GAS TURBINE ENGINE COMPONENT PLATFORM COOLING

Abstract

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion. At least one augmentation feature is disposed on at least one of the leading edge portion and the trailing edge portion of the outer surface of the platform.

Description

This invention was made with government support under Contract No. F33615-03-D-2354-0017 awarded by the United States Air Force and Contract No. N00421-99-C-1270-0011 awarded by the United States Navy.

BACKGROUND

This disclosure relates generally to a gas turbine engine, and more particularly to a component that can be incorporated into a gas turbine engine. The component can include platform cooling augmentation features for cooling the platform of the component.

Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.

Both the compressor and turbine sections of the gas turbine engine may include alternating rows of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate to extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes prepare the airflow for the next set of blades.

Blades and vanes are examples of components that may need cooled in order to withstand the relatively high temperature of the hot combustion gases that are communicated along the core flow path. Typically, cooling is achieved by communicating a dedicated cooling airflow to select portions of the components.

SUMMARY

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion. At least one augmentation feature is disposed on at least one of the leading edge portion and the trailing edge portion of the outer surface of the platform.

In a further non-limiting embodiment of the foregoing gas turbine engine, the platform is an inner diameter platform.

In a further non-limiting embodiment of either of the foregoing gas turbine engines, the component is a turbine vane.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the outer surface is a non-gas path side of the platform and the inner surface is a gas path side of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, an airfoil extends from the inner surface of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the at least one augmentation feature includes a plurality of trip strips.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the plurality of trip strips are disposed at the trailing edge portion of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, each of the plurality of trip strips are angled relative to opposing mate faces of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, each of the plurality of trip strips include a first portion and a second portion that is transverse to the first portion.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the first portions are angled at a first angle relative to a mate face of the platform and the second portions are angled at a second angle different from the first angle relative to the mate face.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the at least one augmentation feature is disposed at the leading edge portion of the platform.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the at least one augmentation feature is disposed on a portion of the outer surface of the platform that axially overlaps a neighboring component of the gas turbine engine.

A gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor section in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor section. At least one of the compressor section and the turbine section includes a first component having a platform that includes an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion and a second component mounted adjacent to the first component and including a platform. A portion of the platform of the first component axially overlaps a portion of the platform of the second component, and the portion of the first platform includes at least one augmentation feature disposed on the outer surface of the platform.

In a further non-limiting embodiment of the foregoing gas turbine engine, the first component is a vane and the second component is a blade.

In a further non-limiting embodiment of either of the foregoing gas turbine engines, the at least one augmentation feature includes a plurality of trip strips.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the at least one augmentation feature is disposed on at least one of the leading edge portion and the trailing edge portion of the platform of the first component.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the at least one augmentation feature is angled relative to a mate face of the platform.

A method of cooling a component of a gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, cooling a platform of the component with a leakage airflow that is communicated from a cavity positioned radially inwardly from the component by circulating the leakage airflow over at least one augmentation feature that is disposed on an outer surface of the platform.

In a further non-limiting embodiment of the foregoing method of cooling a component of a gas turbine engine, the leakage airflow is not a dedicated cooling airflow that is communicated inside of the component.

In a further non-limiting embodiment of either of the foregoing methods of cooling a component of a gas turbine engine, the at least one augmentation feature is disposed on a trailing edge portion of the platform.

The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic, cross-sectional view of a gas turbine engine.

FIG. 2 illustrates a cross-section of a portion of a gas turbine engine.

FIG. 3 illustrates a component that can be incorporated into a gas turbine engine.

FIG. 4 illustrates another view of the component of FIG. 3.

FIG. 5 illustrates a platform of a component.

FIG. 6 illustrates another platform of a component.

FIG. 7 illustrates augmentation features that can be incorporated into a component of a gas turbine engine.

FIG. 8 illustrates additional augmentation features that can be incorporated into a component of a gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The exemplary gas turbine engine 20 is a two-spool turbofan engine that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmenter section (not shown) among other systems for features. The fan section 22 drives air along a bypass flow path B, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26. The hot combustion gases generated in the combustor section 26 are expanded through the turbine section 28. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures.

The gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine centerline longitudinal axis A. The low speed spool 30 and the high speed spool 32 may be mounted relative to an engine static structure 33 via several bearing systems 31. It should be understood that other bearing systems 31 may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 34 that interconnects a fan 36, a low pressure compressor 38 and a low pressure turbine 39. The inner shaft 34 can be connected to the fan 36 through a geared architecture 45 to drive the fan 36 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 35 that interconnects a high pressure compressor 37 and a high pressure turbine 40. In this embodiment, the inner shaft 34 and the outer shaft 35 are supported at various axial locations by bearing systems 31 positioned within the engine static structure 33.

A combustor 42 is arranged between the high pressure compressor 37 and the high pressure turbine 40. A mid-turbine frame 44 may be arranged generally between the high pressure turbine 40 and the low pressure turbine 39. The mid-turbine frame 44 can support one or more bearing systems 31 of the turbine section 28. The mid-turbine frame 44 may include one or more airfoils 46 that extend within the core flow path C.

The inner shaft 34 and the outer shaft 35 are concentric and rotate via the bearing systems 31 about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor 38 and the high pressure compressor 37, is mixed with fuel and burned in the combustor 42, and is then expanded over the high pressure turbine 40 and the low pressure turbine 39. The high pressure turbine 40 and the low pressure turbine 39 rotationally drive the respective high speed spool 32 and the low speed spool 30 in response to the expansion.

Each of the compressor section 24 and the turbine section 28 may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotating blades 25, while each vane assembly can carry a plurality of vanes 27 that extend into the core flow path C. The blades 25 of the rotor assemblies create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine 20 along the core flow path C. The vanes 27 of the vane assemblies direct the core air flow to the blades 25 to either add or extract energy.

Various components of the gas turbine engine 20, such as the blades 25 and the vanes 27 on the compressor section 24 and/or the turbine section 28, may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section 28 is particularly subjected to relatively extreme operating conditions. Therefore, some components may need cooled during engine operation. Example cooling features that can be incorporated into the components to improve cooling efficiency are described below.

FIG. 2 schematically illustrates a portion 100 of a gas turbine engine, such as the gas turbine engine 20 of FIG. 1. In this exemplary embodiment, the portion 100 represents part of the turbine section 28 of the gas turbine engine 20. However, it should be understood that other portions of the gas turbine engine 20 could benefit from the teachings of this disclosure, including but not limited to, the compressor section 24.

In this exemplary embodiment, the portion 100 includes a first component 50 and a second component 52 positioned adjacent to the first component 50. For example, the first component 50 may represent a vane (a generally static structure) and the second component 52 may represent a blade mounted for rotation about the engine centerline longitudinal axis A. Although only a single vane and a single blade are illustrated in FIG. 2, multiple vanes and blades could be circumferentially disposed about the engine centerline longitudinal axis A to provide vane and rotor assemblies. The portion 100 could also include additional, alternating rows of vanes and blades. FIG. 2 is highly schematic, and it should be understood that the various features depicted by this figure are not necessarily drawn to the scale they would be in practice.

In this embodiment, the first component 50 establishes a radially outer and radially inner flow path boundary of the core flow path C and directs the hot combustion gases communicated along the core flow path C to the second component 52. The second component 52 rotates to extract energy from the hot combustion gases that are communicated through the gas turbine engine 20.

The first component 50 and the second component 52 are mounted within the portion 100 such that a gap 56 extends between the first component 50 and the second component 52. A positive pressure can be maintained within the portion 100 by communicating a leakage airflow 60 into the gap 56. The leakage airflow 60 is communicated through a cavity 58 (positioned radially inwardly from the first component 50 and the second component 52) and then through the gap 56 to keep the hot combustion gases of the core flow path C from entering through the gap 56 and potentially damaging components. The leakage airflow 60 may be communicated from the compressor section 24 or some other upstream location of the gas turbine engine 20.

As is discussed in greater detail below, the leakage airflow 60 can also be used to cool portions of one or both of the first component 50 and the second component 52. In other words, the leakage airflow 60 that may be used to cool portions of the first component 50 and/or the second component 52 is not a dedicated cooling airflow that serves no other purpose other than to cool the component(s) 50, 52. In this disclosure, the term “dedicated cooling airflow” may refer to air which feeds the inside of the component(s) 50, 52 and the term “leakage airflow” may refer to airflow that bypasses the inside of the component(s) 50, 52, such as for purging cavities or preventing ingestion.

The first component 50 and/or the second component 52 can include cooling features that increase a local heat transfer effect of the first and/or second component 50, 52 without requiring a large flow pressure ratio. For example, in one embodiment, the first component 50 includes a platform 64A and the second component 52 includes a platform 64B. Each of the platforms 64A, 64B includes an outer surface 74 and an inner surface 76. In one embodiment, at least a portion of the second component 52 extends radially inward from, or under, the first component 50.

The platform 64A of the first component may include one or more augmentation features 78 disposed on the outer surface 74 for increasing the heat transfer effect between the platform 64A and the leakage airflow 60. The platform 64A can be cooled by circulating the leakage airflow 60 over the augmentation features 78. In one embodiment, the augmentation features 78 are disposed on a portion 55 of the platform 64A of the first component 50 that axially overlaps the platform 64B of the second component 52. The platform 64B could also include one or more augmentation features on its outer surface 74, although not shown in this embodiment.

FIGS. 3 and 4 illustrate a component 150 that can be incorporated into a gas turbine engine, such as the gas turbine engine 20 of FIG. 1. In this embodiment, the component 150 is a turbine vane similar to the first component 50 of FIG. 2. However, the various features described herein with respect to the component 150 could extend to other components of the gas turbine engine 20, including but not limited to blades (i.e., the second component 52 of FIG. 2).

The component 150 of this embodiment includes an inner diameter platform 62, an outer diameter platform 64 and an airfoil 66 that extends between the inner diameter platform 62 and the outer diameter platform 64. Each of the inner diameter platform 62 and the outer diameter platform 64 includes a leading edge portion 68, a trailing edge portion 70 and opposing mate faces 71, 73. The inner diameter and outer diameter platforms 62, 64 axially extend between the leading edge portion 68 and the trailing edge portion 70 and circumferentially extend between the opposing mate faces 71, 73. The opposing mate faces 71, 73 can be mounted relative to corresponding mate faces of adjacent components of a gas turbine engine to provide a full ring assembly, such as a full ring vane assembly that can be circumferentially disposed about the engine centerline longitudinal axis A (see FIG. 1).

The inner diameter and outer diameter platforms 62, 64 can also include an outer surface 74 (for example, a non-gas path side) and an inner surface 76 (a gas path side). In other words, when the component 150 is mounted within a gas turbine engine 20, the outer surfaces 74 are positioned on a non-core flow path side of the component 150, while the inner surfaces 76 establish the outer boundaries of the core flow path C of the gas turbine engine 20.

One or both of the inner diameter platform 62 and the outer diameter platform 64 can include one or more augmentation features 78 disposed on the outer surfaces 74 of the inner diameter platform 62 and/or the outer diameter platform 64. In this embodiment, the augmentation features 78 are positioned at the trailing edge portion 70 of the inner diameter platform 62 (see FIG. 2 and FIG. 4). In another embodiment, the augmentation features 78 may be disposed at the leading edge portion 68 of the inner diameter platform 62 and/or the outer diameter platform 64 (see FIG. 5). In yet another embodiment, the augmentation features 78 may be formed on both the leading edge portion 68 and the trailing edge portion 70 of the inner diameter platform 62 and/or the outer diameter platform 64 (see FIG. 6). The augmentation features 78 may be disposed on any portion of the inner diameter platform 62 and/or the outer diameter platform 64, including any portion of the inner or outer diameter platforms 62, 64 that axially overlap a neighboring component (see portion 55 of FIG. 2, for example).

The exemplary augmentation features 78 can include any heat transfer augmentation features including but not limited to trip strips, pin fins, chevron trip strips, or any combination of features. In this embodiment, the augmentation features 78 include trip strips. The augmentation features 78 turbulate the flow of the leakage airflow 60 that comes into contact with the component 150 (as shown in FIG. 2) and adds surface area to platform(s) 62, 64 to enhance heat transfer between the leakage airflow 60 and the platform(s) 62, 64.

FIG. 7 illustrates additional features of the augmentation features 78 described above. The illustrated platform 62, 64 could be representative of either an inner diameter platform or an outer diameter platform. In this embodiment, the augmentation features 78 include a plurality of trip strips 78A that are circumferentially spaced between the opposing mate faces 71, 73 at a trailing edge portion 70 of the platform 62, 64. The trip strips 78A extend parallel to one another and may each be angled at an angle α relative to the opposing mate faces 71, 73. The value of the angle α can vary depending on design specific criteria, including but not limited to the amount of heat transfer required to cool the platform 62, 64. The leakage airflow 60 may be circulated over the trip strips 78A in both a circumferential direction CD as well as an axial direction AD to cool the platform 62, 64.

FIG. 8 illustrates another plurality of augmentation features 178 that can be incorporated into a platform 62, 64 of a component 150. In this embodiment, the augmentation features 178 include a plurality of trip strips 178A that are circumferentially spaced between the opposing mate faces 71, 73 at a trailing edge portion 70 of the platform 62, 64. The plurality of trip strips 178A of this embodiment include a first portion 80 and a second portion 82 that is transverse to the first portion 80. The first portions 80 of the plurality of trip strips 178A extend parallel to one another and may each be angled at an angle α relative to the opposing mate faces 71, 73. The second portions 82 of the plurality of trip strips 178A also extend parallel to one another and may be angled at an angle β relative to the opposing mate faces 71, 73. In this embodiment, the angle β is a different angle than the angle α. The second portions 82 of the plurality of trip strips 178A direct the leakage airflow 60 into the core flow path C by directing the leakage airflow 60 circumferentially along the direction of rotation of a neighboring component (shown schematically via arrow 110), such as a blade (see the second component 52 of FIG. 2), thereby providing improved heat transfer and reducing mixing loss.

Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the non-limiting embodiments.

It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure.

The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would recognize that various modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.

Claims

1. A component for a gas turbine engine, comprising:

a platform having an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion; and

at least one augmentation feature disposed on at least one of said leading edge portion and said trailing edge portion of said outer surface of said platform.

2. The component as recited in claim 1, wherein said platform is an inner diameter platform.

3. The component as recited in claim 1, wherein the component is a turbine vane.

4. The component as recited in claim 1, wherein said outer surface is a non-gas path side of said platform and said inner surface is a gas path side of said platform.

5. The component as recited in claim 1, comprising an airfoil that extends from said inner surface of said platform.

6. The component as recited in claim 1, wherein said at least one augmentation feature includes a plurality of trip strips.

7. The component as recited in claim 6, wherein said plurality of trip strips are disposed at said trailing edge portion of said platform.

8. The component as recited in claim 6, wherein each of said plurality of trip strips are angled relative to opposing mate faces of said platform.

9. The component as recited in claim 6, wherein each of said plurality of trip strips include a first portion and a second portion that is transverse to the first portion.

10. The component as recited in claim 9, wherein said first portions are angled at a first angle relative to a mate face of said platform and said second portions are angled at a second angle different from said first angle relative to said mate face.

11. The component as recited in claim 1, wherein said at least one augmentation feature is disposed at said leading edge portion of said platform.

12. The component as recited in claim 1, wherein said at least one augmentation feature is disposed on a portion of said outer surface of said platform that axially overlaps a neighboring component of the gas turbine engine.

13. A gas turbine engine, comprising:

a compressor section;

a combustor section in fluid communication with said compressor section;

a turbine section in fluid communication with said combustor section; and

wherein at least one of said compressor section and said turbine section includes: a first component having a platform that includes an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion; a second component mounted adjacent to said first component and including a platform; wherein a portion of said platform of said first component axially overlaps a portion of said platform of said second component, and said portion of said first platform includes at least one augmentation feature disposed on said outer surface of said platform.

14. The gas turbine engine as recited in claim 13, wherein said first component is a vane and said second component is a blade.

15. The gas turbine engine as recited in claim 13, wherein said at least one augmentation feature includes a plurality of trip strips.

16. The gas turbine engine as recited in claim 13, wherein said at least one augmentation feature is disposed on at least one of said leading edge portion and said trailing edge portion of said platform of said first component.

17. The gas turbine engine as recited in claim 13, wherein said at least one augmentation feature is angled relative to a mate face of said platform.

18. A method of cooling a component of a gas turbine engine, comprising the steps of:

cooling a platform of the component with a leakage airflow that is communicated from a cavity positioned radially inwardly from the component by circulating the leakage airflow over at least one augmentation feature that is disposed on an outer surface of the platform.

19. The method as recited in claim 18, wherein the leakage airflow is not a dedicated cooling airflow that is communicated inside of the component.

20. The method as recited in claim 18, wherein the at least one augmentation feature is disposed on a trailing edge portion of the platform.