Adjustable flow split platform cooling for gas turbine engine

A turbine assembly may comprise a platform defining a platform cavity cooled by a platform flow with a vane extending from the platform. A platform-fed through cavity is defined by the vane and cooled by a first portion of the platform flow. A direct-fed through flow cavity is defined in the vane and cooled by a direct-fed through flow. The direct-fed through flow cavity and the platform-fed through cavity meet at an outlet to expel an outgoing through flow from the outlet. A platform-fed serpentine cavity may be defined in the vane and separated from the platform-fed through cavity by a divider. The platform-fed serpentine cavity is cooled by a second portion of the platform flow.

FIELD

The disclosure relates generally to gas turbine engines and, more particularly, to cooling airflow internal to turbomachinery components.

BACKGROUND

Engine components in gas turbines are exposed to high temperatures during use. Many components are internally cooled to counteract high ambient conditions. In conventional airfoil design cooling flow is fed through one airfoil and may be used to cool the next. These cooling feeds either come 100% from the platform cooling passages or 100% from the direct feed source. The cooling flows present difficulties when taking into consideration the airfoil body temperature with the temperature that is supplied to the airfoil downstream (e.g., exiting the airfoil). Controlling exit temperatures and/or pressures from cooling chambers may thus be difficult when the cooling flow is supplied 100% from the cooling cavities or 100% from a direct source.

SUMMARY

A turbine assembly is provided. The turbine assembly may comprise a platform defining a platform cavity cooled by a platform flow with a vane extending from the platform. A platform-fed through cavity is defined by the vane and cooled by a first portion of the platform flow. A direct-fed through flow cavity is defined in the vane and cooled by a direct-fed through flow. The direct-fed through flow cavity and the platform-fed through cavity meet at an outlet to expel an outgoing through flow from the outlet. A platform-fed serpentine cavity may be defined in the vane and separated from the platform-fed through cavity by a divider. The platform-fed serpentine cavity is cooled by a second portion of the platform flow.

In various embodiments, a blade may be disposed aft of the vane and be cooled by the outgoing through flow. A trailing edge cavity may be in fluid communication with the platform-fed serpentine cavity. The second portion of the platform flow is exhausted from the trailing edge cavity into a core flow. The through flow and the first portion of the platform flow mix at the outlet to form the outgoing through flow. The platform-fed through cavity may be in fluid communication with the platform cavity. The through flow may be extracted from a compressor stage of a gas turbine engine.

A gas turbine engine is also provided. A compressor may be configured to compress a core flow with a combustor in fluid communication with the compressor and configured to combust the core flow. A turbine may be disposed aft of the combustor and configured to expand the core flow. The turbine may include a platform defining a platform cavity cooled by a body flow. A vane may extend from the platform. A platform-fed through cavity may be defined by the vane and cooled by a first portion of the platform flow. A direct-fed through cavity may be defined in the vane and cooled by a through flow. The direct-fed through cavity and the platform-fed through cavity meet at an outlet. A platform-fed serpentine cavity may be defined in the vane and separated from the platform-fed through cavity. The platform-fed serpentine cavity is cooled by a second portion of the platform flow.

In various embodiments, a may be disposed blade aft of the vane in the turbine and cooled by the outgoing through flow. A trailing edge cavity may be in fluid communication with the platform-fed serpentine cavity. The second portion of the platform flow may be exhausted from the trailing edge cavity into the core flow. The through flow and the first portion of the platform flow mix at the outlet to form the outgoing through flow. The platform-fed through cavity may be in fluid communication with the platform cavity. The cooling flow source for direct-fed through cavities and/or for the platforms may be bled from the core flow in the compressor section.

A vane is also provided. The vane may have a vane platform defining a platform cavity, and a platform-fed through cavity defined by the vane and in fluid communication with the platform cavity. A direct-fed through cavity may be defined in the vane. The direct-fed through cavity and the platform-fed through cavity may meet at an outlet. A platform-fed serpentine cavity may be defined in the vane and separated from the platform-fed through cavity by a divider. The platform-fed serpentine cavity may be in fluid communication with the platform cavity.

In various embodiments, a trailing edge cavity may be fluid communication with the platform-fed serpentine cavity. The trailing edge cavity may be configured to exhaust a body flow. The outlet may be configured to mix a through flow from the direct-fed through cavity and a body flow from the platform-fed serpentine cavity. The platform-fed through cavity and the platform-fed serpentine cavity may be cooled by a body flow exiting the platform cavity. The direct-fed through cavity may be cooled by a direct cooling source.

DETAILED DESCRIPTION

The present disclosure relates to cooling components in gas turbine engines. The internally cooled components may draw cooling flow from two or more sources that meet and pass through the component as through flow. The use of separate cooling-flow sources enables balance of the pick-up of the head load between the two sources. If the platform-fed coolant flow through the airfoil is too hot, then some heat load may be absorbed by the direct-fed through flow. If the direct-fed through flow is too hot, then some of the platform heat load can be absorbed by the platform-fed through flow.

With reference toFIG. 1, a gas turbine engine20is shown according to various embodiments. Gas turbine engine20may be a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines may include, for example, an augmentor section among other systems or features. In operation, fan section22can drive coolant (e.g., air) along a path of bypass airflow B while compressor section24can drive coolant along a core flowpath C for compression and communication into combustor section26then expansion through turbine section28. Although depicted as a turbofan gas turbine engine20herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

Gas turbine engine20may generally comprise a low speed spool30and a high speed spool32mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure36or engine case via several bearing systems38,38-1, and38-2. Engine central longitudinal axis A-A′ is oriented in the z direction on the provided x-y-z axes. It should be understood that various bearing systems38at various locations may alternatively or additionally be provided, including for example, bearing system38, bearing system38-1, and bearing system38-2.

Low speed spool30may generally comprise an inner shaft40that interconnects a fan42, a low pressure compressor44and a low pressure turbine46. Inner shaft40may be connected to fan42through a geared architecture48that can drive fan42at a lower speed than low speed spool30. Geared architecture48may comprise a gear assembly60enclosed within a gear housing62. Gear assembly60couples inner shaft40to a rotating fan structure. High speed spool32may comprise an outer shaft50that interconnects a high pressure compressor52and high pressure turbine54. A combustor56may be located between high pressure compressor52and high pressure turbine54. A mid-turbine frame57of engine static structure36may be located generally between high pressure turbine54and low pressure turbine46. Mid-turbine frame57may support one or more bearing systems38in turbine section28. Inner shaft40and outer shaft50may be concentric and rotate via bearing systems38about the engine central longitudinal axis A-A′, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.

The airflow of core flowpath C may be compressed by low pressure compressor44then high pressure compressor52, mixed and burned with fuel in combustor56, then expanded over high pressure turbine54and low pressure turbine46. Turbines46,54rotationally drive the respective low speed spool30and high speed spool32in response to the expansion. The cooling flow source for direct-fed through cavities and/or for the platform cavities described herein may be bled from the core flow in compressor section24.

Gas turbine engine20may be, for example, a high-bypass ratio geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine20may be greater than about six (6). In various embodiments, the bypass ratio of gas turbine engine20may be greater than ten (10). In various embodiments, geared architecture48may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture48may have a gear reduction ratio of greater than about 2.3 and low pressure turbine46may have a pressure ratio that is greater than about five (5). In various embodiments, the bypass ratio of gas turbine engine20is greater than about ten (10:1). In various embodiments, the diameter of fan42may be significantly larger than that of the low pressure compressor44, and the low pressure turbine46may have a pressure ratio that is greater than about five (5:1). Low pressure turbine46pressure ratio may be measured prior to inlet of low pressure turbine46as related to the pressure at the outlet of low pressure turbine46prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other gas turbine engines including direct drive turbofans. A gas turbine engine may comprise an industrial gas turbine (IGT) or a geared aircraft engine, such as a geared turbofan, or non-geared aircraft engine, such as a turbofan, or may comprise any gas turbine engine as desired.

Referring now toFIG. 1and toFIG. 2, according to various embodiments, each of low pressure compressor44, high pressure compressor52, low pressure turbine46, and high pressure turbine54in gas turbine engine20may comprise one or more stages or sets of rotating blades and one or more stages or sets of stationary vanes axially interspersed with the associated blade stages but non-rotating about engine central longitudinal axis A-A′. Each compressor stage and turbine stage may comprise multiple interspersed stages of blades70and vanes90. The blades70rotate about engine central longitudinal axis A-A′, while the vanes90remain stationary with respect to engine central longitudinal axis A-A′. Blades70and vanes90may be referred to as airfoils100. For example,FIG. 2schematically shows, by example, a portion of an engine section68, which is illustrated as a turbine section28of gas turbine engine20.

With reference toFIG. 2, a schematic view of a portion of engine section68is shown, in accordance with various embodiments. Engine section68may include a circumferential array of blades70coupled about a circumference of a generally circular disk74. Disk74may be disposed radially inward of core flowpath C and centered on the rotation axis of the gas turbine engine. Disk74with blades70may be configured to rotate about engine central longitudinal axis A-A′. Each blade70may include an airfoil body76with a platform disposed at an inner diameter end wall72of the blade70. A disk cavity80may be defined between a forward disk and an aft disk. Upstream (forward) and downstream (aft) of blades70are circumferential arrays of vanes90configured to guide core flowpath C through the engine section68.

Each vane90may include an airfoil body96with an inner diameter platform94disposed at an inner diameter end wall92of vane90and with an outer diameter platform98disposed at an outer diameter end wall102of vane90. Outer diameter platform98may be coupled to engine case structure36. Inner diameter platform94and/or outer diameter platform98may be coupled to or integral with vane90. Airfoils100may be internally cooled engine components. Airfoils100may thus define internal cooling passages having serpentine geometry.

With reference toFIG. 3, cooling system120internal to vane90is shown, in accordance with various embodiments. Vane90may comprise an array of internal cooling cavities including a direct-fed through cavity124, a platform-fed through cavity126separated by the platform-fed serpentine cavity128by divider130, and a trailing edge cavity134.

In various embodiments, platform cavity125may be in fluid communication with platform-fed through cavity126and platform-fed serpentine cavity128. Platform-fed flow127may cool the platform surrounding platform cavity125and flow into platform-fed through cavity126and platform-fed serpentine cavity128defined in airfoil100. The relative percentages flowing into platform-fed through cavity126and platform-fed serpentine cavity128may be determined by divider130between the two cavities. Divider130may be positioned to meter the platform flow into the platform-fed flow cavities based on the relative cross-sectional area of the platform-fed through cavity126and platform-fed serpentine cavity128. Divider130may thus be a wall formed between platform-fed through cavity126and platform-fed serpentine cavity128.

In various embodiments, a platform-fed serpentine flow133may flow into platform-fed serpentine cavity128towards trailing edge cavity134. Platform-fed serpentine flow133may exit from trailing edge cavity134as exit flow136into core flow path C (ofFIG. 1).

In various embodiments, direct-fed through flow122from a direct coolant source may flow into direct-fed through cavity124towards outlet139. A platform-fed through flow132may flow into the platform-fed through cavity126towards outlet139. Platform-fed through flow132and direct-fed through flow122may meet and mix at outlet139and exit vane90as outgoing through flow138. Outgoing through flow138may be directed to an adjacent blade70(ofFIG. 2) to cool the blade.

In various embodiments, components cooled by the mixed flow of the present disclosure may allow the flexibility to reduce or increase through flow exit temperature, while the coolant temperature to the serpentine would respond in an opposite manner. The through flow may absorb additional heat load in response to the platform-fed flow temperature elevating to relatively high temperatures. The platform-fed flow may absorb additional heat load in response to the through flow temperature elevating to relatively high temperatures.

In various embodiments, cooling system120in airfoils may be made using an additive manufacturing technique such as direct metal laser sintering, selective laser sintering, selective laser melting, electron-beam melting, or electron-beam freeform fabrication. Casting may also be used to form cooling cavities of airfoil100. To cast an airfoil100or another internally cooled component having cooling passages, a core may be formed. The core of the component wall may have features of the cooling cavities. In that regard, cooling cavities may be formed in a core as positive material. The core may then be placed in a mold, and the material to form the internally cooled component may be deposited in the mold. The core may later be removed from the internally cooled component, leaving a cavity with the desired geometry. Airfoil100(as well as other internally cooled components) may be made from an austenitic nickel-chromium-based alloy or other materials capable of withstanding exhaust temperatures.