compressor discharge air passes through apertures

20

formed along and about a transition piece impingement sleeve

22

for flow in an annular region or annulus

24

(or, second flow annulus) between the transition piece

10

and the radially outer transition piece impingement sleeve

22

. The remaining approximately 50% of the compressor discharge flow passes into flow sleeve holes

34

and mixes with the air from the transition piece from annulus

30

and eventually mixes with the gas turbine fuel in the combustor.

FIG. 2

illustrates the connection between the transition piece

10

and the combustor flow sleeve

28

as it would appear at the far left hand side of FIG.

1

. Specifically, the impingement sleeve

22

(or, second flow sleeve) of the transition piece

10

is received in a telescoping relationship in a mounting flange

26

on the aft end of the combustor flow sleeve

28

(or, first flow sleeve), and the transition piece

10

also receives the combustor liner

12

in a telescoping relationship. The combustor flow sleeve

28

surrounds the combustor liner

12

creating a flow annulus

30

(or, first flow annulus) therebetween. It can be seen from the flow arrow

32

in

FIG. 2

, that crossflow cooling air traveling in the annulus

24

continues to flow into the annulus

30

in a direction perpendicular to impingement cooling air flowing through the cooling holes

34

(see flow arrow

35

) formed about the circumference of the flow sleeve

28

(while three rows are shown in

FIG. 2

, the flow sleeve may have any number of rows of such holes).

The impingement cooling flow in the first row of holes

34

in the flow sleeve (the row of holes closest to the mounting flange

26

) is particularly subject to disruption by the crossflow from the annulus

24

. The negative impact can be seen in FIG.

3

. The test impingement sleeve/combustor liner configuration shown in

FIG. 3

is slightly different than the arrangement in

FIG. 2

, but similar reference numerals are used for ease of understanding. The cross flow impacts on the first row cooling jets exiting the holes

34

, bending them over and degrading their ability to impinge upon the liner

12

. Depending on the relative strengths of the cross flow and jets, the jet flow may not even reach the surface of the combustor liner

12

. Because the impingement jets are high velocity, there is a characteristic zone of low static pressure behind the jets and near the flow sleeve entrance hole. The cross flow accelerates toward the low pressure zone, leading to a velocity gradient across the flow sleeve/liner annulus

30

. The resulting low velocity and thickened boundary layer near the liner surface has very poor heat transfer effectiveness.

To neutralize the negative impact of the crossflow on the cooling jets, cooling thimbles

36

as shown in

FIG. 4

are employed. These thimbles may comprise tubes

38

of circular cross section, with a flat ring or flange

40

welded to its top. The dimensions of the thimbles

36

are set by the liner/flow sleeve gap, the tolerance of this dimension, the jet and cross flow momentum, the geometric constraints of the thimble, and the specific cooling requirements. A single thimble

36

is inserted into each of the first row flow sleeve holes

34

and welded at the ring or flange

40

to the outside of the flow sleeve

28

. There are

20

thimbles mounted in a like number of holes

34

around the circumference of the flow sleeve

28

in at least the first row adjacent mounting flange

26

, but they could be added to the second and/or third row as well. The critical location to be cooled in one exemplary embodiment, has been determined to be about 30 inches from the opposite end of the flow sleeve (not shown).

The cross section of the tube

38

is shown to be circular but other cross-sectional shapes could be used, e.g., square, triangular, airfoil-shaped, semi-circular and the like. An open channel-shape could also be used.

FIG. 6

shows a cross section revealing the flow field in the liner/flow sleeve annulus with the thimbles

36

in place. This figure may be contrasted with

FIG. 3

to identify the benefits of enhancing first row jets by use of flow sleeve thimbles. It is particularly evident that impingement cooling of the liner

12

is enhanced at the first row of holes

36

, and that the crossflow has been blocked at least to the extent of allowing cooling jets from the impingement cooling holes

34

to reach the surface of the liner

12

. In fact, it has been established that the thimbles add 150 BTU/HR*ft

2

*F to the liner heat transfer coefficient at the first row jet centerline, 29.2 inches from the opposite end of the flow sleeve.

The device and process in accordance with this invention can be used in any impingement cooling application where there is a crossflow present at the first row of jets.

While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.

Claims

1. In a combustor for a turbine:a combustor liner; a first flow sleeve surrounding said combustor liner with a first flow annulus therebetween, said first flow sleeve having a plurality of rows of cooling holes formed about a circumference of said first flow sleeve for directing cooling air into said first flow annulus; a transition piece connected to said combustor liner and said first flow sleeve and adapted to carry hot combustion gases to a stage of the turbine; said transition piece surrounded by a second flow sleeve formed with an array of cooling apertures for directing cooling air into a second flow annulus between the second flow sleeve and the transition piece, said first flow annulus connecting to said second flow annulus; wherein a first of said plurality of rows of cooling holes in said first flow sleeve is located adjacent said transition piece; and further wherein one or more cooling conduits are mounted in the cooling holes of at least said first of said plurality of rows of cooling holes in said first flow sleeve for directing cooling air into said first flow annulus to impingement cool the combustor liner and mix with cooling air from said second flow annulus.
2. The combustor of claim 1 wherein said cooling conduits each include a tube with a radial flange at one end thereof.
3. The combustor of claim 2 wherein each said cooling conduit is welded to said flow sleeve, with an opposite end of said tube spaced from said liner by a predetermined space.
4. The combustor of claim 2 wherein one of said plurality of cooling conduits are circular in cross-section.
5. A method of cooling a combustor liner of a gas turbine combustor, said combustor liner having a substantially circular cross-section, and a first flow sleeve surrounding said liner in substantially concentric relationship therewith creating a first flow annulus therebetween for feeding air to the gas turbine combustor, and wherein a transition piece is connected to said combustor liner, with the transition piece surrounded by a second flow sleeve, thereby creating a second flow first annulus in communication with said first flow first annulus; the method comprising:a) providing a plurality of axially spaced rows of cooling holes in said flow sleeve, each row extending circumferentially of said flow sleeve, a first of said rows adjacent said transition piece and said second flow sleeve; b) locating a plurality of cooling conduits in the cooling holes of at least said first of said rows, said cooling conduits extending radially inwardly towards, but not engaging, said combustor liner; c) supplying cooling air to said cooling holes such that said cooling conduits direct said cooling air against said combustor liner; and d) thereafter, directing said cooling air axially within said first flow annulus.
6. The method of claim 5 wherein said cooling conduits each include a tube with a radial flange at one end thereof.
7. The method of claim 6 wherein said tube has a circular cross section.
8. A gas turbine combustor assembly comprising a combustor liner and a transition piece connected to a downstream end of the combustor liner, the transition piece adapted to carry hot combustion gases in a first flow direction to a first stage of a gas turbine; a first flow sleeve surrounding the combustor liner with a first flow annulus radially therebetween; a second flow sleeve surrounding the transition piece with a second flow annulus radially therebetween, with said first annulus communicating directly with said second flow annulus; a plurality of rows of impingement cooling holes in said first flow sleeve, at least one of said plurality of rows located adjacent said transition piece at said downstream end of the combustor liner with a plurality of cooling conduits located in the impingement cooling holes of at least a first of said plurality of rows in said first flow sleeve and projecting toward said combustor liner; wherein, in use, cooling air flows through said cooling conduits to impinge on said combustor liner and mix with cooling air flowing through said first and second flow annuli in a second flow direction counter to said first flow direction.

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4339925	Eggmann et al.	Jul 1982	A
4719748	Davis, Jr. et al.	Jan 1988	A
4872312	Iizuka et al.	Oct 1989	A
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“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, vol. II, “Heat Pipe Turbine Vane Cooling”, Langston et al., pp. 566-572, Oct., 1995.
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“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “ATS and the Industries of the Future”, Denise Swink, p. 1, Nov., 1996.
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“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Power Needs in the Chemical Industry”, Keith Davidson, pp. 17-26, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Advanced Turbine Systems Program Overview”, David Esbeck, p. 27-34, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Westinghouse's Advanced Turbine Systems Program”, Gerard McQuiggan, p. 35-48, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Progam Review Meeting”, “Overview of GE's H Gas Turbine Combined Cycle”, Cook et al., p. 49-72, Nov., 1996.
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“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “The AGTSR Industry-University Consortium”, Lawrence P. Golan, p. 95-110, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “NOX and CO Emissions Models for Gas-Fired Lean-Premixed Combustion Turbines”, A. Mellor, p. 111-122, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Methodologies for Active Mixing and Combustion Control”, Uri Vandsburger, p. 123-156, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Combustion Modeling in Advanced Gas Turbine Systems”, Paul O. Hedman, p. 157-180, Nov., 19967.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Manifold Methods for Methane Combustion”, Stephen B. Pope, p. 181-188, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “The Role of Reactant Unmixedness, Strain Rate, and Length Scale on Premixed Combustor Performance”, Scott Samuelsen, p. 189-210, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Effect of Swirl and Momentum Distribution on Temperature Distribution in Premixed Flames”, Ashwani K. Gupta, p. 211-232, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Combustion Instability Studies Application to Land-Based Gas Turbine Combustors”, Robert J. Santoro, p. 233-252.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, Active Control of Combustion Instabilities in Low NOX Turbines, Ben T. Zinn, p. 253-264, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Life Prediction of Advanced Materials for Gas Turbine Application,” Sam Y. Zamrik, p. 265-274, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Combustion Chemical Vapor Deposited Coatings for Thermal Barrier Coating Systems”, W. Brent Carter, p. 275-290, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Compatibility of Gas Turbine Materials with Steam Cooling”, Vimal Desai, p. 291-314, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Bond Strength and Stress Measurements in Thermal Barrier Coatings”, Maurice Gell, p. 315-334, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Advanced Multistage Turbine Blade Aerodynamics, Performance, Cooling and Heat Transfer”, Sanford Fleeter, p. 335-356, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Flow Characteristics of an Intercooler System for Power Generating Gas Turbines”, Ajay K. Agrawal, p. 357-370, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Improved Modeling Techniques for Turbomachinery Flow Fields”, B. Lakshiminarayana, p. 371-392, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Development of an Advanced 3d & Viscous Aerodynamic Design Method for Turbomachine Components in Utility and Industrial Gas Turbine Applications”, Thong Q. Dang, p. 393-406, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Advanced Turbine Cooling, Heat Transfer, and Aerodynamic Studies”, Je-Chin Han, pp. 407-426, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Heat Transfer in a Two-Pass Internally Ribbed Turbine Blade Coolant Channel with Vortex Generators”, S. Acharya, p. 427-446.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Experimental and Computational Studies of Film Cooling with Compound Angle Injection”, R. Goldstein, p. 447-460, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Study of Endwall Film Cooling with a Gap Leakage Using a Thermographic Phosphor Fluorescence Imaging System”, Mingking K. Chyu, p. 461-470, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Steam as a Turbine Blade Coolant: External Side Heat Transfer”, Abraham Engeda, p. 471-482, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Flow and Heat Transfer in Gas Turbine Disk Cavities Subject to Nonuniform External Pressure Field”, Ramendra Roy, p. 483-498, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Closed-Loop Mist/Steam Cooling for Advanced Turbine Systems”, Ting Wang, p. 499-512, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Heat Pipe Turbine Vane Cooling”, Langston et al., p. 513-534, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “EPRI's Combustion Turbine Program: Status and Future Directions”, Arthur Cohn, p. 535,-552 Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “ATS Materials Support”, Michael Karnitz, p. 553-576, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Land Based Turbine Casting Initiative”, Boyd A. Mueller, p. 577-592, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Turbine Airfoil Manufacturing Technology”, Charles S. Kortovich, p. 593-622, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Hot Corrosion Testing of TBS's”, Norman Bornstein, p. 623-631, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Ceramic Stationary Gas Turbine”, Mark van Roode, p. 633-658, Nov., 1996.
“Proceedings of the Advanced Turbine Systems Annual Program Review Meeting”, “Western European Status of Ceramics for Gas Turbines”, Tibor Bornemisza, p. 659-670, Nov., 1996.
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Combustor liner cooling thimbles and related method

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