Patent Application
20190048730
The invention relates to components for combustion turbine engines, with ceramic matrix composite (“CMC”) structures that are in turn insulated by a thermal barrier coating (“TBC”), and methods for making such components. More particularly, the invention relates to engine components for combustion turbines, with ceramic matrix composite (“CMC”) structures, having graded fiber-reinforced ceramic substrates. An inner layer fiber pattern provides structural support for the component and an outer layer fiber pattern anchors the TBC to the CMC structure.
CMC structures comprise a solidified ceramic substrate, in which are embedded ceramic fibers. The embedded ceramic fibers within the ceramic substrate of the CMC improve elongation rupture resistance, fracture toughness, thermal shock resistance, and dynamic load capabilities, compared to ceramic structures that do not incorporate the embedded fibers. The CMC embedded fiber orientation also facilitates selective anisotropic alteration of the component's structural properties. CMC structures are fabricated by laying-up or otherwise orienting ceramic fibers, also known as “rovings”, into fabrics, filament windings, tows, or braids. Fiber-reinforced ceramic substrate fabrication for CMCs is comparable to what is done to form fiber-reinforced polymer structural components for aircraft wings or boat hulls. Unless the ceramic fibers are pre-impregnated with a resin containing ceramic material, they are subsequently impregnated with ceramic material by such techniques as gas deposition, melt infiltration, preceramic polymer pyrolysis, chemical reactions, sintering, or electrophoretic deposition of ceramic powders, creating a solid ceramic structure with embedded, oriented ceramic fibers.
Ceramic matrix composite (“CMC”) structures are being incorporated into gas turbine engine components as insulation layers and/or structural elements of such components, such as insulating sleeves, vanes and turbine blades. These CMCs provide better oxidation resistance, and higher temperature capability, in the range of approximately 1150 degrees Celsius (“C”) for oxide based ceramic matrix composites, and up to around 1350 C for Silicon Carbide fiber-Silicon Carbide core (“SiC—SiC”) based ceramic matrix composites, whereas nickel or cobalt based superalloys are generally limited to approximately 950 to 1000 degrees Celsius under similar operating conditions within engines. While 1150 C (1350 C for SiC—SiC based CMCs) operating capability is an improvement over traditional superalloy temperature limits, mechanical strength (e.g., load bearing capacity) of CMCs is also limited by grain growth and reaction processes with the matrix and/or the environment at 1150 C/1350 C and higher. With desired combustion turbine engine firing temperatures as high as 1600-1700 C, the CMCs need additional thermal insulation protection interposed between themselves and the combustion gasses, to maintain their temperature below 1150 C/1350 C.
CMCs are receiving additional thermal insulation protection by application of overlayer(s) of thermal barrier coats or coatings (“TBCs”), as has been done in the past with superalloy components. However, TBC application over CMC or superalloy substrates presents new and different thermal expansion mismatch and adhesion challenge. During gas turbine engine operation superalloy, CMC and TBC materials all have different thermal expansion properties. In the case of TBC application over superalloy substrates, the superalloy material expands more than the overlying TBC material, which in extreme cases leads to crack formation in the TBC layer and its delamination from the superalloy surface. Along with thermal mismatch challenges, metallic substrate/TBC interfaces have adhesion challenges. While TBC material generally adheres well to a fresh metallic superalloy substrate, or in an overlying metallic bond coat (“BC”) substrate, the metals generate oxide surface layers, which subsequently degrade adhesion to the TBC at the respective layer interface.
TBC/metallic substrate interface integrity is maintained by use of the inventions in International Application No. PCT/US15/16318, entitled “TURBINE COMPONENT THERMAL BARRIER COATING WITH CRACK ISOLATING ENGINEERED GROOVE FEATURES”; and International Application No. PCT/US15/16331, entitled “TURBINE COMPONENT THERMAL BARRIER COATING WITH CRACK ISOLATING ENGINEERED SURFACE FEATURES”, both of which are in the priority chain of this application and which are incorporated by reference during National Phase prosecution in those jurisdictions that allow such incorporation. Some embodiments described in these priority applications incorporate engineered surface features (“ESFs”) on the substrate surface of the metallic superalloy substrate, or in an overlying metallic bond coat (“BC”), or a combination in both metallic surfaces. The ESFs at the metal surface/TBC layer interface mechanically anchor the TBC material, to inhibit delamination or at least confine delamination damage to boundaries defined by adjacent ESFs. Other embodiments in the priority applications incorporate engineered groove features (“EGFs”) on the TBC layer outer surface, to control surface crack propagation. Additional embodiments in those applications incorporate both ESFs and EGFs. Therefore, as the metal material is heated (forming surface oxides) and expands during engine operation, the lesser expanding TBC material is mechanically interlocked with the metal, despite degradation of interlayer adhesion.
Turning back to CMC/TBC thermal expansion mismatch and general interlayer adhesion challenges, relative layer expansion is opposite that experienced by superalloy/TBC components. The TBC material tends to expand more than underlying CMC material. As the TBC heats, it tends to lose adhesion with and delaminate from the CMC surface. Many CMC materials already contain oxides in the solidified ceramic core and in their embedded ceramic fibers, which adversely affect inter-layer adhesion at the CMC/TBC interface. In the case of SiC—SiC composites, the thermal barrier coatings can react with the underlying Silicon based matrix to form new chemical compounds, more brittle than the matrix or coating. Therefore, application of the TBC on the CMC surface of the component without subsequent delamination during engine operation is difficult. Depending upon the local macro roughness of the embedded ceramic fibers in the preform, and the infiltration characteristics of the ceramic material, which embed the preform into the solidified ceramic core, the adhesion of TBC coatings, is generally poorer than that of TBC coating on metallic substrates. TBC/CMC adhesion is particularly poor where the ceramic substrate's embedded fibers are oriented parallel to the component surface. TBC layer thickness is limited to that which will maintain adhesion to the CMC surface, despite its higher rate of thermal expansion. In other words, TBC layer thickness is kept below a threshold that accelerates the TBC/CMC thermal expansion delamination, within the already relatively limited bounds of TBC/CMC material adhesion capabilities. Unfortunately, limiting the TBC layer thickness undesirably limits its insulation properties. Generally, a thicker TBC layer offers more insulation protection to the underlying CMC substrate/layer than a thinner layer.
Exemplary embodiments described herein enhance TBC retention on CMC components in combustion turbine engines, by utilizing graded fiber or graded patterned fabric embedded in different zones within the CMC ceramic substrate. Inner fibers, in the more inwardly facing zone of the ceramic substrate provide greater structural strength of the component, than the outer fibers along the outer surface of the substrate, which interface with the TBC layer's inner surface. The outer fiber patterns have anchoring voids between fibers and/or fiber bundles for retention and anchoring of the TBC layer as the latter is applied to the ceramic core. In some embodiments, the outer fiber patterns have textured surfaces, including in some embodiments three-dimensional textured surfaces, for anchoring of the TBC layer within peaks and valley voids formed in the fabric weave. Other embodiments include fiber strands and/or fiber loops that project from the outer fabric weave pattern, for additional TBC layer anchoring. The outer fabric weaves voids and/or textured surface features mechanically interlock the CMC structure, and in particular, the fibers, to the TBC, and provide increased surface area and additional interlocking for interlayer adhesion. Optionally, as described in the incorporated by reference priority document, PCT/US16/18224, filed Feb. 17 2016, and entitled “CERAMIC MATRIX COMPOSITE TURBINE COMPONENT WITH ENGINEERED SURFACE FEATURES RETAINING A THERMAL BARRIER COAT”, engineered surface features (“ESFs”) are cut into the outer surface ceramic core and fibers of the preform. A thermally sprayed or vapor deposited or solution/suspension plasma sprayed TBC is applied over and coupled to the ceramic substrate outer surface and any cut ESFs. Increased adherence capabilities afforded by the outer fiber pattern voids and/or projections facilitate application of thicker TBC layers to the component, which increases insulation protection for the underlying CMC structure/layer. The increased adhesion surface area and added mechanical interlocking of the respective materials facilitates application of greater TBC layer thickness to the CMC substrate without risk of TBC delamination. The greater TBC layer thickness in turn provides more thermal insulation to the CMC structure, for higher potential engine operating temperatures and efficiency. In some embodiments, the CMC component covers an underlying substrate, such as a superalloy metallic substrate. In other embodiments, the CMC component is a sleeve over a metallic substrate. In other embodiments, the CMC component has no underlying metallic substrate, and provides its own internal structural support within the fiber-reinforced ceramic substrate. In additional embodiments, a plurality of CMC components are joined together to form a larger, composite CMC component, such as a laminated turbine blade or vane. In other embodiments, the CMC component is a unistructural, non-laminated, turbine blade, or vane.
The CMC component is made by laying-up ceramic fibers into a layered structure, having an inner layer, for structural support, and an outer layer, for TBC anchoring. If the layed-up fabric structure is not already pre-impregnated with ceramic material prior to laying them up, non-impregnated fibers are subsequently infiltrated with ceramic material, forming a solidified ceramic core. The TBC is then applied to the core outer surface. The outer fabric layer voids and projections assist in anchoring the TBC layer to the ceramic substrate's outer surface, in order to resist the aforementioned oxide layer and thermal expansion induced delamination challenges inherent in CMC/TBC components for gas turbine engines.
Exemplary embodiments feature a ceramic matrix composite (“CMC”) component for a combustion turbine engine, which has a solidified ceramic substrate, with ceramic fibers, embedded therein. The fiber-reinforced ceramic substrate has an inner layer of fibers, for enhancing structural strength of the component. The substrate also has an outer layer of fibers outboard of the inner layer, which defines voids therein. A thermally sprayed, or vapor deposited, or solution/suspension plasma sprayed thermal barrier coat (“TBC”) is applied over and coupled to the substrate's outer layer fibers, filling the voids. The voids provide increased surface area and mechanically interlock the TBC, improving adhesion between the fiber-reinforced ceramic substrate and the TBC. The TBC outer surface is suitable for combustion gas exposure when installed in an operating gas turbine engine.
Other exemplary embodiments feature a component for a combustion turbine engine, which component includes a metallic member for structural support. The metallic member is circumscribed by a ceramic matrix composite (“CMC”), such that the inner layer of fibers also circumscribe the metallic member. In some embodiments, the CMC layer also functions as a support substrate, with or without an additional metallic member. The CMC layer includes the solidified ceramic fiber-reinforced substrate, with a substrate inner surface that is shaped to conform to and abut the metallic member's surface profile. In some embodiments, engineered surface features (“ESFs”) are cut into the ceramic substrate's outer surface and its outer layer fibers. A thermally sprayed or vapor deposited or solution/suspension plasma sprayed thermal barrier coat (“TBC”), including a TBC inner surface, is applied over and coupled to the ceramic substrate's outer surface, anchored by voids in the outer layer fibers. In some embodiments, the outer layer fibers define a textured, surface profile, having height variation greater than the diameter of any single fiber, or bundle of fibers therein, for increasing contact surface area with the TBC inner surface. In some embodiments, fiber strands or fiber strand loops project outwardly from the second layer, for increasing contact surface area with the TBC inner surface. In other embodiments, an intermediate layer of fibers is interposed between the inner and outer fiber layers. In some embodiments, the intermediate layer has a pattern that defines a third density and cross sectional area less than those of the first layer do, and greater than those of the second layer.
Other exemplary embodiments feature methods for manufacturing a CMC component for a combustion turbine engine. A three-dimensional preform is fabricated with ceramic fibers. Ceramic fibers are layed-up into a layered structure, which includes an inner layer, for structural strength of the component. The layered structure also has an outer layer of fibers outboard of the inner layer, having a second weave pattern that defines voids therein. In some embodiments, the inner and outer layer fibers are pre-impregnated with ceramic material prior to being layed-up into the layered structure. If not already pre-impregnated with ceramic material, the layered structure is infiltrated with ceramic material, forming a solidified, fiber-reinforced ceramic substrate, which defines a substrate outer surface. A thermally sprayed, or vapor deposited, or solution/suspension plasma sprayed thermal barrier coat (“TBC”) is applied over and coupled to the substrate outer surface and its outer layer fibers, filling the voids. The voids provide increased surface area and mechanically interlock the TBC, improving adhesion between the ceramic substrate and the TBC. The TBC outer surface is suitable for combustion gas exposure when installed in an operating gas turbine engine. In some method embodiments, the outer layer fibers have a textured surface profile, having height variation greater than the diameter of any single fiber, or bundle of fibers therein, for increasing contact surface area with the TBC inner surface. In other embodiments, the outer layer fibers have outwardly projecting fiber strands or fiber loops, for increasing contact surface area with the TBC inner surface. In yet other embodiments, the fiber-reinforced ceramic substrate is fabricated with an intermediate layer of fibers interposed between the inner and outer fiber layers, having a third weave pattern that defines a third weave density and cross sectional area less than those of the inner layer weave pattern, and greater than those of the outer layer weave pattern.
The respective features of the exemplary embodiments of the invention that are described herein may be applied jointly or severally in any combination or sub-combination.
The exemplary embodiments are further described in the following detailed description in conjunction with the accompanying drawings, in which:
To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures. The figures are not drawn to scale.
Exemplary embodiments herein are utilized in combustion turbine engines. Embodiments of the CMC components of the invention are combined to form composite structures, such as turbine blades or vanes, which are structurally self-supporting. In other embodiments, the CMC components cover other structural elements, such as internal metallic (e.g., superalloy metal) members, including by way of example structural reinforcement ribs or other types of supports. In some embodiments, the ceramic matrix composite (“CMC”) components of the invention are utilized as insulative covers or sleeves for other structural components, such as metallic superalloy components or other types of metallic support members. In other embodiments, the CMC component is entirely structurally self-supporting, relying on internally embedded fibers to provide additional strength to its fiber-reinforced, ceramic substrate. Embodiments of the CMC components of the invention have a solidified, fiber-reinforced ceramic substrate, with ceramic fibers embedded therein. The fiber-reinforced ceramic substrate utilizes a graded fiber or graded patterned fabric embedded in different zones within the CMC substrate. Inner fibers in the more inwardly facing zone of the ceramic substrate have relatively higher fiber density and cross section, for greater structural support of the component, than the outer fibers along the outer surface of the core, which interface with the TBC layer's inner surface. The outer fiber patterns have voids between fibers and/or fiber bundles for retention and anchoring of the TBC layer as the latter is applied to the fiber-reinforced ceramic substrate. In some embodiments, the outer fiber patterns have textured surfaces, including in other embodiments textured three-dimensional surfaces, for anchoring of the TBC layer within peaks and valley voids, or fiber-spacing voids formed in the fabric pattern or weave. Other embodiments include fiber strands and/or fiber loops that project from the outer fabric pattern or weave (including by further example knitted fabric weaves), for additional TBC layer anchoring. The outer fabric voids and surface features mechanically interlock the CMC structure, and in particular, the fibers, to the TBC, and provide increased surface area and additional interlocking for interlayer adhesion. In some embodiments, engineered surface features (“ESFs”) are cut into an outer surface of the fiber-reinforced ceramic substrate. A thermally sprayed or vapor deposited or solution/suspension plasma sprayed TBC is applied over and coupled to the fiber-reinforced ceramic substrate's outer surface and any cut ESFs.
The outer fabric layer voids and surface features provide increased surface area, and mechanically interlock the TBC, improving adhesion between the ceramic fiber-reinforced ceramic substrate and the TBC. The mechanical interlocking and improved adhesion afforded by the voids and surface features within the outer fabric layer facilitate application of relatively thick TBC layers, from 0.5 mm to 2.0 mm. Because of the thick TBC application, embodiments of the CMC components of the invention are capable of operation in combustion environments up to 1950 degrees Celsius, with the thick TBC limiting the CMC ceramic core temperature to below 1150/1350 degrees Celsius.
In accordance with method embodiments of the invention, the CMC component is made by laying-up ceramic fibers into a layered structure. If the ceramic fibers are not already pre-impregnated with ceramic material prior to their laying-up, they are subsequently infiltrated with ceramic material, forming a solidified, fiber-reinforced ceramic substrate. In some embodiments, engineered surface features (“ESFs”) are cut into the ceramic substrate's outer surface and its outer layer fibers. The TBC is then applied to the ceramic substrate's outer surface and any ESFs. If the CMC component is structurally self-supporting, the ceramic substrate's inner fabric layer provides structural support to the component, such as a blade or vane of a gas turbine engine. If the CMC component is an insulative cover for another structural component, such as a metallic member, superalloy substrate, the component is dimensioned to cover, or otherwise circumscribe, the metallic member. In some applications, the CMC component or a plurality of CMC components are configured as insulative sleeves to cover the metallic member. In some embodiments, a plurality of such sleeves are stacked and laterally joined over a metallic member or other metallic substrate, prior to TBC application. In other embodiments, the CMC component is a unistructural, self-supporting blade, or vane for a gas turbine engine.
A schematic cross section of an exemplary engine component, a turbine blade 60, is shown in
The turbine blade 80 embodiment of
The graded fabric plies, which are incorporated into the layered structure of the fiber-reinforced ceramic substrate 88, are shown schematically in
Referring to
Referring to
As shown in
As shown in the embodiment of
A thermally sprayed or vapor deposited or solution/suspension plasma sprayed thermal barrier coat (“TBC”) 90 is applied over and coupled to the CMC ceramic substrate 88 outer surface and its outer fabric layer 120. The TBC 90 bonds and anchors to the outer fabric layer 120, with its relatively large surface area along the bonding zone, compared to a relatively flat planar bonding zone, which would otherwise be formed by the weave pattern of the inner layer fabric 110. Experience has shown that TBC tends to delaminate and spall from a flat CMC outer surface, especially if the reinforcing fibers, such as those of the inner fabric layer 110, are oriented parallel to the CMC ceramic substrate 88 outer surface. In embodiments herein, the voids or interstices 126, including the exemplary three-dimensional voids and interstices, skew orientation of the fibers 122 and 124 relative to the TBC layer 90, which creates abutting interfaces, rather than the parallel interfaces of the inner fabric layer 110. Optional engineered groove features (EGFs) 91 are cut into the TBC outer surface.
Referring to the component embodiments of
Referring to
Exemplary methods for manufacturing a ceramic matrix composite (“CMC”) component for a combustion turbine engine are now described. Such components include the oxide fiber-oxide ceramic core CMC components 60, 80, 130 and 150 of
The graded fiber layers in the CMC component are selected to vary locally structural strength, as well as to enhance impregnated ceramic slurry material or TBC anchoring capabilities. The layered fabric's surface texture (e.g., within a two- or three-dimensional weave pattern fabric or non-woven scrim fabric) can be selectively varied during its laying-up or prior to the lay-up by selecting fabrics with desired fiber patterns. In some embodiments, the layed-up fiber surface texture is varied through application of different scrim fabric fiber spacing and/or fiber thickness, or weave/tow patterns within woven fabrics. This allows selective alteration of fiber orientation and anisotropic structural strength in some layers or zones within the fiber-reinforced ceramic substrate, and for future bonding with an applied TBC in other fabric layers or zones within the ceramic substrate. For example, in some embodiments, the fabric layers within the layed-up layered structure can be varied to accommodate future cut ESF orientation between fiber bundles or outwardly jutting projections in the completed fiber-reinforced ceramic substrate.
In some embodiments, the fiber-reinforced ceramic substrate 88, within the CMC-composite turbine blade 80, is made from: (i) oxide ceramic fibers (e.g., yttrium aluminum garnet (“YAG”) fibers commercially available under the trademarks NEXTEL® 440, NEXTEL® 610, and NEXTEL® 720), or alternatively, zirconium oxide (“ZrO2”); (ii) glass or glassy fibers (e.g., commercially available under the trademarks NEXTEL® 312, Fiberglass, E-glass); or (iii) non-oxide ceramic fibers (silicon carbide (“SiC”), or alternatively, silicon carbon nitride (“SiCN”)). Oxide ceramic fiber composites are typically formed using oxide ceramic slurry, such as alumina, mullite, zirconia, or zirconia toughened alumina (“ZTA”). Glass fiber composites typically have a glassy matrix. Non-oxide fiber ceramics (typically SiC, commercially available under trademarks SYLRAMIC®, HI-NICALON®, TYRANO®) are formed using a non-oxide ceramic matrix (SiC, SiCN) from ceramic powders, ceramic precursors (silicon polyborosilazane), chemical vapor infiltration, or melt infiltrated processing.
In some embodiments, the fibers used to lay-up the layered structure that will be incorporated into the fiber-reinforced ceramic substrate 88 are pre-impregnated with ceramic material (“pre-preg” fiber or fabrics). After the pre-preg lay-up is completed, it is cured into the solidified and hardened fiber-reinforced ceramic substrate 88, which is in turn processed into the final CMC component, such as the turbine blade 80. If pre-preg fiber material is not utilized, it is layed-up into a layered structure, which is subsequently impregnated with ceramic material prior to curing, solidification and hardening into the fiber-reinforced ceramic substrate 88. Exemplary ceramic materials used to impregnate the layered structure, for subsequent solidification into the fiber-reinforced ceramic substrate 88, include alumina silicate, alumina zirconia, alumina, yttria stabilized zirconia, silicon, or silicon carbide polymer precursors. The post lay-up infiltration is performed, by any known technique, including gas deposition, melt infiltration, chemical vapor infiltration, slurry infiltration, preceramic polymer pyrolysis, chemical reactions, sintering, or electrophoretic deposition of ceramic powders, creating a solid, fiber-reinforced ceramic structure with embedded, graded ceramic fiber layers 100, 120, and in some embodiments 110.
Optional engineered surface features (“ESFs”) are cut into the outer surface of the fiber-reinforced ceramic substrate, and into its embedded fibers 120, with any known cutting technique, including mechanical machining, ablation by laser or electric discharge machining, grid blasting, or high pressure fluid. While general CMC fabrication generally disfavors cutting fibers within a preform, for fear of structural weakening, cutting fibers proximate the outer surface of the fiber-reinforced ceramic substrate, such as those incorporated within the CMC components 60, 80, 130, and 150 of
A known composition, thermally sprayed, or vapor deposited, or solution/suspension plasma sprayed thermal barrier coat (“TBC”) is applied over the fiber-reinforced ceramic substrate 88. Exemplary TBC compositions include single layers of 8-weight percent yttria stabilized zirconia (“8YSZ”), or 20-weight percent yttria stabilized zirconia (“20YSZ”). For pyrochlore containing thermal barrier coatings, an underlayer of 8YSZ is required to form a bilayer 8YSZ/59 weight percent gadolinium stabilized zirconia (8YSZ/59GZO) coating, or a bilayer 8YSZ/30-50 weight percent yttria stabilized zirconia (“30-50 YSZ”) coating, or combinations thereof. The TBC adheres to the outer surface of the ceramic substrate, including the outer layer fibers 120 and any optional ESFs. The outer layer fibers 120 and any optional ESFs increase surface area for TBC to ceramic substrate adhesion, and provide mechanical interlocking of the materials. Cut ceramic fiber ends along sides of the optional ESFs, as well as the fiber strands 136 of the CMC component 130 of
Increased ceramic substrate/TBC adhesion, attributable to increased adhesion surface area, mechanical interlocking, and exposed outer layer ceramic fiber/TBC adhesion facilitate application of thicker TBC layers in the range of 0.5 mm to 2.00 mm, which would otherwise potentially delaminate from a comparable flat surface TBC/ceramic substrate interface. Thicker TBC increases insulation protection to the underlying ceramic substrate and fibers of the CMC component, such as a blade or vane for a combustion turbine engine. Exemplary simulated turbine component structures, fabricated in accordance with embodiments described herein, withstand TBC outer layer exposure to 1950 degrees Celsius combustion temperatures, while maintaining the underlying fiber-reinforced ceramic substrate and its embedded fiber layers temperatures below 1150 degrees/1350 degrees Celsius. As previously discussed, exposure of the underlying fiber-reinforced ceramic substrate and its embedded fiber layers within CMC components to temperatures above 1150 C/1350 C, within a combustion turbine engine, thermally degrade those components.
Although various embodiments that incorporate the invention have been shown and described in detail herein, others can readily devise many other varied embodiments that still incorporate the claimed invention. The invention is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The invention is capable of other embodiments and of being practiced or of being carried out in various ways. In addition, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of “including,” “comprising,” or “having” and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items. Unless specified or limited otherwise, the terms “mounted”, “connected”, “supported”, and “coupled”, and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further, “connected” and “coupled” are not restricted to physical, mechanical, or electrical connections or couplings.
This application claims priority to International Application No. PCT/US16/18224, filed Feb. 17 2016, and entitled “CERAMIC MATRIX COMPOSITE TURBINE COMPONENT WITH ENGINEERED SURFACE FEATURES RETAINING A THERMAL BARRIER COAT”, which claims priority to and the benefit of International Application No. PCT/US15/16318, filed Feb. 18, 2015, and entitled “TURBINE COMPONENT THERMAL BARRIER COATING WITH CRACK ISOLATING ENGINEERED GROOVE FEATURES”, the entire contents of which are incorporated by reference herein.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US2016/031607 | 5/10/2016 | WO | 00 |
| Number | Date | Country | |
|---|---|---|---|
| Parent | PCT/US2016/018224 | Feb 2016 | US |
| Child | 16076922 | US | |
| Parent | PCT/US2015/016318 | Feb 2015 | US |
| Child | PCT/US2016/018224 | US |
