WO2018203924A1

WO2018203924A1 - Gas turbine engine with a rim seal

Info

Publication number: WO2018203924A1
Application number: PCT/US2017/055550
Authority: WO
Inventors: Todd A. Ebert
Original assignee: Florida Turbine Technologies, Inc.
Priority date: 2017-05-03
Filing date: 2017-10-06
Publication date: 2018-11-08

Abstract

A rim seal for a turbine in a gas turbine engine, the rim seal formed between a rotor and a stator of the turbine and includes a first seal with an expansion cavity and a second seal formed radially inward of the first seal. A buffer cavity is formed between the first seal and the second seal, and a trench is formed between the first seal and a hot gas path of the turbine.

Description

GAS TURBINE ENGINE WITH A RIM SEAL

GOVERNMENT LICENSE RIGHTS

This invention was made with Government support under contract number DE-FC26-05NT42644 awarded by Department of Energy. The Government has certain rights in the invention.

TECHNICAL FIELD

The present invention relates generally to a gas turbine engine, and more specifically to a rim seal between a rotor and a stator in a gas turbine of an industrial gas turbine engine.

BACKGROUND

In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a gas turbine to produce mechanical work. The turbine includes one or more rows of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine - and therefore the engine - can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.

The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the hot gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.

A seal must be formed between the rotor and the stator of the turbine in a gas turbine engine in order to prevent ingestion of the hot gas stream into the rim cavity which can affect the rotor disk. Labyrinth or angle wing seals are used in the prior art to produce a seal between the rotating part and the static part. However, the seal gaps change during engine operation because of the thermal growth on the parts that forms the seal.

Rim seals are axial extensions of a turbine rotor blade, i.e. a bucket, which form a seal by overlapping with vane (nozzle) seal lands forming part of the fixed component of a gas turbine. The rim seal inhibits ingestion of hot gas from the flow path into gas turbine wheel spaces. Typically, rim seals are cast integrally as part of the blade or bucket, or are multiple overlays having multiple angel wings.

Conventional airfoil platform seals have such a shape that the vane cannot be removed from the turbine without also removing the adjacent rotor blade because of the overlapping of adjacent platforms, i.e. the platform extending from the vane overlaps with the platform extending from the blade. Multiple overlap rim seals are assembled axially, and therefore the vanes cannot be removed radially from the casing due to interference with platforms on the blades that form the rim seal. U.S. Patent No. 5,236,302 issued to Weisgerber et al. on August 17, 1993 shows a turbine disc interstage seal system in which an air seal is formed between adjacent platforms of the blade and the vane, where a finger of the vane platform extends in-between a space formed between two fingers extending from the blade platform. The vane in the Weisgerber invention cannot be removed from the turbine without removing the blade, since the fingers on the platforms interfere with each other.

Gas turbine engines also produce circumferential static pressure variations downstream from the airfoils. In a typical gas turbine, the gas stream flows past the airfoils both rotating and stationary, and the static pressure exiting the airfoil passage varies between two extreme pressures. This variation in static pressure acts across the rim seal at the platforms, and will cause undesirable hot gas ingestion into the wheel space without the presence of a rim seal. Multiple overlaps create a desirable buffer cavity or volume to dissipate this circumferential pressure variation.

SUMMARY

A rim seal for a gas turbine in a gas turbine engine, where the rim seal includes a first seal with an expansion cavity formed between the rotor and the stator, and a second seal formed between the rotor and the stator radially inward of the first seal. A buffer cavity is formed between the first seal and the second seal. A trench is formed between the first seal and a hot gas path of the turbine.

The first seal includes two radial projections that form labyrinth seals between the expansion cavity, which is formed mostly within the stator. The second seal includes a single angle wing labyrinth seal extending from the rotor. Pressure variations from the hot gas stream will be buffered with passage through the expansion cavity and into the buffer cavity such that flow passes from the rim cavity into the buffer cavity and therefore prevents any hot gas flow from leaking into the rim cavity to produce a very effective rim seal for the gas turbine. In one embodiment, a gas turbine engine includes: a rotor with a rotor blade; a stator with a stator vane; a first seal point formed between the rotor and the stator; and a second seal point formed between the rotor and the stator, the first seal point including an expansion cavity formed in both the rotor and the stator; and the expansion cavity including sharp corners on an upstream side of the expansion cavity.

In one aspect of the embodiment, the expansion cavity includes an inner expansion cavity on the rotor and an outer expansion cavity on the stator. In one aspect of the embodiment, the outer expansion cavity is formed on a radial inward extension of the rotor; and the inner expansion cavity is formed on an axial extension of the stator. In one aspect of the embodiment, the second seal point is located radially outward from the first seal point; the second seal point is formed from a radial outward extension from the stator; and a buffer cavity is formed between the radial inward extension and the radial outward extension. In one aspect of the embodiment, a trench cavity formed between the stator and the radial inward extension in direct communication with a hot gas path of the gas turbine engine.

In one aspect of the embodiment, the inner expansion cavity) is shorter than the outer expansion cavity. In one aspect of the embodiment, the first seal point is located radially outward from the second seal point. In one aspect of the embodiment, the second seal point is an angel wing seal. In one aspect of the embodiment, a buffer cavity formed between the first seal point and the second seal point. In one aspect of the embodiment, an axial extension from the stator separates the buffer cavity into an inner buffer cavity and an outer buffer cavity. In one aspect of the embodiment, both the inner expansion cavity and the outer expansion cavity are formed by sharp corners on an inlet side and an outlet side of the expansion cavity. BRIEF DESCRIPTION OF THE DRAWINGS

A more complete understanding of the present invention, and the attendant advantages and features thereof, will be more readily understood by reference to the following detailed description when considered in conjunction with the accompanying drawings wherein:

FIG. 1 shows a rim seal of the present invention according to a first embodiment as a steady state condition after HCO;

FIG. 2 shows the rim seal of the present invention according to a first embodiment at a steady state condition prior to HCO;

FIG. 3 shows the rim seal of the present invention according to a second embodiment at a steady state condition prior to HCO; and

FIG. 4 shows a rim seal of the present invention according to a second embodiment as a steady state condition after HCO.

DETAILED DESCRIPTION

The present invention is a rim seal in a gas turbine of an industrial gas turbine engine, especially for an industrial gas turbine engine that includes a hydraulic clearance optimization (HCO) where the rotor is shifted in an axial direction to change the clearance between blade tips or seals in the turbine section. FIG. 1 shows a rim seal for a turbine between the rotor which includes a rotor blade 11 with a platform 12 and an attachment 13, such as a fir tree arrangement, and a stator which includes a vane 14 extending from an inner endwall 31. A cover plate 25 is secured to the rotor. The cover plate 25 includes an axial extension with radially extending projections 22 and 23 (which may be referred to as the first or forward radially extending projection 22 and the second or aft radially extending projection 23) that form the rim seal with surfaces formed on the inner endwall 31.

The inner endwall 31 includes an expansion cavity 32 (which may also be referred to as an outer expansion cavity 32) formed by a forward seal forming surface or first radially extending projection 34 and an aft seal forming surface or second radially extending projection 33. The cover plate 25 includes an expansion cavity 24 (which may also be referred to as an inner expansion cavity 24) formed by the forward or first radially extending projection 22 and the second or aft radially extending projection 23. The forward 34 and aft 33 seal forming surfaces/radially extending projections and the forward 22 and aft 23 radially extending projections together form the rim seal for the turbine. The two expansion cavities 32 and 24 are opposed to each other and together form a first seal point 39 that includes an expansion cavity 40 (that is, expansion cavity 40 is formed from the two expansion cavities 32 and 24). In one embodiment, the expansion cavity 32 of the stator is larger than the expansion cavity 24 of the rotor and, therefore, the expansion cavity 40 formed of the two expansion cavities 24, 32 is formed mostly in the stator.

Another or third radially extending projection 35 extends from the cover plate 25 and forms a second seal point 38. Buffer cavities 36 (for example, an inner buffer cavity and an outer buffer cavity) are formed between the inner endwall 31 of the stator and the cover plate 25 of the rotor between the first seal point 39 and the second seal point 38. The second seal point 38 is located radially inward from the first seal point 39. In one embodiment, the second seal point 38 is an angel wing seal with an axial extension 57, the third radially extending projection 35 being at an end of the axial extension. The angel wing seal forms at least a portion of the buffer cavity 36, in which the rotor produces a vortex flow in the buffer cavity 36 due to rotation of the rotor such that any leakage flow across the angle wing seal is minimized. In one embodiment, the buffer cavity 36 includes an outer buffer cavity and an inner buffer cavity located axially inward from the outer buffer cavity, and the inner and outer buffer cavities are separated from each other by an axial extension 58 from the stator that forms a gap with the rotor. A trench cavity 37 is formed between the inner endwall 31 and the cover plate 25 on the rotor outward from the first seal point 39. Thus, the trench cavity 37 is located between the first seal point 39 and the hot gas path of the turbine.

The expansion cavity 40 formed by the two expansion cavities 32 and 24 is formed with sharp corners at the inlet side and the outlet side of the expansion cavity 40 so that the flow into the cavity will quickly expand due to the flow detaching from the surfaces. The expansion cavity 40 is designed to both expand and compress flow as it enters and exits through this area at 90 degree angle. This separation causes mixing to further dampen the inlet pressure signal. The buffer cavities 36 then further dampen the pressure fluctuations to prevent any hot gas ingestion into the rim cavity.

FIG. 2 shows the rim seal of the present invention in a steady state condition after the HCO. The difference between FIG. 2 and FIG. 1 is that the rotor shaft has been shifted in the axial direction. The contraction and expansion and contraction process greatly reduces unsteady circumferential pressure variation, which is the major mechanism causing hot gas ingestion into the rim cavity.

A second embodiment of the rim seal is shown in FIGS. 3 and 4. FIG. 3 shows a rim seal for a turbine between the rotor which includes a rotor blade 41 with a platform 42 with an attachment such as a fir tree arrangement extending from the platform 42, and a stator with includes a stator vane 43 extending from an inner endwall 44. FIG. 4 shows the rotor shifted axially prior to HCO. The rotor platform 42 includes a radial inward extension 45 that extends inward with two radially extending projections 48 and 49 (which may be referred to herein as a first or forward radially extending projection 48 and a second or aft radially extending projection 49) extending from the ends of the radial inward extension 45 that forms a rotor expansion cavity 50. The stator includes an axial extension 47 and a radial upward extension 46 with a stator expansion cavity 51 formed on the axial extension 47. The stator expansion cavity 51 is bordered on either side by a first or forward seal forming surface 59 and a second or aft seal forming surface 60.

FIG. 4 shows the rim seal in a steady state condition after HCO when the rotor has been shifted axially. In the FIG. 4 condition, the rim seal is fully engaged. The radial upward extension 46 of the stator forms a HCO pad 54 with the underside of the rotor platform 42. The radial inward extension 45 of the rotor forms a seal point 55 with the top surface of the axial extension 47. The first 48 and second 49 radial projections of the radial inward extension 45 of the rotor form close gaps with the first 59 and second 60 seal forming surfaces of the axial extension 47 of the stator, respectively. The rotor expansion cavity 50 and the stator expansion cavity 51 are opposed to each other and aligned to form an expansion cavity 56 for the seal point 55. The side of the inner endwall 44 and the side of the radial inward extension 45 of the rotor form a trench cavity 52. Thus, the trench cavity 52 is located between the seal point 55 and the hot gas path of the turbine. A buffer cavity 53 is formed between the radial inward extension 45 of the rotor and the radial upward extension 36 of the stator, between the seal point 55 and the HCO pad 54. The rotor produces a vortex flow in the buffer cavity 53 due to rotation of the rotor, thereby minimizing any leakage flow across HCO pad 54. The seal point 55 is located radially inward from the HCO pad 54. Put another way, the HCO pad 54 is located radially outward from the seal point 55.

The expansion cavity 56 formed by the rotor expansion cavity 50 and the stator expansion cavity 51 includes sharp edges that are designed to both expand and compress flow as it enters and exits the cavity through this area at a 90 degree angle. This quick expansion causes the flow to separate from the surface which results in mixing within the cavity to further dampen the inlet pressure signal.

Two seal points are formed in the FIG. 4 design: seal point 55 is one seal and the HCO pad 54 forms a second seal to significantly reduce hot gas ingestion into the rim cavity from the hot gas path through the turbine. Thus, the seal point 55 may be referred to as a first seal point 55 and the HCO pad 54 may be referred to as a second seal point 54. The expansion cavity 56 formed by the rotor expansion cavity 50 and the stator expansion cavity 51 in the first seal point 55 further improves the seal capability.

The rim seal design of the second embodiment (for example, as shown in FIGS. 3 and 4) of the present invention not only provides for a very effective seal for the rim cavity, but also allows for radial assembly and disassembly of the rotor blades without the need to remove any of the adjacent stator vanes. In the prior art, to remove a rotor blade, one or more adjacent stator vanes would have to be removed first in order to axially slide out the rotor blades. With the design of FIGS. 3 and 4, each rotor blade can be removed or installed in a radial direction without removal of any adjacent stator vanes. Industrial gas turbine engines are designed for assembly of the vanes in a radial direction due to the engine being of a split case design with an upper case and a lower case.

FIG. 4 shows the rim seal of the present invention in a steady state condition after the HCO. The difference between FIG. 4 and FIG. 3 is that the rotor shaft has been shifted in the axial direction. The contraction and expansion and contraction process greatly reduces unsteady circumferential pressure variation which is the major mechanism causing hot gas ingestion into the rim cavity.

In one embodiment, a gas turbine engine includes: a rotor with a rotor blade; a stator with a stator vane; a first seal point (39, 55) formed between the rotor and the stator; and a second seal point (38, 54) formed between the rotor and the stator, the first seal point (39, 55) including an expansion cavity (40, 56) formed in both the rotor and the stator; and the expansion cavity including sharp corners on an upstream side of the expansion cavity (40, 56).

In one aspect of the embodiment, the expansion cavity (40, 56) includes an inner expansion cavity (24, 51) on the rotor and an outer expansion cavity (32, 50) on the stator. In one aspect of the embodiment, the outer expansion cavity (50) is formed on a radial inward extension (45) of the rotor; and the inner expansion cavity (51) is formed on an axial extension (47) of the stator. In one aspect of the embodiment, the second seal point (54) is located radially outward from the first seal point (56); the second seal point (54) is formed from a radial outward extension (46) from the stator; and a buffer cavity (53) is formed between the radial inward extension (45) and the radial outward extension (46). In one aspect of the embodiment, a trench cavity (52) formed between the stator and the radial inward extension (45) in direct

communication with a hot gas path of the gas turbine engine.

In one aspect of the embodiment, the inner expansion cavity (24) is shorter than the outer expansion cavity (32). In one aspect of the embodiment, the first seal point (39) is located radially outward from the second seal point (38). In one aspect of the embodiment, the second seal point (38) is an angel wing seal. In one aspect of the embodiment, a buffer cavity (36) formed between the first seal point (39) and the second seal point (35). In one aspect of the embodiment, an axial extension (58) from the stator separates the buffer cavity (36) into an inner buffer cavity and an outer buffer cavity.

In one aspect of the embodiment, both the inner expansion cavity (24, 51) and the outer expansion cavity (32, 50) are formed by sharp corners on an inlet side and an outlet side of the expansion cavity (40, 56).

It will be appreciated by persons skilled in the art that the present invention is not limited to what has been particularly shown and described herein above. In addition, unless mention was made above to the contrary, it should be noted that all of the accompanying drawings are not to scale. A variety of modifications and variations are possible in light of the above teachings without departing from the scope and spirit of the invention, which is limited only by the following claims.

Claims

What is claimed is:

1. A gas turbine engine comprising:

a rotor with a rotor blade;

a stator with a stator vane;

a first seal point (39, 55) formed between the rotor and the stator; and a second seal point (38, 54) formed between the rotor and the stator, the first seal point (39, 55) including an expansion cavity (40, 56) formed in both the rotor and the stator; and

the expansion cavity including sharp corners on an upstream side of the expansion cavity (40, 56).

2. The gas turbine engine of claim 1, wherein:

the expansion cavity (40, 56) includes an inner expansion cavity (24, 51) on the rotor and an outer expansion cavity (32, 50) on the stator.

3. The gas turbine engine of claim 2, wherein:

the outer expansion cavity (50) is formed on a radial inward extension (45) of the rotor; and

the inner expansion cavity (51) is formed on an axial extension (47) of the stator.

4. The gas turbine engine of claim 3, wherein:

the second seal point (54) is located radially outward from the first seal point (56); the second seal point (54) is formed from a radial outward extension (46) from the stator; and a buffer cavity (53) is formed between the radial inward extension (45) and the radial outward extension (46).

5. The gas turbine engine of claim 4, further comprising:

a trench cavity (52) formed between the stator and the radial inward extension

(45) in direct communication with a hot gas path of the gas turbine engine.

6. The gas turbine engine of claim 2, wherein:

the inner expansion cavity (24) is shorter than the outer expansion cavity (32).

7. The gas turbine engine of claim 6, wherein:

the first seal point (39) is located radially outward from the second seal point (38).

8. The gas turbine engine of claim 7, wherein:

the second seal point (38) is an angel wing seal.

9. The gas turbine engine of claim 8, and further comprising:

a buffer cavity (36) formed between the first seal point (39) and the second seal point (35).

10. The gas turbine engine of claim 9, wherein:

an axial extension (58) from the stator separates the buffer cavity (36) into an inner buffer cavity and an outer buffer cavity.

11. The gas turbine engine of claim 2, wherein:

both the inner expansion cavity (24, 51) and the outer expansion cavity (32, 50) are formed by sharp corners on an inlet side and an outlet side of the expansion cavity (40, 56).