US20010025476A1

US20010025476A1 - Cooling supply system for stage 3 bucket of a gas turbine

Info

Publication number: US20010025476A1
Application number: US09/773,369
Authority: US
Inventors: Sacheverel Eldrid; James Burns; Gene Palmer; Sal Leone; Gary Drlik; Edward Gibler
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1999-04-15
Filing date: 2001-02-01
Publication date: 2001-10-04
Anticipated expiration: 2019-04-15
Also published as: DE60043131D1; EP1045114B1; JP4778603B2; KR20000071653A; EP1045114A3; US6397604B2; EP1045114A2; JP2000310127A

Abstract

In a land based gas turbine including a compressor, a combustor and turbine section including at least three stages, an improvement comprising an inlet into a third stage nozzle from the compressor for feeding cooling air from the compressor to the third stage nozzle; at least one passageway running substantially radially through each airfoil of the third stage nozzle and an associated diaphragm, into an annular space between the rotor and the diaphragm; and passageways communicating between the annular space and individual buckets of the third stage.

Description

TECHNICAL FIELD

The present invention relates generally to turbines, particularly to land-based gas turbines for power generation, employing compressor air for cooling the buckets of the third turbine stage.

BACKGROUND OF THE INVENTION

Steam cooling of hot gas path components of a gas turbine (for example, the buckets), has been proposed in the past and found viable in land-based power generating plants. While gas turbines are typically air cooled (for example, jet turbines employ compressor discharge air for cooling the hot gas path components), steam cooling is more efficient in that the losses associated with the use of steam as a coolant are not as great as the losses realized by extracting compressor bleed air for cooling. In land based gas turbines and especially those in combined cycle systems, steam cooling is particularly advantageous because the heat energy imparted to the steam as it cools the gas turbine components is recovered as useful work in driving the steam turbine in the combined cycle operation. However, while steam is preferred for cooling the first and second turbine stages, air is required to cool this third stage bucket, and (optionally) to purge the aft portion of the turbine rotor.

BRIEF SUMMARY OF THE INVENTION

In accordance with this invention, air is extracted from the twelfth stage of the compressor and is carried through extraction piping outside the gas turbine, and then supplied through the turbine shell to the stage 3 nozzle. In order to reduce the cycle performance penalty of cooling the third stage bucket, relatively low pressure twelfth stage air is used. The traditional method of bringing air flow from the machine center line practiced by the assignor of this invention is not possible as the forward wheel cavities require high pressure air to drive their purge circuits. Even with steam cooling of the first two bucket stages, air is required to bathe the turbine wheels to control their temperature during transient and start-up operations. In other words, since the forward rotor cavities are filled with high pressure air, a new technique had to be devised for supplying low pressure compressor extraction air for air cooling the third stage. As a result, the bucket cooling air is supplied radially inwardly through the adjacent stator structure, i.e., the third stage nozzle, and then routed to the third stage bucket. In addition, access to relatively low pressure air also provides an optional air source for purge flow in the aft portion of the turbine rotor, with reduced cycle performance penalty. Thus, the invention seeks to introduce low pressure extraction air to the turbine rotor at a low temperature relative to the rotor for use in cooling the third stage bucket. The invention also provides at least an option to make use of the above mentioned air flow to purge the aft section of the turbine rotor, but this is not a preferred arrangement.

In accordance with the invention, a nozzle inducer system comprises a system of tubes carrying the compressor extraction air from the turbine shell through the nozzle airfoils and into the nozzle diaphragm. At the outer end, this tube system penetrates the turbine shell at twenty-two circumferential locations in the exemplary embodiment. Once inside the turbine shell, the piping is split into two conduits, thereby introducing air into forty-four nozzle vanes or airfoils. At the radially inner end of each nozzle vane, the air enters a passage in a diaphragm seal segment which directs the cooling air tangentially into a cavity surrounding the rotor. This passage is configured to accelerate the air in the direction of wheel rotation into this circumferential open area so as to substantially match the tangential velocity of the rotor spacer wheel located radially inwardly of the nozzle. The air is then fed into discrete sets of axial pipes which deliver the air to the shank passages of the stage three buckets. The air then flows radially outwardly through internal passages in the buckets and exits at the bucket tips, into the hot combustion gas path.

The air delivery system in accordance with the invention has several advantages. For example, the use of separate tubing for rotor delivery air minimizes heat transfer to the air from the hot nozzle airfoils. It also allows the use of lower pressure air to pressurize the outer side wall cavities and nozzle cooling circuits which reduces parasitic leakage, improving machine efficiency. In addition, due to the reduction in relative velocity between the rotor spacer wheel and the air, and the drop in air static temperature due to its tangential acceleration, a significantly lower temperature is available for bucket cooling compared to a design where air is simply fed radially into the rotor area.

Accordingly, in its broader aspects, the present invention relates to a land based gas turbine comprising a compressor, a combustor and at least three turbine stages fixed to a rotor, and specifically to an improvement which includes an air cooling circuit for the third turbine stage comprising an inlet into a third stage nozzle from a compressor for feeding cooling air from the compressor to the third stage nozzle; at least one passageway running substantially radially through each airfoil of the third stage nozzle and an associated diaphragm, into an annular space between the rotor and the diaphragm; and passageways communicating between the annular space and individual buckets of the third stage.

The present invention also relates to a method of cooling one stage of a gas turbine comprising a) extracting cooling air from a turbine compressor; b) supplying cooling air to a stationary nozzle adjacent the one stage of the gas turbine; c) establishing a path for the cooling air from the stationary nozzle to a plurality of buckets in the one turbine stage; and d) flowing the cooling air radially outwardly through the plurality of buckets and exhausting the cooling air from radially outer tips of the buckets.

Additional features of the subject invention will become apparent from the detailed description which follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a fragmentary longitudinal cross sectional view of a turbine section of a gas turbine, illustrating the environment of the present invention; [0009]
FIG. 2 is a simplified enlarged detail illustrating the air flow inlet to the third stage nozzle and air flow outlet from a third stage bucket in accordance with this invention; [0010]
FIG. 3 simplified cross section illustrating the cooling air flow path from the third stage nozzle to the shank portion of a third stage bucket in accordance with the invention; [0011]
FIG. 4 is a cross section illustrating the axial cooling passages in the stage [0012] 2-3 spacer;
FIG. 5 is a section taken along the line [0013] 5-5 of FIG. 4;
FIG. 6 is a simplified end view of the third stage nozzle and diaphragm illustrating the cooling air flow path in a pair of adjacent nozzle airfoils; and [0014]
FIG. 7 is an enlarged detail of the tangential cooling air flow passages at the base of the seal segments mounted in the third stage nozzle diaphragm.[0015]

DETAILED DESCRIPTION OF THE INVENTION

With reference to FIG. 1, the [0016] turbine section 10 of a gas turbine is partially illustrated. At the outset, it should be appreciated that the gas turbine of this invention is advantageously utilized in a combined cycle system in which the exhaust gases exiting the gas turbine enter a heat recovery steam generator in which water is converted to steam in the manner of a boiler. Steam thus produced drives one or more steam turbines in which additional work is extracted to drive an additional load, such as a second generator, which, in turn, produces additional electric power.
The [0017] turbine section 10 of the gas turbine is downstream of the turbine combustor 11 and includes a rotor, generally designated R, with four successive stages comprising turbine wheels 12, 14, 16 and 18 mounted to and forming part of the rotor shaft assembly for rotation therewith. Each wheel carries a row of buckets B1, B2, B3 and B4, the blades of which project radially outwardly into the hot combustion gas path of the turbine. The buckets are arranged alternately between fixed nozzles N1, N2, N3 and N4. Alternately, between the turbine wheels from forward to aft are spacers 20, 22 and 24, each located radially inwardly of a respective nozzle. An aft disk 26 forms an integral part of the aft shaft 28 on the aft side of the last stage turbine wheel 18. It will be appreciated that the wheels and spacers are secured to one another by a plurality of circumferentially spaced axially extending bolts 30 (one shown), as in conventional gas turbine construction.
While not per se part of the present invention, a [0018] bore tube assembly 32 forms part of the rotor R and rotates with the rotor about the rotor axis A. The bore tube assembly includes outer and inner tubes 34 and 36 defining annular steam cooling supply passage 38 and spent stream return passage 40. These passages communicate steam to and from the outer rim of the rotor through sets of radial conduits 42, 44 and axially extending conduits (one shown at 46) circumferentially spaced about the rotor rim for supplying cooling steam to the first and second stage buckets B1 and B2. Return or spent cooling steam flows through similar axially and radially extending conduits, respectively, for flow coaxially from the rotor bore via return passage 40. The steam cooling circuit per se, however, forms no part of this invention.
In the exemplary embodiment of this invention, the third stage nozzle N[0019] 3 includes twenty-two part annular segments 48 (see FIG. 6), each having two stationary vanes or airfoils 50, 52. An air manifold 54 outside the turbine shell is designed to supply air from compressor 55 to twenty-two individual pipes (one shown at 56) which penetrate the turbine shell and which are connected to the twenty-two respective segments. For simplicity, the compressor 55 and manifold 54 are shown schematically in FIG. 2. Inside the shell, the pipe 56 feeds two supply pipes 58, 58 a, etc. for each of the forty-four vanes or airfoils (see FIG. 6). Pipes 58, 58 a are connected by flexible connector couplings shown at 59. For convenience, only one flow circuit need be described in detail.
With specific reference to FIGS. 2, 3 and [0020] 6, a passage or conduit 60 is shown extending radially within the vane or airfoil 50, with a generally radially extending, flexible coupling or connector 62 (incorporating a carbon bushing, not shown) carrying the air within the diaphragm 64. At its radially inner end, the connector is operatively connected to a diaphragm insert 66 by means of a spoolie device 68. The latter, having generally spherically shaped opposite ends, in combination with the flexible coupling 62, accommodate any relative movement between the insert 66 and the diaphragm 64. As is well known, the diaphragm inserts 66 comprise a plurality of part annular segments extending circumferentially about the rotor, with labyrinth seals 70 engaged with cooperating seals 72 on the rotor spacer wheel 22 to prevent leakage of air along the rotor.
Within the [0021] insert 66, the air passage changes direction via elbow passage 70 and substantially straight passage 72 to direct the air tangentially (at an angle α of about 22-23°) into an annular rotor cavity 74, as best seen in FIGS. 6 and 7. Passage 70 tapers in the flow direction through an elbow portion to a smaller diameter at passage 72, thereby causing acceleration of the cooling air as it is fed into the annular cavity 74. As a result of this inducer arrangement, the air as supplied to cavity 74 is relatively “still” vis-a-vis the rotor. In other words, the air is fed tangentially at a speed substantially the same as the rotational speed of the rotor. This results in cooler air being available for the third stage buckets, due to the reduction in relative velocity between the rotor spacer wheel and the air, and the drop in air static temperature due to its tangential acceleration.
From the [0022] annular rotor cavity 74, the cooling air moves axially through multiple sets of three passages 76 each (see FIGS. 3, 4 and especially 5), with access to the passages permitted by forming the spacer wheel 22 with scalloped areas 78 about its periphery as best seen in FIG. 5.
Note in this regard that the individual sets of [0023] passages 76 are located circumferentially between axial steam supply and return passages 46 and radially outwardly of bores 80 (one shown) for bolts 30.
The air then moves radially outwardly at the interface of [0024] spacer 22 and the third stage wheel 16, to an axial supply passage 82 between the wheel rim and the bucket shank. From here, the air travels radially outwardly in one or more radial passages 86, and then vents into the hot gas path at the bucket tips (see the flow arrows in FIGS. 1 and 2). In order to prevent leakage of cooling air between the spacer 22 and wheel 16, an annular wire seal 86 is located within a groove formed in the radially outermost edge of spacer 22. Since the wheel 16 and spacer 22 are rotating together with the rotor, there is no relative frictional movement between the seal 86 and the wheel 16.
While the invention as described relates to air cooling in land based turbines, it can be applied to aircraft turbines as well. [0025]
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. [0026]

Claims

What is claimed is:

1. In a land based gas turbine comprising a compressor, a combustor and turbine section including at least three stages, a cooling circuit comprising:

an inlet into a third stage nozzle from the compressor for feeding cooling air from the compressor to the third stage nozzle; at least one passageway running substantially radially through each airfoil of said third stage nozzle and an associated diaphragm, into an annular space between the rotor and the diaphragm; and passageways communicating between said annular space and individual buckets of said third stage.

2. The gas turbine of

claim 1

wherein said at least one passageway includes a portion in said diaphragm configured to feed said cooling air into said annular space substantially tangent to the rotor.

3. The gas turbine of

claim 2

wherein said portion of said at least one passageway contracts in the flow direction to accelerate the cooling air as it enters said annular space.

4. The gas turbine of

claim 2

wherein said portion of at least one passageway in said diaphragm is provided in a part annular labyrinth seal segment secured to said diaphragm and cooperating with a corresponding seal on a rotor spacer wheel located radially inwardly of said third stage nozzle.

5. The gas turbine of

claim 4

wherein means are provided to accommodate relative motion or mismatch between said portion of said at least one passageway in said seal segment and said diaphragm.

6. The gas turbine of

claim 1

wherein said inlet to said third stage nozzle includes a manifold external to a casing of said gas turbine.

7. The gas turbine of

claim 4

wherein said passageways communicating between said annular space and said individual buckets include plural sets of axial passages through said spacer wheel.

8. The gas turbine of

claim 1

wherein said third stage nozzle includes a plurality of part annular segments, each segment having two nozzle airfoils, and wherein said inlet includes a pipe feeding cooling air to each segment, said pipe supplying cooling air to each of said two nozzle airfoils.

9. The gas turbine of

claim 1

wherein stages 1 and 2 are primarily steam cooled.

10. The gas turbine of

claim 4

including an annular seal between said spacer wheel and said third stage buckets to prevent leakage of cooling air passing into said third stage buckets.

11. A method of cooling one stage of a gas turbine comprising:

a) extracting cooling air from a turbine compressor;

b) supplying cooling air to a stationary nozzle adjacent said one stage of the gas turbine;

c) establishing a path for said cooling air from said stationary nozzle to a plurality of buckets in said one turbine stage; and

d) flowing said cooling air radially outwardly through said plurality of buckets and exhausting said cooling air from radially outer tips of said buckets.

12. The method of

claim 11

wherein during step c), said cooling air is fed tangentially into an annular space surrounding a rotor of said gas turbine.

13. The method of

claim 12

wherein said cooling air is accelerated into said annular space.

14. The method of

claim 11

wherein said one turbine stage is a third stage.

15. The method of

claim 11

wherein said cooling air is supplied to said stationary nozzle via a path outside said gas turbine.

16. In a land based gas turbine comprising a compressor, a combustor and turbine section including at least three stages, a cooling circuit comprising:

means for supplying cooling air from a gas turbine compressor to a stationary nozzle; and

means for establishing a cooling air flow path from said nozzle to individual buckets of a turbine stage downstream and adjacent said stationary nozzle.