Enhanced adhesion thermal barrier coating

A gas turbine engine including a compressor section, a combustor section fluidly connected to the compressor section, a turbine section fluidly connected to the combustor section, and a plurality of gas path components exposed to a primary fluid flowpath through the compressor section, the combustor section and the turbine section. At least one of the gas path components includes an exterior facing surface, a lattice structure extending outward from the exterior facing surface, the lattice structure being integral to the exterior facing surface, and a thermal barrier coating adhered to at least a portion of the exterior facing surface and the lattice structure.

TECHNICAL FIELD

The present disclosure relates generally to thermal barrier coatings for gas powered turbine components, and more specifically to a system and process for enhancing adhesion between the thermal barrier coating and an underlying surface of a component for a gas powered turbine

BACKGROUND

In order to increase efficiencies of gas turbine engines, gas turbine engine manufacturers rely on extreme turbine inlet temperatures to provide a boost to the overall engine performance. In some modern gas turbine engine applications, the gas path temperatures within the turbine exceed the melting point of the constituent materials from which the underlying components of the gas path are constructed. To address the extreme heat, cooling systems are used to cool the gas path components in the turbine.

One exemplary mechanism for cooling turbine gas path components is a ceramic thermal barrier coating (TBC) adhered to an exterior surface of the gas path component. The presence of the thermal barrier coating significantly reduces the operating temperature of the component and allows for lower cooling flow requirements.

SUMMARY OF THE INVENTION

In one exemplary embodiment a gas turbine engine component includes an exterior facing surface, a lattice structure extending outward from the exterior facing surface, the lattice structure being integral to the exterior facing surface, and a thermal barrier coating adhered to at least a portion of the exterior facing surface and the lattice structure.

In another exemplary embodiment of the above described gas turbine engine component the lattice structure extends a first distance outward from the exterior facing surface and wherein the first distance is less than a thickness of the thermal barrier coating.

In another exemplary embodiment of any of the above described gas turbine engine components the lattice structure is at least partially an artifact of a manufacturing process.

In another exemplary embodiment of any of the above described gas turbine engine components the manufacturing process is an additive metal manufacturing process.

In another exemplary embodiment of any of the above described gas turbine engine components the exterior facing surface includes a plurality of through holes, and wherein the thermal barrier is characterized by not obstructing the through holes.

In another exemplary embodiment of any of the above described gas turbine engine components the exterior facing surface and the lattice structure have a combined surface roughness (Ra) in the range of 100-600.

In another exemplary embodiment of any of the above described gas turbine engine components a surface roughness of an exterior surface of the thermal barrier coating is less than 100 Ra.

In another exemplary embodiment of any of the above described gas turbine engine components the combined surface roughness (Ra) is greater than 180.

In another exemplary embodiment of any of the above described gas turbine engine components the lattice structure is disposed across less than 100% of the exterior surface.

In another exemplary embodiment of any of the above described gas turbine engine components the component is one of a blade outer air seal, a blade and a vane.

An exemplary method for manufacturing a gas path component for a gas powered turbine includes additively manufacturing a first component structure and a support structure, wherein the support structure is integral to the first component structure, partially removing the support structure via a finishing process such that an artifact of the support structure remains integral to the first component structure, and applying a thermal barrier coating to the first component structure, wherein the thermal barrier coating is adhered to the first component structure and the artifact of the support structure.

In another example of the above described method for manufacturing a gas path component for a gas powered turbine applying the thermal barrier includes applying a metallic bondcoat and applying a ceramic topcoat.

In another example of any of the above described exemplary methods for manufacturing a gas path component for a gas powered turbine additively manufacturing the first component structure and the support structure includes constructing a lattice structure on at least one external surface of the first component structure.

In another example of any of the above described exemplary methods for manufacturing a gas path component for a gas powered turbine the thickness of the thermal barrier coating is larger than a height of the artifact of the support structure normal to the surface of the first component structure on which the artifact of the support structure is disposed.

In another example of any of the above described exemplary methods for manufacturing a gas path component for a gas powered turbine partially removing the support structure via the finishing process comprises applying the finishing process until a combined surface roughness of the first component structure and the artifact of the support structure has a combined surface roughness (Ra) in the range of 100-600.

In another example of any of the above described exemplary methods for manufacturing a gas path component for a gas powered turbine partially artifact of the support structure is disposed across less than 100% of the exterior surface of the first component structure.

In one exemplary embodiment a gas turbine engine includes a compressor section, a combustor section fluidly connected to the compressor section, a turbine section fluidly connected to the combustor section, and a plurality of gas path components exposed to a primary fluid flowpath through the compressor section, the combustor section and the turbine section. At least one of the gas path components includes an exterior facing surface, a lattice structure extends outward from the exterior facing surface, the lattice structure being integral to the exterior facing surface, and a thermal barrier coating adhered to at least a portion of the exterior facing surface and the lattice structure.

In another exemplary embodiment of the above described gas turbine engine the at least one of the gas path components includes one of a blade outer air seal, a blade, and a vane.

In another exemplary embodiment of any of the above described gas turbine engines the lattice structure is at least partially an artifact of an additive manufacturing process.

In another exemplary embodiment of any of the above described gas turbine engines a combined surface roughness of the exterior facing surface and the lattice structure is a first magnitude and a combined surface roughness of the thermal barrier coating, the exterior facing surface, and the lattice structure has a second magnitude, the second magnitude being less than the first magnitude.

DETAILED DESCRIPTION OF AN EMBODIMENT

Turbine engine components, alternately referred to as gas path components, that are exposed to the combustion products formed in the combustor section26are exposed to extreme temperatures during operation of the gas powered turbine. In order to prevent damage to the gas path components, a ceramic thermal barrier coating is adhered to surfaces of the gas path components that are exposed to the gas path. The ceramic thermal barrier coating reduces the operating temperature of the component and allows lower cooling flow requirements for active cooling systems used to cool the gas path components.

In some exemplary systems, ceramic thermal barrier coatings are susceptible to breakage due to mechanical spallation. The breakage in such examples is accelerated in regions with high surface curvature and/or high heat load. By way of example, the leading edge of an airfoil is typically curved in such a manner.

With continued reference toFIG. 1,FIG. 2schematically illustrates a turbine inlet100immediately aft of a combustor110. Combustion products112are expelled from the combustor110along a gas path102. Disposed within the gas path102are a vane120and a blade130. The blade130is supported by a disk132according to standard turbine blade constructions. Radially outward of the blade130is a blade outer air seal140. Both the blade outer air seal140and the vane120are supported by an engine case structure150. Each of the surfaces of the vane120, blade130, and blade outer air seal140that are exposed to the combustion products112in the gas path102are at least partially coated with a thermal barrier coating.

With continued reference toFIG. 2,FIG. 3schematically illustrates the blade130ofFIG. 2isolated from the gas path structure. Similarly,FIG. 4illustrates a cross sectional view of the blade130ofFIG. 2along cross section lines A-A. The blade130includes a leading edge210and a trailing edge220. A root section230connects the blade130to the disk132(illustrated inFIG. 2). A thermal barrier coating240is applied to the leading edge210of the blade130. In alternative examples, the thermal barrier coating240can be applied to the entire blade130, instead of being limited to the leading edge210. Some example gas path components include holes disposed on the surface connecting an internal cavity to the exterior of the gas path component. In such examples, the thermal barrier coating does not cover or impede the through holes.

Due to the curvature of the leading edge210, adherence of the thermal barrier coating240to the exterior facing surface of the blade structure is reduced under extreme heat loads, and mechanical spallation can result. In order to increase the adherence of the thermal barrier coating240to the exterior facing surface of the blade structure, a lattice structure (illustrated inFIG. 5) is disposed on the leading edge210of the blade130. The lattice structure protrudes outward from the exterior facing surface of the blade130and is covered by the thermal barrier coating240.

With continued reference toFIGS. 2-4,FIG. 5illustrates the lattice structure400positioned on the exterior facing surface of the blade130.FIG. 6illustrates a cross sectional view of the lattice structure400ofFIG. 5along view lines C-C. The lattice structure400is formed of a series of interconnected walls410extending outward from an exterior facing surface420of the blade130. In the example lattice structure400, each of the walls410extends approximately normal to the exterior facing surface420on which the lattice structure400is disposed. In alternative examples, the lattice structure400can extend at an angle to the surface, with a component of the angle being normal to the exterior facing surface420. While illustrated inFIG. 5as a connected pattern of hexagons, practical implementations can utilize any interconnected shape, and the lattice structure400is not limited to the illustrated hexagonal shape.

A thermal barrier coating430is applied over the lattice structure400. The thermal barrier coating430in this example is applied by first applying a metallic bondcoat and then applying a ceramic topcoat. In alternative examples, alternative thermal barrier coating application techniques can be utilized to similar effect. A depth432of the thermal barrier coating430normal to the exterior facing surface420is less than a height412of the normal component of the lattice structure400. The surface area generated by the sides of the lattice structure400provides a greater surface for the thermal barrier coating to adhere to and reduces the occurrence of spallation.

In some examples, prior to application of the thermal barrier coating, the lattice structure400provides a surface roughness (Ra) of the exterior facing surface420in the range of 100-600 Ra. In other examples, the surface roughness provided by the lattice structure400is in the range of 180-600 Ra. As the thermal barrier coating430has a depth432greater than the height412of the lattice structure400, the resultant surface roughness of the component after application of the thermal barrier coating430is significantly lower than the surface roughness of the exterior facing surface420. The specific surface roughness of the thermal barrier coating430can be designed to a needed roughness according to known thermal barrier coating techniques. In some examples, the resultant surface roughness of the thermal barrier coating is less than 100 Ra.

With continued reference toFIGS. 2-6,FIG. 7illustrates a completed blade610constructed in an additive metal manufacturing device600.

Additive metal manufacturing typically utilizes lasers or electron beams to sinter particles in a 2D powder bed. Parts, such as the blade610, are made by successively sintering layers upwards to form the component. In order to create overhanging structures using the additive metal manufacturing process, support material and frameworks must be utilized to maintain the orientation of the part. When removed from the part, the support structures leave remnant structures that are attached to, and integral to, the constructed part. The remnant structures are alternately referred to as artifacts. In existing systems the artifacts are removed in a finishing process.

In the illustrated example, the blade610is constructed with a leading edge620approximately parallel to a base602of the additive metal manufacturing device600. Support structures630extend upward and contact/support the blade610as it is being constructed. As will be understood by one of skill in the art, the support structures630are constructed as a part of the additive metal manufacturing process.

In the illustrated example, the blade610and support structures630are constructed layer by layer beginning at the base and extending in a direction indicated by the arrow640. Each of the support structures630contacts only the leading edge region of the blade610. After the blade610is constructed via the additive metal manufacturing process, the support structures630are removed from the blade610via a finishing process. By halting the finishing process while a portion of the support structure630remains connected to the leading edge, the artifacts from the support structure630form the lattice structure described above. One of skill in the art, having the benefit of this disclosure, will understand methods by which to finish the blade610to a desired surface roughness such that an appropriately sized lattice structure remains.

In alternative examples, such as those where a lattice structure is desired across the entire surface of the blade610, the support structures630can extend across the entire blade surface. In yet further alternative structures, a lattice structure can be constructed along any desired exterior surface of the blade610during the additive metal manufacturing process independent of the support structure.

While described above with general reference to application of a thermal barrier coating to a leading edge of a turbine blade, one of skill in the art will understand that the above described lattice structure applied to an exterior facing surface of a component can be utilized to increase adhesion of a thermal barrier coating to the exterior facing surface of any component and any surface, and is not limited to a leading edge of a turbine blade.