Patent Description:
The present invention relates to gas turbine engines and, more particularly, to systems and methods used to cool airfoils within gas turbine engines.

A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.

Turbine section components, such as turbine blades and vanes, are operated in high temperature environments. To avoid deterioration in the components resulting from their exposure to high temperatures, cooling circuits are typically employed within the components. Turbine blades and vanes are subjected to high thermal loads on both the suction and pressure sides of the airfoil portions and at both the leading and trailing edges. The regions of the airfoils having the highest thermal loads can differ depending on engine design and specific operating conditions.

Minicore technology offers the potential to provide higher specific cooling passages for turbine components such as blade and vane airfoils, blade outer air seals and combustor or gas path panels. Minicore technology typically employs refractory metal cores to allow cooling circuits to be placed just under the surface of the hot wall through which cooling air flows before being expelled into the gas path. State of the art cooling circuits made using investment casting techniques and refractory metal cores may, however, contain defects or artifacts following manufacture. One such defect or artifact is a linear ridge that can form at the exit portion of a diffuser hole connected to the minicore. The linear ridge may, however, disrupt the smooth flow of air as it passes out of the diffuser hole and into in the main gas path. In order to prevent such flow disruption, additional fabrication steps are typically required to remove the linear ridge, resulting in increased time required and cost incurred during the fabrication of the turbine components.

<CIT> discloses an apparatus and method relating to a cooling hole of a component of a turbine engine.

<CIT> discloses a gas turbine engine component including a wall having first and second wall surfaces and a cooling hole extending through the wall.

<CIT> discloses a method for producing a diffusion cooling hole extending between a wall having a first wall surface and a second wall surface including forming a cooling hole inlet at the first wall surface, forming a cooling hole outlet at the second wall surface, forming a metering section downstream from the inlet and forming a multi-lobed diffusing section between the metering section and the outlet.

<CIT> discloses a component for a gas turbine engine including a wall and a cooling hole.

According to an aspect of the invention, there is provided a component for a gas turbine engine. The component includes a cooling passage; an outer wall separating a core flow path from the cooling passage; a minicore exit aperture in fluid communication with the cooling passage and opening into the core flow path, the minicore exit aperture being characterized by a linear ridge on a downstream end of the minicore exit aperture; and a thermal barrier coating covering the outer wall and the linear ridge. The linear ridge includes an upstream facing side that is characterized by a height, the height having a value within a range of between five one-hundredths and seventy-five one-hundredths of a depth of the cooling passage.

In various embodiments, the minicore exit aperture defines a rectangular shape in a direction normal to the outer wall and the linear ridge extends perpendicular to the cooling passage along the downstream end of the minicore exit aperture. In various embodiments, the thermal barrier coating includes a first portion upstream of the linear ridge, the first portion extending from the cooling passage and being characterized by a first radius of curvature and the thermal barrier coating includes a second portion, the second portion extending from the first portion and over the linear ridge and being characterized by a second radius of curvature.

In various embodiments, the thermal barrier coating is configured to transition from the first portion to the second portion at an inflection line. In various embodiments, the minicore exit aperture includes an upstream wall, the upstream wall extending from the cooling passage to the outer wall and being characterized by a third radius of curvature, the third radius of curvature being greater than the depth of the cooling passage. In various embodiments, the first radius of curvature is greater than the third radius of curvature. In various embodiments, the second radius of curvature is less than the first radius of curvature.

Referring now to the drawings, <FIG> schematically illustrates a gas turbine engine <NUM>, in accordance with various embodiments. The fan section <NUM> drives air along a bypass flow path B in a bypass duct defined within a nacelle <NUM>, while the compressor section <NUM> drives air along a primary or core flow path C for compression and communication into the combustor section <NUM> and then expansion through the turbine section <NUM>. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it will be understood that the concepts described herein are not limited to use with two-spool turbofans, as the teachings may be applied to other types of gas turbine engines, including, for example, architectures having three or more spools or only a single spool.

The gas turbine engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that various bearing systems at various locations may alternatively or additionally be provided and the location of the several bearing systems <NUM> may be varied as appropriate to the application. The inner shaft <NUM> may be directly connected to the fan <NUM> or through a speed change mechanism, such as, for example, a fan drive gear system configured to drive the fan <NUM> at a lower speed than that of the low speed spool <NUM>. The high speed spool <NUM> generally includes an outer shaft <NUM> that interconnects a high pressure compressor <NUM> and a high pressure turbine <NUM>. A combustor <NUM> is arranged in the gas turbine engine <NUM> between the high pressure compressor <NUM> and the high pressure turbine <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the several bearing systems <NUM> about the engine central longitudinal axis A, which is collinear with longitudinal axes of the inner shaft <NUM> and the outer shaft <NUM>.

The air in the core flow path C is compressed by the low pressure compressor <NUM> and then the high pressure compressor <NUM>, mixed and burned with fuel in the combustor <NUM>, and then expanded over the high pressure turbine <NUM> and the low pressure turbine <NUM>. The low pressure turbine <NUM> and the high pressure turbine <NUM> rotationally drive the respective low speed spool <NUM> and the high speed spool <NUM> in response to the expansion. It will be appreciated that each of the positions of the fan section <NUM>, the compressor section <NUM>, the combustor section <NUM>, the turbine section <NUM>, and the fan drive gear system, if present, may be varied. For example, the fan drive gear system may be located aft of the combustor section <NUM> or even aft of the turbine section <NUM>, and the fan section <NUM> may be positioned forward or aft of the location of the fan drive gear system.

Referring now to <FIG>, a pair of turbine vanes <NUM> (e.g., a first turbine vane <NUM> and a second turbine vane <NUM>) are schematically illustrated. The pair of turbine vanes <NUM> are representative of the vanes present in either of the low pressure turbine <NUM> and the high pressure turbine <NUM> described above with reference to <FIG>. While the present disclosure will be described with respect to its application to a turbine vane, the disclosure could also be utilized in a rotating structure such as a turbine blade (e.g., the turbine blades present in either of the low pressure turbine <NUM> and the high pressure turbine <NUM>) or other static turbine components such as blade outer air seals, turbine exhaust cases and struts. Additional uses of the cooling scheme may include combustor liners and flame holders as well as nozzle liners and flaps. In various embodiments, each of the pair of turbine vanes <NUM> includes an outer wall <NUM> through which are formed minicore slots or exit apertures <NUM> for exhausting a cooling air (or other fluid) from a plurality of minicores <NUM> formed in the outer wall <NUM> of each of the pair of turbine vanes <NUM>. Between adjacent pairs of minicores is a rib or a web <NUM>. A baffle <NUM> having a plurality of baffle apertures <NUM> formed therethrough is disposed within each of the pair of turbine vanes <NUM>. Each of the pair of turbine vanes <NUM> also includes a baffle inlet <NUM> at a radial end. The baffle apertures <NUM> are arranged and positioned in order to direct a cooling fluid CF directly onto the web <NUM>, and are sized in order to allow sufficient fluid flow to fill the plurality of minicores <NUM>.

Referring now to <FIG>, with continued reference to <FIG>, a sectional view through one of the pair of turbine vanes <NUM> is illustrated. The plurality of minicores <NUM> (e.g., a first minicore <NUM> and a second minicore <NUM>) is formed in the outer wall <NUM> of each of the pair of turbine vanes <NUM>. The baffle <NUM> is disposed inside each of the pair of turbine vanes <NUM>, and spaced inwardly of the outer wall <NUM>. Referring now to <FIG>, with continued reference to <FIG> and <FIG>, operation of the baffle <NUM> and each of the plurality of minicores <NUM> inside the pair of turbine vanes <NUM> is described. The cooling fluid CF (e.g., a high-pressure flow of air bled from the compressor section <NUM> of the gas turbine engine <NUM> described above with reference to <FIG>), is directed into the baffle inlet <NUM> of the baffle <NUM>. The cooling fluid CF is then directed from the baffle apertures <NUM> through the baffle <NUM> and directly onto the web <NUM> in the outer wall <NUM> between adjacent pairs of the plurality of minicores <NUM>. The cooling fluid CF then flows into each of the plurality of minicores <NUM> through a minicore inlet aperture <NUM>. Inside each of the plurality of minicores <NUM>, the cooling fluid CF flows generally parallel to an outer surface of the outer wall <NUM> before being exhausted through a minicore exit aperture <NUM> associated with each of the plurality of minicores <NUM>. Although each of the plurality of minicores <NUM> is shown having a single minicore inlet aperture and a single minicore exit aperture, each of the plurality of minicores <NUM> could have a plurality of minicore inlet apertures or a plurality of minicore exit apertures. The dimensions and spacing of the baffle apertures <NUM> (one shown) are such that the heat transfer coefficients generated provide a heat flux that is comparable to that achieved by each of the plurality of minicores <NUM>. As a result, the outer wall <NUM> exhibits lower overall thermal gradients, which tends to reduce thermal mechanical fatigue and increase oxidation life by lowering peak surface temperatures. As illustrated in <FIG>, and as described in more detail below with reference to <FIG>, a linear ridge <NUM> that extends laterally across the minicore exit aperture <NUM> and has an upstream facing face <NUM> may be formed during fabrication of the turbine vane and, in various embodiments, is employed to help anchor a thermal barrier coating <NUM> following fabrication.

Referring now to <FIG>, a cross-sectional view of a portion of a turbine vane <NUM> (e.g., one of the pair of turbine vanes <NUM> described above with reference to <FIG>) is illustrated. The turbine vane <NUM> includes a minicore <NUM> having a minicore exit aperture <NUM> opening from an outer wall <NUM> of the turbine vane <NUM> and into a core flow path C. In various embodiments, the outer wall <NUM> defines a hot side surface <NUM> in contact with the core flow path C and a cold side surface <NUM> that faces a cooling fluid CF. In various embodiments, the cold side surface <NUM> can form a portion of a cooling passage <NUM> of the minicore <NUM>. In various embodiments, the minicore exit aperture <NUM> forms a diffuser in fluid communication with the cooling passage <NUM>, the diffuser having a rectangular shape in a direction normal to the outer wall <NUM>. In various embodiments, a thermal barrier coating <NUM> is applied to the outer wall <NUM> at locations both upstream and downstream of the minicore exit aperture <NUM>, with the outer surface of the thermal barrier coating forming the hot side surface <NUM>. In various embodiments, the thermal barrier coating <NUM> may comprise, for example, a coating having yttria-stabilized zirconia, ytterbium zirconium, fully-stabilized gadolinia zirconia, alumina, pyrochlores, or combinations thereof. In various embodiments, the coating further includes a bonding layer, for example, a MCrAlY alloy (where M identifies one or more of Fe, Ni, and Co), intermetallic aluminide, or any other suitable material. The coating may be applied to the outer wall <NUM> by any suitable process, such as, for example, physical vapor deposition, chemical vapor deposition, cold spray, or a combination thereof.

As described previously, during fabrication of the turbine vane <NUM>, a defect or an artifact in the form of a linear ridge <NUM> (or a slot exit ridge) can form along the exit portion, generally perpendicular to the cooling passage <NUM> or to the flow of the cooling fluid CF through the cooling passage <NUM> at a downstream end of the minicore exit aperture <NUM>. In various embodiments, the linear ridge <NUM> includes an upstream facing side <NUM> that is generally perpendicular to the hot side surface <NUM> and an outer wall side <NUM> that is generally parallel to and contiguous with the outer surface of the outer wall <NUM>. Rather than machine the ridge away to improve the flow aerodynamics of the cooling fluid CF as it exits the minicore exit aperture <NUM>, the thermal barrier coating <NUM> is applied to the outer wall <NUM> to cover both the outer surface of the outer wall <NUM> and the linear ridge <NUM>, with the linear ridge <NUM> helping to anchor the thermal barrier coating <NUM> in the vicinity of the minicore exit aperture <NUM>. In various embodiments, the linear ridge <NUM> may result as a defect or an artifact from the casting process or may be intentionally incorporated into the design of the casting process. Without the linear ridge <NUM> being present, a continuous slope <NUM> would transition a downstream wall <NUM> of the minicore exit aperture <NUM> to the outer wall <NUM>.

Still referring to <FIG>, in various embodiments, the upstream facing side <NUM> of the ridge <NUM> is characterized by a height <NUM> (H). The height <NUM> may be further characterized in relation to a depth <NUM> (D) of the cooling passage <NUM>. According to the invention, the height <NUM> is configured within a range such that <NUM>. 05D < H < <NUM>. 75D; or, in various embodiments, the height <NUM> may be configured within a range such that <NUM>. 10D < H < <NUM>. 65D; or, in various embodiments, the height <NUM> may be configured within a range such that H ≈ <NUM>. Further, in various embodiments, the geometry of the ridge, as described above, facilitates incorporation of a first portion <NUM> of the thermal barrier coating <NUM> characterized by a first radius of curvature <NUM> (R<NUM>). The first portion <NUM> merges smoothly, through an inflection point (or an inflection line extending along a width of the minicore exit aperture), with a second portion <NUM> of the thermal barrier coating <NUM> that is characterized by a second radius of curvature <NUM> (R<NUM>) that is typically less than the first radius of curvature <NUM>. In various embodiments, for example, R<NUM> > 10R<NUM>; or, in various embodiments, R<NUM> > 5R<NUM>; or, in various embodiments, R<NUM> ≈ 2R<NUM>. The smooth transition between the first portion <NUM> and the second portion <NUM> facilitates smooth flow of the cooling fluid CF along the thermal barrier coating <NUM>, which extends a distance of between about 1D and about 2D into the minicore exit aperture <NUM> along the downstream wall <NUM> of the aperture. In various embodiments, an upstream wall <NUM> of the minicore exit aperture <NUM> is characterized by a third radius of curvature <NUM> (R<NUM>). The third radius of curvature <NUM> also facilitates smooth flow of the cooling fluid CF along the upstream wall <NUM> as the cooling fluid CF exits the minicore exit aperture <NUM>. In various embodiments, the third radius of curvature <NUM> is configured such that, for example, R<NUM> > 3D; or, in various embodiments, R<NUM> > 2D; or, in various embodiments, R<NUM> ≈ D. In addition, in order to configure the minicore exit aperture <NUM> as a diffuser, R<NUM> will generally be greater than R<NUM>, such that, for example, in various embodiments, R<NUM> > 5R<NUM>; or, in various embodiments, R<NUM> > 3R<NUM>; or, in various embodiments, R<NUM> ≈ 2R<NUM>.

The above disclosure provides improved manufacturability and reproducibility of casted components (e.g., airfoils of rotors or stators or guide vanes) having film cooling diffuser exits. The disclosure is particularly applicable to such components having ridge defects at a downstream end of the diffuser exit, where the ridge defects result from the casting process, and obviates the need to machine away or otherwise remove the defect prior to completing further steps in the fabrication process, such as, for example, applying a thermal barrier coating, thereby reducing manufacturing cost. The dimensional ranges described above also result in enhanced adherence of the film cooling fluid to the surface of the component (e.g., a reduction in the tendency for the cooling fluid to blow off the surface of the component).

Numbers, percentages, or other values stated herein are intended to include that value, and also other values that are about or approximately equal to the stated value, as would be appreciated by one of ordinary skill in the art encompassed by various embodiments of the present disclosure. A stated value should therefore be interpreted broadly enough to encompass values that are at least close enough to the stated value to perform a desired function or achieve a desired result. The stated values include at least the variation to be expected in a suitable industrial process, and may include values that are within <NUM>%, within <NUM>%, within <NUM>%, within <NUM>%, or within <NUM>% of a stated value. Additionally, the terms "substantially," "about" or "approximately" as used herein represent an amount close to the stated amount that still performs a desired function or achieves a desired result. For example, the term "substantially," "about" or "approximately" may refer to an amount that is within <NUM>% of, within <NUM>% of, within <NUM>% of, within <NUM>% of, and within <NUM>% of a stated amount or value.

Finally, it should be understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. Although various embodiments have been disclosed and described, one of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims.

Claim 1:
A component for a gas turbine engine (<NUM>), comprising:
a cooling passage (<NUM>);
an outer wall (<NUM>; <NUM>) separating a core flow path (C) from the cooling passage (<NUM>);
a minicore exit aperture in fluid communication with the cooling passage (<NUM>) and opening into the core flow path (C), the minicore exit aperture being characterized by a linear ridge (<NUM>; <NUM>) on a downstream end of the minicore exit aperture; and
a thermal barrier coating (<NUM>; <NUM>) covering the outer wall (<NUM>; <NUM>) and the linear ridge (<NUM>;<NUM>),
characterised in that the linear ridge (<NUM>; <NUM>) includes an upstream facing side (<NUM>) that is characterized by a height (<NUM>), the height (<NUM>) having a value within a range of between five one-hundredths and seventy-five one-hundredths of a depth (<NUM>) of the cooling passage (<NUM>).