Gas turbine engine component providing prioritized cooling

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having a non-gas path surface and a gas path surface, a cover plate positioned relative to the non-gas path surface and a cooling passage that extends between the cover plate and the non-gas path surface. At least one cooling entrance is formed through the cover plate and configured to bias the flow of a cooling fluid toward a prioritized location of the non-gas path surface.

BACKGROUND

This disclosure relates generally to a cooled gas turbine engine component. More particularly, this disclosure is directed to a gas turbine engine component capable of providing prioritized cooling at preferred locations (i.e., relatively hotter surfaces) of the component prior to cooling less preferred locations (i.e., relatively cooler surfaces) 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 a 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. The turbine vanes direct the combustion gases at a preferred angle of entry into the downstream row of blades. Blades and vanes are examples of components that may need cooled by a dedicated source of cooling air in order to withstand the relatively high temperatures of the hot combustion gases they are exposed to.

SUMMARY

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having a non-gas path surface and a gas path surface, a cover plate positioned relative to the non-gas path surface and a cooling passage that extends between the cover plate and the non-gas path surface. At least one cooling entrance is formed through the cover plate and configured to bias the flow of a cooling fluid toward a prioritized location of the non-gas path surface.

In a further non-limiting embodiment of the foregoing component, the platform is an outer diameter platform or an inner diameter platform.

In a further non-limiting embodiment of either of the foregoing components, the cover plate is a separate component from the platform.

In a further non-limiting embodiment of any of the foregoing components, the component is static hardware of the gas turbine engine.

In a further non-limiting embodiment of any of the foregoing components, an airfoil extends from the gas path side of the platform.

In a further non-limiting embodiment of any of the foregoing components, the cover plate is received at least partially about an opening of an internal core of the airfoil.

In a further non-limiting embodiment of any of the foregoing components, the at least one cooling entrance is positioned exclusively at a first location of the cover plate and the cover plate excludes cooling openings outside of the first location.

In a further non-limiting embodiment of any of the foregoing components, a plurality of cooling entrances are positioned near a leading edge of the cover plate.

In a further non-limiting embodiment of any of the foregoing components, the plurality of cooling entrances are triangular shaped.

In a further non-limiting embodiment of any of the foregoing components, the prioritized location is at a leading edge of the platform.

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor section and a turbine section downstream from the compressor section. A component is positioned within at least one of the compressor section and the turbine section. The component includes a platform having a non-gas path surface and a gas path surface. A cover plate is positioned adjacent to the non-gas path surface and a cooling passage extends between the cover plate and the non-gas path surface. A plurality of cooling entrances are positioned exclusively at a first location of the cover plate. The cover plate excludes cooling openings outside of the first location.

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

In a further non-limiting embodiment of either of the foregoing gas turbine engines, the component is a blade outer air seal (BOAS).

In a further non-limiting embodiment of any of the gas turbine engines, the plurality of cooling entrances are configured to bias the flow of a cooling fluid toward a prioritized location of the non-gas path surface.

In a further non-limiting embodiment of any of the gas turbine engines, at least one opening is configured to expel a cooling fluid from the cooling passage to either a core flow path of the gas turbine engine or a second cooling passage.

A method of cooling a component of a gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, communicating a cooling fluid through at least one cooling entrance of a cover plate and into a first cooling passage and cooling a prioritized location of a platform with the cooling fluid. After the step of cooling, the cooling fluid is communicated to a non-prioritized location of the platform and expelled from the cooling passage.

In a further non-limiting embodiment of the foregoing method, the step of communicating the cooling fluid through the at least one cooling entrance includes altering a flow characteristic of the cooling fluid.

In a further non-limiting embodiment of either of the foregoing methods, the step of communicating the cooling fluid to the non-prioritized location includes circulating the cooling fluid to convectively cool the non-prioritized location.

In a further non-limiting embodiment of any of the foregoing methods, the step of expelling includes communicating the cooling fluid to a core flow path of the gas turbine engine.

In a further non-limiting embodiment of any of the foregoing methods, the step of expelling includes communicating the cooling fluid to a second cooling passage.

The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following descriptions and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.

DETAILED DESCRIPTION

This disclosure relates to a gas turbine engine component that prioritizes cooling at specific locations of the component. For example, the gas turbine engine components described by this disclosure may employ a cover plate having a plurality of cooling entrances for directing a cooling fluid to preferred locations of a platform. The shape and location of the cooling entrances may be configured such that the cooling fluid is supplied first to the relatively hotter surfaces of the component before passing over its relatively cooler surfaces. These and other features are described in detail herein.

FIG. 1schematically illustrates a gas turbine engine20. The exemplary gas turbine engine20is a two-spool turbofan engine that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section22drives air along a bypass flow path B, while the compressor section24drives air along a core flow path C for compression and communication into the combustor section26. The hot combustion gases generated in the combustor section26are expanded through the turbine section28. Although depicted as a turbofan gas turbine engine in this 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 engine20generally includes a low speed spool30and a high speed spool32mounted for rotation about an engine centerline longitudinal axis A. The low speed spool30and the high speed spool32may be mounted relative to an engine static structure33via several bearing systems31. It should be understood that other bearing systems31may alternatively or additionally be provided.

The low speed spool30generally includes an inner shaft34that interconnects a fan36, a low pressure compressor38and a low pressure turbine39. The inner shaft34can be connected to the fan36through a geared architecture45to drive the fan36at a lower speed than the low speed spool30. The high speed spool32includes an outer shaft35that interconnects a high pressure compressor37and a high pressure turbine40. In this embodiment, the inner shaft34and the outer shaft35are supported at various axial locations by bearing systems31positioned within the engine static structure33.

A combustor42is arranged between the high pressure compressor37and the high pressure turbine40. A mid-turbine frame44may be arranged generally between the high pressure turbine40and the low pressure turbine39. The mid-turbine frame44can support one or more bearing systems31of the turbine section28. The mid-turbine frame44may include one or more airfoils46that extend within the core flow path C.

The inner shaft34and the outer shaft35are concentric and rotate via the bearing systems31about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor38and the high pressure compressor37, is mixed with fuel and burned in the combustor42, and is then expanded over the high pressure turbine40and the low pressure turbine39. The high pressure turbine40and the low pressure turbine39rotationally drive the respective high speed spool32and the low speed spool30in response to the expansion.

The pressure ratio of the low pressure turbine39can be measured prior to the inlet of the low pressure turbine39as related to the pressure at the outlet of the low pressure turbine39and prior to an exhaust nozzle of the gas turbine engine20. In one non-limiting embodiment, the bypass ratio of the gas turbine engine20is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor38, and the low pressure turbine39has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines, including direct drive turbofans.

In this embodiment of the exemplary gas turbine engine20, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section22of the gas turbine engine20is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine20at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section22without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine20is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [Tram° R)/(518.7° R)]0.5. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine20is less than about 1150 fps (351 m/s).

Each of the compressor section24and the turbine section28may 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 blades25, while each vane assembly can carry a plurality of vanes27that extend into the core flow path C. The blades25create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine20along the core flow path C. The vanes27direct the core airflow to the blades25to either add or extract energy.

Various components of the gas turbine engine20, including but not limited to the airfoil and platform sections of the blades25and vanes27of the compressor section24and the turbine section28, may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section20is particularly subjected to relatively extreme operating conditions. Therefore, some components may require dedicated internal cooling circuits to cool the parts during engine operation. This disclosure relates to gas turbine engine components having cooling configurations capable of providing prioritized cooling at specific locations of a component, as is further discussed below.

FIG. 2illustrates a component60that can be incorporated into a gas turbine engine, such as the gas turbine engine20ofFIG. 1. In this non-limiting embodiment, the component60is represented as a turbine vane. However, the teachings of this disclosure are not limited to turbine vanes and could extend to other components of a gas turbine engine, including but not limited to, compressor vanes, blade outer air seals (BOAS) (see, for example, the BOAS illustrated inFIG. 6), or any other static hardware.

In one embodiment, the component60includes an outer platform62, an inner platform64, and an airfoil66that extends between the outer platform62and the inner platform64. The outer platform62connects the component60to an engine casing (not shown), and the inner platform64affixes a radially inboard portion of the component60to securely position the component60within the core flow path C.

The outer platform62and the inner platform64each include a non-gas path surface68and a gas path surface70. In other words, when mounted in a gas turbine engine, the non-gas path surfaces68are positioned on a non-core flow path side of the component60(i.e., remotely from the hot combustion gases of the core flow path C). The gas path surfaces70establish the radially inner and outer flow boundaries of the core flow path C.

The component60may additionally include one or more internal cores72. The internal cores72are hollow passages formed inside of the component60that channel a cooling fluid F (e.g., a bleed airflow from the compressor section24) to remove heat from the component60in a heat transfer process. The internal cores72may extend through the airfoil66to cool the airfoil66. In one embodiment, as best shown with respect to the outer platform62, the internal cores72may jut outwardly from the outer platform62and the inner platform64to define openings65for receiving the cooling fluid F.

The outer platform62and the inner platform64may additionally require dedicated cooling. Accordingly, a cover plate74may be positioned relative to each of the outer platform62and the inner platform64. The cover plate74may be a separate component form the outer and inner platforms62,64or could be cast integrally with each of the outer and inner platforms62,64.

In one embodiment, the cover plates74are positioned relative to the non-gas path surfaces68of the outer and inner platforms62,64. Once positioned relative to the non-gas path surface68, the cover plates74establish a platform cooling passage (see, for example, platform cooling passage86ofFIG. 4) that extends between an interior surface94of the cover plate74and the non-gas path surface68. The cooling fluid F may be communicated through the cover plate74and into a platform cooling passage to cool portions of the outer and inner platforms62,64. In one embodiment, as discussed in greater detail below, the component60may prioritize cooling at preferred locations (i.e., relatively high temperature surfaces) of the component60prior to cooling less preferred locations (i.e., relatively low temperature surfaces) of the component60.

FIGS. 3 and 4illustrate an exemplary cover plate74that can be utilized to establish a platform cooling passage86associated with a platform62(shown inFIG. 4). In this disclosure, the term “platform” is intended to include any portion of a component that has a surface that is exposed to the core flow path C of a gas turbine engine and therefore may require dedicated cooling. For example, the platform62could be part of a vane (seeFIGS. 2 and 4), a BOAS (seeFIG. 6), or other static hardware within the scope of this disclosure.

In one embodiment, the cover plate74includes a leading edge76, a trailing edge78and a main body80that extends between the leading edge76and the trailing edge78. An opening82may be formed in the main body80for accommodating an opening in a platform, such as the opening65of an internal core72(seeFIG. 2).

In one embodiment, the leading edge76is angled relative to the main body80. In other words, in a non-limiting embodiment, the leading edge76is elevated relative to the main body80. However, the cover plate74can be configured to accommodate the size and shape of any platform of any gas turbine engine component.

The cover plate74may additionally include one or more cooling entrances84for directing cooling fluid F into the platform cooling passage86. In one embodiment, the cooling entrances84are triangular shaped and are arranged in multiple rows. However, the cooling entrances84may include any shape and may be configured in any pattern. In one embodiment, the cooling entrances84are configured to alter a flow characteristic of the cooling fluid F. For example, the cooling entrances84may impart a swirling motion to the cooling fluid F, thereby increasing its cooling effectiveness.

In one non-limiting embodiment, a plurality of cooling entrances84are disposed exclusively at a first location L1of the cover plate74. The first location L1may be near the leading edge76of the cover plate74. However, other locations are also contemplated depending on the cooling requirements of the component60. The remainder of the cover plate74, including the main body80and the trailing edge78, completely excludes any cooling openings for directing the cooling fluid F into the platform cooling passage86. In this way, the cooling fluid F that is directed through the cooling entrances84is communicated directly to a prioritized location P1, such as the hottest locations of the platform62that are most susceptible to thermal distresses, prior to being used to cool other, non-prioritized locations P2of the platform62that are generally located at relatively cooler locations of the platform62.

Referring toFIG. 4, the cover plate74is positioned relative to the non-gas path surface68of the platform62to establish the platform cooling passage86. Although illustrated as an outer platform, the platform62could include an inner platform or any other section of a component. The cover plate74may be positioned such that a portion of the leading edge76rests against a rail88of the platform62and a portion of the trailing edge78rests against a land90that extends from the non-gas path surface68to another rail92. The main body80of the cover plate74is spaced from the non-gas path surface68to define the platform cooling passage86between an interior surface94of the cover plate74and the non-gas path surface68.

In one non-limiting embodiment, the non-gas path surface68of the platform62of the component60may be cooled by communicating cooling fluid F through the plurality of cooling entrances84formed in the cover plate74. The cooling fluid F is first directed to the prioritized location P1of the platform62. This can be accomplished by locating the cooling entrances84exclusively in the vicinity of the prioritized location P1. In one non-limiting embodiment, the prioritized location P1is at a leading edge96of the platform62and the cooling entrances84are therefore disposed at the leading edge76of the cover plate74. The cooling fluid F may at least partially impinge upon the prioritized location P1.

After first cooling the prioritized location P1, the cooling fluid F can be communicated to cool the non-prioritized locations P2of the platform62. In this case, the cooling fluid F is axially circulated toward a trailing edge98of the platform62in order to convectively cool the non-prioritized locations P2of the non-gas path surface68. One or more augmentation features100may be protrude from the non-gas path surface68within the platform cooling passage86in order to increase the effective heat transfer between the cooling fluid F and the platform62. Suitable augmentation features include but are not limited to domes, trip strips, pins or fins.

Finally, the cooling fluid F may be expelled from the platform cooling passage86. In one embodiment, the cooling fluid F is expelled through one or more film cooling holes102and is returned to the core flow path C. Alternatively, as shown inFIG. 5, the cooling fluid F may be expelled through an opening104into an adjacent cooling passage106in order to cool other surfaces of the platform62.

FIG. 6illustrates another component160that can be incorporated into a gas turbine engine. In this embodiment, the component160is represented as a BOAS. The component160includes a platform162and a cover plate174that can be positioned relative to a non-gas path surface168of the platform162to define a platform cooling passage186. Similar to the component60, the cover plate174may include a plurality of cooling entrances184that are configured to bias a flow of a cooling fluid F toward a prioritized location P1of the platform162prior to directing the cooling fluid F toward the non-prioritized locations P2of the platform162.

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 other 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.