Patent Description:
Ceramic coatings are used for several purposes in modern gas turbine engines (broadly inclusive of turbofans, turbojets, turboprops, turboshafts, and industrial gas turbines). Key purposes include being barrier coatings (environmental (EBC) and/or thermal (TBC)) and abradable coatings (e.g., in sliding engagement with an abrading member). Exemplary abradable coatings are used on the gaspath-facing inner diameter (ID) surface of blade outer air seals (BOAS) in circumferential sliding engagement with blade tips (e.g., airfoil tips or tip shrouds).

Other locations for abradable coatings include rotor-to-stator interactions other than at blade tips. One example involves sealing between an inter-blade stage area of a rotor and inner diameter (ID) tips of stator airfoils or inner diameter sealing surfaces of ID shrouds of vanes.

Delamination or spalling of thermal barrier coatings from their underlying substrates is a significant problem. A principal driver of delamination is differential thermal expansion/contraction of the coating and the underlying substrate. This can be exacerbated by coating contamination such as calcium magnesium alumino-silicate (CMAS) attack, also known as molten sand attack. CMAS attack reduces the ability of the coating to accommodate differential thermal expansion.

<CIT>, "Segmented thermally insulating coating" (the `<NUM> patent), discloses combatting spalling of ceramic coatings via creation of faults to accommodate thermal deformations. To initiate the faults, the substrate is provided with an array of recesses along the region to be coated. The boundaries between recessed and unrecessed surface provide initiation sites for the faults. With typical plasma spray coatings, the initiation may be the creation of boundaries/gaps between regions of the as-applied coating.

<CIT>, "SEGMENTED CERAMIC COATING INTERLAYER" (the `<NUM> publication), discloses further use of a ceramic interlayer.

<CIT>, "EDGE TREATMENT FOR GAS TURBINE ENGINE COMPONENT" (the `<NUM> publication), discusses the particular problems of spallation at coating edges. A particular example is at the circumferential (circumferential end) edges of BOAS segments.

<CIT> discloses a prior art article according to the preamble of claim <NUM>.

One aspect of the disclosure provides a gas turbine engine component as set forth in claim <NUM>.

Another aspect of the disclosure provides a method as set forth in claim <NUM>.

BOAS inter-segment edges are an area of high temperature and resultant high coating spallation due to several factors: the edges are free and exposed to heat on two sides; blade passage causes disruption of film cooling; pressure fluctuations associated with blade passage may cause hot gasses to pump in and out of intersegment gaps; and the <NUM>° edge creates stress concentration. Various implementations disclosed below may addresses one or more of heat transfer, tolerance to thermally induced sintering shrinkage and the stress concentrations.

<FIG> schematically illustrates an example gas turbine engine <NUM> that includes a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM>, and a turbine section <NUM>. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section <NUM> drives air along a bypass flow path B while the compressor section <NUM> draws air in along a core flow path C where air is compressed and communicated to the combustor section <NUM>. In the combustor section <NUM>, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section <NUM> where energy is extracted and utilized to drive the fan section <NUM> and the compressor section <NUM>.

Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, 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 such as turbojets, turboprops, turboshafts, and industrial gas turbines (IGT).

The example 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 <NUM> at various locations may alternatively or additionally be provided.

The low speed spool <NUM> generally includes an inner shaft <NUM> that connects a fan <NUM> and a low pressure (or first) compressor section <NUM> to a low pressure (or first) turbine section <NUM>. The example engine is a geared turbofan where the inner shaft <NUM> drives the fan <NUM> through a speed change device, such as a geared architecture <NUM> (e.g., epicyclic transmission), to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The high speed spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure (or second) compressor section <NUM> and a high pressure (or second) turbine section <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine central longitudinal axis A.

The combustion section <NUM> comprises a combustor <NUM> between the high pressure compressor <NUM> and the high pressure turbine <NUM>. The example combustor is an annular combustor. Alternative combustors include can-type combustor arrays. In one example, the high pressure turbine <NUM> includes at least two stages to provide a double stage high pressure turbine <NUM>. In another example, the high pressure turbine <NUM> includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

A mid-turbine frame <NUM> of the engine static structure <NUM> is generally between the high pressure turbine <NUM> and the low pressure turbine <NUM>. The mid-turbine frame <NUM> further supports bearing systems <NUM> in the turbine section <NUM> as well as setting airflow entering the low pressure turbine <NUM>.

The core airflow C is compressed by the low pressure compressor <NUM>, then by the high pressure compressor <NUM>, mixed with fuel and ignited in the combustor <NUM> to produce high speed exhaust gases that are then expanded through the high pressure turbine <NUM> and low pressure turbine <NUM>. The mid-turbine frame <NUM> includes vanes <NUM>, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine <NUM>. Utilizing the vane <NUM> of the mid-turbine frame <NUM> as the inlet guide vane for low pressure turbine <NUM> decreases the length of the low pressure turbine <NUM> without increasing the axial length of the mid-turbine frame <NUM>. Reducing or eliminating the number of vanes in the low pressure turbine <NUM> shortens the axial length of the turbine section <NUM>. Thus, the compactness of the gas turbine engine <NUM> is increased and a higher power density may be achieved.

<FIG> illustrates a portion <NUM> of a gas turbine engine, such as the gas turbine engine <NUM> of <FIG>. In this exemplary embodiment, the portion <NUM> represents the high pressure turbine <NUM>. However, it should be understood that other portions of the gas turbine engine <NUM> could benefit from the teachings of this disclosure, including but not limited to, the compressor section <NUM>, the combustor section <NUM>, and the low pressure turbine <NUM>.

In this exemplary embodiment, a rotor disk <NUM> (only one shown, although multiple disks could be axially disposed within the portion <NUM>) is mounted to the outer shaft <NUM> and rotates as a unit with respect to the engine static structure <NUM>. The portion <NUM> includes alternating rows of rotating blades <NUM> (mounted to the rotor disk <NUM>) and vanes 70A and 70B of vane assemblies <NUM> that are also supported within an outer casing <NUM> of the engine static structure <NUM>.

Each blade <NUM> of the rotor disk <NUM> includes a blade tip 68T that is positioned at a radially outermost portion of the blades <NUM>. The blade tip 68T extends toward a blade outer air seal (BOAS) assembly <NUM>. The BOAS assembly <NUM> may find beneficial use in many industries including aerospace, industrial, electricity generation, naval propulsion, pumps for gas and oil transmission, aircraft propulsion, vehicle engines and stationery power plants.

The BOAS assembly <NUM> is disposed in an annulus radially between the outer casing <NUM> and the blade tip 68T. The BOAS assembly <NUM> generally includes a support structure <NUM> and a multitude of BOAS segments <NUM> (only one shown in <FIG>). The BOAS segments <NUM> may form a full ring hoop assembly that encircles associated blades <NUM> of a stage of the portion <NUM>. The support structure <NUM> is mounted radially inward from the outer casing <NUM> and includes forward and aft flanges 78A, 78B that mountably receive the BOAS segments <NUM>. The forward flange 78A and the aft flange 78B may be manufactured of a metallic alloy material and may be circumferentially segmented for the receipt of the BOAS segments <NUM>.

The support structure <NUM> may establish a cavity <NUM> that extends axially between the forward flange 78A and the aft flange 78B and radially between the outer casing <NUM> and the BOAS segment <NUM>. A secondary cooling airflow S may be communicated into the cavity <NUM> to provide a dedicated source of cooling airflow for cooling the BOAS segments <NUM>. The secondary cooling airflow S can be sourced from the high pressure compressor <NUM> or any other upstream portion of the gas turbine engine <NUM>.

<FIG> illustrate one exemplary embodiment of a BOAS segment <NUM> that may be incorporated into a gas turbine engine, such as the gas turbine engine <NUM>. The BOAS segment <NUM> is an example type of gas turbine engine component. Portions of the BOAS segment <NUM> includes areas that are exposed to high temperature and other relatively harsh conditions.

Other types of gas turbine engine components face similar harsh conditions. Such components may include vane and blade platforms, burner liner segments, fuel nozzle guides, bulkhead segments, combustors, etc. Although the examples of this disclosure are described with reference to the BOAS segment <NUM>, the other types of gas turbine engine components could also benefit from the teachings of this disclosure, particularly those components having ceramic coating such as thermal barrier coatings and abradable coatings.

The BOAS segment <NUM> may include a seal body <NUM> having a radially inner face <NUM> (or faces) that faces toward the blade tip 68T a radially outer face <NUM> (or faces) that faces toward the cavity <NUM> (See <FIG>). The radially inner face <NUM> (inboard or inner diameter (ID)) and the radially outer face <NUM> (outboard or outer diameter (OD)) circumferentially extend between a first mate face (circumferential end) <NUM> and a second mate face (circumferential end) <NUM> and axially extend between a leading edge face or end <NUM> and a trailing edge face or end <NUM>. A pair of fore-and-aft mounting lugs <NUM> extend outward from the outer face <NUM>.

The first mate face <NUM> meets the radially inner face <NUM> at an intersection <NUM>. There are similar intersections or junctions between the radial interface and the mate face <NUM>, leading edge face <NUM>, and trailing edge face <NUM>. However, the interaction of the blades make the intersection <NUM> particularly relevant. As is discussed below, the intersection includes a bevel rather than reflecting a right angle intersection of substrate surfaces. <FIG> shows a blade sweep direction <NUM>. Thus, there may be the highest contact forces as the blade tips first contact the inner face <NUM> at the intersection <NUM>. This may present some of the highest mechanical forces and shocks. Nevertheless, other locations are relevant. For example, near the leading end there are mechanical forces associated with the impingement of the gas flow including any particulate therein. At the intersection of the inner face with the second mate face <NUM>, there are kinetics associated with the releasing of the force from blade tip contact as the blade tip passes out of engagement with one segment and into engagement with the next segment.

<FIG> shows an intersection between two adjacent BOAS segments <NUM>. <FIG> shows a ceramic abradable coating system <NUM> atop a metallic substrate <NUM>. The exemplary substrate material is nickel-based superalloy. Although the exemplary coating is shown as a single ceramic layer, there may be multiple ceramic layers. Additionally, there may be a bondcoat <NUM> (<FIG>, e.g., an MCrAlY (e.g., NiCoCrAlY applied such as by physical vapor deposition (PVD) or thermal spray (e.g., high velocity air fuel (HVAF)), an aluminide (e.g. a platinum aluminide applied such as by plating platinum and spraying or packing aluminum followed by diffusion), or the like) atop the substrate as well as a thermally grown oxide, diffusion zone, or the like. Exemplary ceramic coatings are stabilized zirconias such as yttria stabilized zirconias (YSZ) and gadolinia stabilized zirconias (GSZ) or combinations.

The ID surface of the ceramic <NUM> provides the inner face <NUM> of the overall blade outer air seal. The substrate <NUM> has a corresponding ID face <NUM>. In general, the segment and the substrate may have corresponding portions with the corresponding portions being one in the same except along the coated areas. However, areas of the segment not covered by ceramic coating may be covered by some other coating such as for corrosion protection. Typically, these other coatings are much thinner than the ceramic coatings. Exemplary such other coatings include chromium conversion coatings.

<FIG> shows first and second circumferential end faces <NUM> and <NUM> of the substrate along the corresponding mate faces <NUM> and <NUM> of the overall segment. Cooling passages <NUM> may have outlets <NUM> (<FIG>) to these surfaces. The cooling passages may be fed from bypass air introduced through the radially outer face <NUM> by conventional means (not shown).

As one example of an interaction between adjacent segments, the substrates have a ship lap junction formed by a circumferential OD protrusion <NUM> from the second mate face <NUM> accommodated in an OD relief <NUM> extending from the first mate face of the adjacent segment. The interaction leaves a gap having a first (inner) radial segment <NUM>, a circumferential segment <NUM>, and a second (outer) radial segment <NUM>.

To help avoid spalling, the ID face <NUM> is provided with an array of recesses <NUM> such as in the `<NUM> patent. The recessing leaves intact protruding material <NUM> between recesses. Exemplary recesses are principally machined circular recesses (e.g., flat bottomed, such as by a plunge end milling) in a regular array (e.g., a square array or a hexagonal array). As in the `<NUM> patent, the boundaries <NUM> between recesses serve as initiation sites for faults <NUM> which help isolate and accommodate stresses. For non-intersecting recesses, the protruding material is continuous encircling the recesses (e.g., circular recesses have an on-center spacing greater than their diameter).

Unlike in the `<NUM> patent, the recessing extends to and is modified along the edges of the coated substrate region. <FIG> shows the substrate <NUM> as having an edge bevel or chamfer having a surface <NUM> and the coating having a corresponding bevel or chamfer having a surface <NUM>. Additionally, the recessing extends along the bevel with a recess base <NUM> (also see <FIG>) shown.

<FIG> shows an exemplary patterning of recesses along the underside. The recesses are shown as circular having diameter D<NUM> and in a hexagonal array leaving a thickness T<NUM> of intact substrate material at the location of minimum thickness. <FIG> further shows a recess depth or height of the protruding material as H<NUM> and an overall coating thickness measured from recess bases of H<NUM>. The substrate bevel on the intact material has a span S<NUM>. The corresponding recess base <NUM> may have a slightly larger span as is reflected, for example, in <FIG>. The coating bevel may be approximately the same size as S<NUM> or may be larger or smaller.

<FIG> shows one aspect of the hexagonal array when applied to a surface having a right angle corner between adjacent edges or ends. In this example, a main direction of the array is parallel to one edge while a secondary direction is parallel to the other. The recesses along the bevels may thus be different for the two edges.

<FIG> shows recesses <NUM> side-by-side along the edge (e.g., mate face <NUM>) aligned with the main direction and recesses <NUM> and <NUM> along the edge (e.g., leading end <NUM>) aligned with the secondary direction. The recesses <NUM> may comprise the combination of an intact portion <NUM> of a circular recess in the array and a portion <NUM> from bevel to edge. In one example using a plunge end mill, the end mill is plunged in to form the portion <NUM> then tilted to rotate its axis normal to the axis and then translated toward and past the edge to form the section <NUM>.

Along the secondary direction of the hexagonal pattern, smaller quill or bit may be used to machine the recesses <NUM>.

An exemplary manufacture process may modify any appropriate existing or yet-developed baseline process. The substrate may be formed by conventional techniques (e.g., investment casting over casting cores (if any) followed by deshelling/decoring and finish machining). The finish machining may leave reference surfaces for the subsequent machining of the recesses as discussed above. After recess machining, bondcoat (if any) may be applied via appropriate techniques including physical vapor deposition. The bondcoat (e.g., metallic), if any, may be applied/formed by conventional techniques.

The ceramic coating may then be applied via vapor deposition and/or spray (as noted above multiple layer coatings are possible). One area of examples involves a pure spray application (e.g., air plasma spray). Alternative spray techniques include high velocity oxy-fuel (HVOF) and flame spray. The spray application may leave excess coating material in order to guarantee sufficient ultimate thickness along the recesses. Thus, after coating application, the recesses may print through to the surface of the as-applied coating. The coating may then, however, be ground down to a smooth contour eliminating any valleys associated with the recesses but leaving the faults.

Returning to <FIG>, the bevel width or span S<NUM> is on the order of the size of the recesses (e.g., between the thickness T<NUM> and 2D<NUM>, more narrowly between T<NUM> and T<NUM>+D<NUM>. For typical aircraft engines, exemplary T<NUM> is <NUM> mil to <NUM> mil (<NUM> to <NUM>), more narrowly <NUM> mil to <NUM> mil (<NUM> to <NUM>). Exemplary D<NUM> is between <NUM> mil to <NUM> mil (<NUM> to <NUM>), more narrowly <NUM> mil to <NUM> mil (<NUM> to <NUM>). Exemplary H<NUM> <NUM> mil to <NUM> mil (<NUM> to <NUM>), more narrowly <NUM> mil to <NUM> mil (<NUM> to <NUM>) or <NUM> mil to <NUM> mil (<NUM> to <NUM>). For very thick coatings such as in large applications like industrial gas turbines, these could be magnified by a factor of three. An exemplary bevel angle θ is <NUM>° to <NUM>° off the ID face, more particularly, <NUM>° to <NUM>°.

The width S<NUM> is chosen to achieve coating thickness H<NUM> (along the bevel - <FIG>) to width S<NUM> ratio of <NUM>:<NUM> to <NUM>:<NUM> which relieves some of the stress from sintering shrinkage and promotes spallation resistance and surface temperature capability. In some less demanding applications, the ratio of thickness H<NUM> to width S<NUM> may be as small as <NUM>:<NUM>. To better illustrate features, <FIG> is not to scale/proportion. Also, although not shown, a further variation could extend the recesses radially farther along the unbeveled portions of the faces <NUM>, <NUM> (particularly where those unbeveled portions also were coated). Also, in the illustrated <FIG> configuration, H<NUM> and H<NUM>, measured normal to the actual surface (and not to the pre-bevel surface) are lower along the bevel than along unbeveled substrate. This is not required. They may be the same.

The pair of bevels brings TBC further radially outboard while beveling the square metallic edge which reduces heat load and base metal temperature and also reduces the temperature of the intersegment edge as the bevel helps to reduce hot gas impingement along the edges of the intersegment gap. The bevel also serves to reduce stress concentration by eliminating the square ground corners of TBC and base metal.

Further improvement in sintering tolerance is achieved by introducing linear features perpendicular to the chamfer in the form of the lateral boundaries of the recesses <NUM>, <NUM>. These boundaries initiate the faults <NUM> to provide ceramic segmentation and expansion joints in the ceramic layer. These faults act as an expansion joint to limit the transfer of stresses due to thermal expansion, sintering shrinkage, etc..

As noted above, depending on coating technique, these faults may be in the as-applied coating as non-crack regions of the coating structure where little or no strength has been developed in the planar directions. This is due to simultaneous deposition on the raised and recessed features which results in the lower coating regions building up alongside the raised regions where little to no bonding occurs. This low strength region acts as an expansion joint to achieve strain tolerance. In other examples (particularly physical vapor deposition (PVD) and suspension plasma spray (SPS)), the faults may form as cracks during thermal cycling after initial deposition.

As noted above, myriad variations are possible. Some variations include omitting the array of recesses <NUM> and only having recessing along the bevel. Other variations may extend an array of recesses only slightly inward along intact surface <NUM> but not along the entire footprint of the BOAS or other article.

Additional variations involve substrate and/or coating materials. Alternative substrates include ceramic matrix composites (CMC) such as SiC-SiC. These may favor alternative coatings such as those comprising (e.g., at least <NUM>% by weight) one or more of hafnia, hafnium silicate and yttrium silicate.

Claim 1:
A gas turbine engine component (<NUM>) comprising:
a body (<NUM>) having:
a first face (<NUM>); and
a first chamfer surface (<NUM>) extending from the first face (<NUM>); and
a ceramic coating (<NUM>) along the first face (<NUM>) and the first chamfer surface (<NUM>),
characterised in that the body (<NUM>) further comprises:
a plurality of first recesses (<NUM>, <NUM>) along the first chamfer surface (<NUM>) extending from the first face (<NUM>); and
linear features perpendicular to the first chamfer surface (<NUM>) in the form of lateral boundaries of the plurality of first recesses (<NUM>, <NUM>).