WO2017142572A1 - Composant de turbine composite à matrice céramique ayant un substrat céramique renforcé par des fibres à gradient - Google Patents

Composant de turbine composite à matrice céramique ayant un substrat céramique renforcé par des fibres à gradient Download PDF

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Publication number
WO2017142572A1
WO2017142572A1 PCT/US2016/031607 US2016031607W WO2017142572A1 WO 2017142572 A1 WO2017142572 A1 WO 2017142572A1 US 2016031607 W US2016031607 W US 2016031607W WO 2017142572 A1 WO2017142572 A1 WO 2017142572A1
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WO
WIPO (PCT)
Prior art keywords
fibers
fiber
layer
tbc
ceramic
Prior art date
Application number
PCT/US2016/031607
Other languages
English (en)
Inventor
Ramesh Subramanian
Christian Xavier Campbell
Original Assignee
Siemens Aktiengesellschaft
Siemens Energy, Inc.
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority claimed from PCT/US2016/018224 external-priority patent/WO2016133990A1/fr
Application filed by Siemens Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Priority to EP16728764.8A priority Critical patent/EP3397840A1/fr
Priority to CN201680081909.1A priority patent/CN108699916A/zh
Priority to US16/076,922 priority patent/US20190048730A1/en
Publication of WO2017142572A1 publication Critical patent/WO2017142572A1/fr

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/284—Selection of ceramic materials
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • 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 partem anchors the TBC to the CMC structure. BACKGROUND
  • 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.
  • 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.
  • C degrees Celsius
  • SiC-SiC Silicon Carbide fiber-Silicon Carbide core
  • 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.
  • TBC application over CMC or superalloy substrates presents new and different thermal expansion mismatch and adhesion challenge.
  • CMC and TBC materials all have different thermal expansion properties.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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 interlay er adhesion.
  • engineered surface features 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 partem 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.
  • the CMC component covers an underlying substrate, such as a superalloy metallic substrate.
  • the CMC component is a sleeve over a metallic substrate.
  • the CMC component has no underlying metallic substrate, and provides its own internal structural support within the fiber-reinforced ceramic substrate.
  • a plurality of CMC components are joined together to form a larger, composite CMC component, such as a laminated turbine blade or vane.
  • 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.
  • 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.
  • TBC thermal barrier coat
  • 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.
  • CMC ceramic matrix composite
  • ESFs engineered surface features
  • a thermally sprayed or vapor deposited or solution/suspension plasma sprayed thermal barrier coat 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.
  • 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.
  • fiber strands or fiber strand loops project outwardly from the second layer, for increasing contact surface area with the TBC inner surface.
  • an intermediate layer of fibers is interposed between the inner and outer fiber layers.
  • 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.
  • 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 partem that defines voids therein.
  • 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 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.
  • 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.
  • the outer layer fibers have outwardly projecting fiber strands or fiber loops, for increasing contact surface area with the TBC inner surface.
  • 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 partem.
  • FIG. 1 is a partial axial, cross sectional view of a gas or combustion turbine engine, incorporating one or more CMC components constructed in accordance with exemplary embodiments of the invention
  • FIG. 2 is a cross sectional elevation view of a metal reinforced, CMC turbine blade component for a combustion turbine engine, in accordance with an exemplary embodiment of the invention
  • FIG. 3 is a cross sectional plan view of the turbine blade of FIG. 2;
  • FIG. 4 is a cross sectional elevation view of another embodiment of a CMC turbine blade component for a combustion turbine engine, which does not have internal metal reinforcement;
  • FIG. 5 is a cross sectional plan view of the turbine blade of FIG. 4;
  • FIG. 6 is a partial cross sectional elevation view through the side wall of the turbine blade of FIGs. 4 and 5, showing the graded fiber layers that are embedded within the solidified, fiber-reinforced ceramic substrate;
  • FIG. 7 is an elevational view of inner layer fibers of the blade side wall of
  • FIG. 8 is an elevational view of an alternate embodiment of the inner layer fibers of the blade side wall of FIG. 6, prior to their incorporation into the fiber- reinforced ceramic substrate, or application of a TBC;
  • FIG. 9 is a detailed elevational view of the inner layer fibers of FIG. 8;
  • FIG. 10 is an elevational view of intermediate layer fibers of the blade side wall of FIG. 6, prior to their incorporation into the fiber-reinforced ceramic substrate, or application of a TBC;
  • FIG. 11 is an elevational view of outer layer fibers of the blade side wall of FIG. 6, prior to their incorporation into the fiber-reinforced ceramic substrate, or application of a TBC;
  • FIG. 12 is a detailed elevational view of the outer layer fibers of FIG. 11 ;
  • FIG. 13 is a partial cross sectional elevation view through the side wall of the turbine blade of FIGs. 2 and 3, showing the blade's metallic support member, and the graded fibers that embedded within the fiber-reinforced ceramic substrate;
  • FIG. 14 is a partial cross sectional elevation view through the side wall of another embodiment of turbine blade, showing fiber strands projecting outwardly from the outer layer fibers, for anchoring the TBC; and
  • FIG. 15 is a partial cross sectional elevation view through the side wall of another embodiment of turbine blade, showing fiber loops projecting outwardly from the outer layer fibers, for anchoring the TBC.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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
  • ESFs engineered surface features
  • 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 0.5mm 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
  • 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.
  • ESFs engineered surface features
  • 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.
  • the CMC component or a plurality of CMC components are configured as insulative sleeves to cover the metallic member.
  • a plurality of such sleeves are stacked and laterally joined over a metallic member or other metallic substrate, prior to TBC application.
  • the CMC component is a unistructural, self-supporting blade, or vane for a gas turbine engine.
  • FIG. 1 shows a gas turbine engine 20, having a gas turbine casing 22, a multi- stage compressor section 24, a combustion section 26, a multi-stage turbine section 28 and a rotor 30.
  • One of a plurality of basket-type combustors 32 is coupled to a downstream transition 34 that directs combustion gasses from the combustor to the turbine section 28.
  • Atmospheric pressure intake air is drawn into the compressor section 24 generally in the direction of the flow arrows F along the axial length of the turbine engine 20.
  • the intake air is progressively pressurized in the compressor section 24 by rows rotating compressor blades 50 and directed by mating compressor vanes 52 to the combustion section 26, where it is mixed with fuel and ignited.
  • the ignited fuel/air mixture now under greater pressure and velocity than the original intake air, is directed through a transition 34 to the sequential vane 56 and blade 50 rows in the turbine section 28.
  • the engine's rotor 30 and shaft retains the plurality of rows of airfoil cross sectional shaped turbine blades 54.
  • Embodiments of the CMC components described herein are designed to operate in engine temperature environments of up to 1950 degrees Celsius.
  • the CMC components are insulative sleeves or coverings for metallic members, including metallic-substrate structural components, such as the subcomponents within the combustors 32, the transitions 34, the blades 54, or the vanes 56.
  • the CMC components of the invention are structurally self-supporting, without the need for metallic members or other supporting metallic substrates.
  • Exemplary self-supporting CMC components include compressor blades 50 or vanes 52 (which do not necessarily require the insulation of a TBC, internal subcomponents of combustors 32 or transitions 34).
  • entire turbine section 28 blades 54 or vane 56 airfoils are CMC structures; with outer surfaces of their fiber- reinforced ceramic substrates having surface textures, through use of graded fabrics, that mechanically interlock a relatively thick TBC layer of 0.5 to 2.0 mm.
  • FIGs. 2 and 3 A schematic cross section of an exemplary engine component, a turbine blade 60, is shown in FIGs. 2 and 3.
  • the turbine blade 60 has a blade root 62 for affixation to the turbine engine's rotor 30 and a distal blade tip 64.
  • the blade internal side wall 65 includes a metallic member, which is a metallic reinforcement rib 66.
  • the rib 66 is covered or otherwise circumscribed by a CMC, fiber-reinforced ceramic substrate 68.
  • Internal structural details of the intemal core within the turbine blade 60 are known, and for brevity are not shown or described in detail.
  • the ceramic substrate 68 includes graded ceramic fibers embedded therein, which are described in detail below.
  • a thermally sprayed or vapor deposited or solution/suspension plasma sprayed thermal barrier coat (“TBC”) 70 is applied over and coupled to the substrate 68 outer surface about the blade side wall pressure side 72 and its suction side 74, including the leading and trailing edges.
  • the turbine blade 80 embodiment of FIGs. 4 and 5 has a blade root 82, blade tip 84, and a self-supporting, CMC fiber-reinforced ceramic substrate 88, which form the blade airfoil side wall, including the internal side wall 85 and an outer portion onto which is applied a TBC 90.
  • the turbine blade 80 has no intemal metallic ribs or other support structure to carry loads imparted on the blade's pressure side 92 or suction side 94.
  • the self-supporting CMC, fiber-reinforced ceramic substrate 88 utilizes graded fibers for performing localized structural functions. Some fibers are primarily incorporated for structural strength within the CMC ceramic substrate, while others are primarily incorporated for anchoring the TBC layer 90.
  • the fibers forming the layered structure of the CMC fiber-reinforced ceramic substrate 88 are incorporated into sheets of fabric, which in some embodiments are stacked in layers or plies, and wrapped about the blade, i.e., outboard of the boundaries of the intemal side wall 85, during blade fabrication, analogous to a cigar wrapper.
  • fiber orientation within the fiber structures, which are embedded within fiber-reinforced ceramic substrate 88 incorporate one or more of woven fibers (two- or three-dimensional weaves, including knit weaves), braided fibers (including tows of braided fibers), and/or uniaxial fibers.
  • the description that follows with respect to FIG. 6 is also applicable to FIG. 13.
  • the schematic of FIG. 13, directed to the blade embodiment of FIGs. 2 and 3, incorporates the metal member, blade-reinforcing rib 66, but is otherwise identical to FIG. 6.
  • a plane view of each exemplary type of fabric sheet within the ceramic core 88 is shown in FIG. 6.
  • an inner layer of parallel fibers such as the woven fabric fibers shown therein, have a first weave partem 100 that defines a first weave density and cross sectional area, providing structural strength to the component ceramic substrate 88 of the blade component 80.
  • the primary function of the inner layer fibers is to provide structural support to the ceramic substrate 88 and the overall blade component 80.
  • the shown inner layer weave pattern has parallel vertical or axially oriented fibers 102 that are woven in bundles or tows 112, which enhance axial and torsional tensile strength of the blade component 80.
  • the inner layer weave pattern 100 also has horizontal fibers 104, which maintain parallel orientation of the vertical/axially-oriented fibers 102.
  • the vertical fibers 102 are tightly packed in the horizontal direction of FIG. 7, with no intentional spacing between them.
  • the inner layer fabric 100A has a tight, flat weave fiber pattern, with an equal distribution and density of vertical fibers 102 A and horizontal fibers 104 A.
  • Both the embodiments 100 and 100 A for the inner layer fabric weave have relatively flat surface profiles. As previously noted, such flat weave profiles do not promote good bonding with TBC overlays, leaving the TBC susceptible to separation from the fiber-reinforced ceramic substrate 88 during thermal expansion or from relatively poor oxide bonding between the ceramic substrate 88 and the TBC material 90.
  • the embodiments 100 and 100A of the inner layer fiber weave are less than optimal for use in the outermost layer of a fiber-reinforced ceramic substrate, such as the turbine blade ceramic substrate 88.
  • the embodiments described herein incorporate graded fibers in the fiber- reinforced ceramic substrate 88, which vary orientation locally within the substrate, and which are chosen for their structural and functional suitability.
  • the fiber-reinforced ceramic substrate 88 incorporates an intermediate layer of fibers, such as the woven fiber fabric 110, which are outboard of the inner woven fiber layer(s) 100.
  • the intermediate layer 110 has a weave pattem that defines a weave density and cross sectional area, less than those of the first weave pattem 100 do.
  • the intermediate layer comprises vertical fibers (one or more fibers alone or in bundles/tows) 112, with lateral spacing distance S2, and horizontal fibers 114 (one or more fibers alone or in bundles/tows), with vertical spacing SI .
  • Voids 116 are formed between the spaced fibers 112 and 114, facilitating bonding and structural integrity within the CMC ceramic core 88.
  • an outer layer of fibers, such as the fabric woven fibers 120 is oriented in the CMC component's fiber-reinforced ceramic substrate 88, outboard of the inner layer 100 and the intermediate layer 110.
  • scrim fabric which generally comprises an open fiber grid of non- woven, fused fibers, is used to form the outer layer fibers 120.
  • the outer layer fibers 120 have a weave pattem that defines a weave density and cross sectional area less than those of the inner layer weave pattem 100 or the intermediate layer weave pattem 110.
  • This outer layer 120 employs a textured surface, with a weave pattern of vertical fibers 122 (one or more fibers alone or in bundles/tows, or a non-woven, scrim-type fabric), with lateral spacing distance S4, and horizontal fibers 124 (one or more fibers alone or in bundles/tows, or in scrim fabric fused fibers), with vertical spacing S3.
  • the spacing S3 or S4 between fibers is between 0.1 mm to 8 mm Voids 126 are formed between the net-like array of spaced fibers 122 and 124, facilitating bonding and structural integrity within the CMC ceramic core 88 and the TBC layer 90.
  • the three-dimensional, textured weave also creates vertical gaps or voids 126 of radial distance or height H, between the fiber strands 122A, 122B, and 122C, in addition to the net-like voids 126 formed between fibers in the open spacing distances S3 and S4.
  • vertical voids in the outer fiber layer defined by spacing between individual fibers or fiber bundles of textured surface profile height H in the outer fiber layer 120 varies between 0.1 to 1.5 mm.
  • TBC thermally sprayed or vapor deposited or solution/suspension plasma sprayed thermal barrier coat
  • 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.
  • 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.
  • 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
  • TBC adhesion to the CMC component's fiber-reinforced ceramic substrate is enhanced by bonding between the TBC material and fibers that project from the outer fabric layer outer surface.
  • the component 130 CMC component's fiber-reinforced ceramic substrate 132 has an outer layer fiber pattern 134, with fiber strands 136 projecting from the outer layer. Cutting ceramic fiber strands in non-structural load bearing outer layer fabric 134 peripheral zones of the CMC component's ceramic substrate 132 does not impair structural integrity of the CMC component 130. The fiber strands 136 provide additional anchoring for the TBC layer 140.
  • optional engineered groove features (“EGFs”) 71, 91 are cut into the respective TBC 70, 90 outer surfaces, as described in the incorporated by reference priority International Application No. PCT/US 16/18224, filed February 17 2016, and entitled "CERAMIC MATRIX COMPOSITE TURBINE COMPONENT WITH ENGINEERED SURFACE FEATURES RETAINING A THERMAL BARRIER COAT".
  • the EGFs 71, 91 are cut in partem arrays, including pattern arrays that intersect engineered surface features ("ESFs”) of the CMC ceramic core, for enhanced spallation isolation.
  • the CMC ceramic substrate outer surfaces and embedded ceramic fibers are cut by milling arrays of dimple- or cylindrical-shaped ESFs into them.
  • Other profile ESFs are optionally formed by selectively varying the ceramic substrate's outer profiles, symmetrically or asymmetrically.
  • Ceramic matrix composite (“CMC”) component for a combustion turbine engine
  • Such components include the oxide fiber-oxide ceramic core CMC components 60, 80, 130 and 150 of FIGs. 6 and 13-15, Using any known technique, graded ceramic fibers are layed-up into a layered structure. Exemplary layered structures are layed-up by orienting ceramic fibers into symmetrical or asymmetrical patterns. In some embodiments the fibers are already incorporated into a two- or three-dimensional fabric weave, or various fabric bundles, or within non-woven scrim fabric, ready to be layed-up into the layered structure.
  • the fiber pattern is selectively varied to provide anisotropic structural properties, for example if the finished CMC component is to function as a self-supporting or partially self-supporting structural element, as opposed to a non-structural insulative cover over a metallic member or another substrate.
  • 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
  • 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.
  • the fabric layers within the layed-up layered structure can be varied to accommodate future cut
  • 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 (“ZrCV); (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 fibers e.g., yttrium aluminum garnet (“YAG) fibers commercially available under the trademarks NEXTEL® 440, NEXTEL® 610, and NEXTEL® 720
  • ZrCV zirconium oxide
  • 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®
  • SiC, SiCN non-oxide ceramic matrix
  • 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 pyro lysis, 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 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 FIGs. 6 and 13-15, have not structurally weakened those components.
  • TBC thermal barrier coat
  • 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.
  • a rough surface ceramic bond coat is applied over the fiber-reinforced ceramic substrate outer surface, including its outer layer fabric and any optional ESFs, by a known deposition process, further enhancing adhesion of the TBC layer to the ceramic substrate.
  • the bond coat material is alumina or YAG, to enable oxidation protection, in case of complete TBC spallation from the ceramic substrate outer surface.
  • 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 0.5mm 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.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne un composant composite à matrice céramique (« CMC » pour Ceramic Matrix Composite), tel qu'une aube de turbine (80) pour un moteur à turbine à combustion (20) qui comporte un substrat (88) céramique solidifié renforcé par des fibres. Le substrat (88) comporte une couche interne de fibres (100) pour améliorer la résistance structurelle du composant (80). Une couche externe (120) de fibres définit des vides (126) en son sein. Une couche barrière thermique (« TBC » pour Thermal Barrier Coat) (90) est appliquée sur la couche externe de fibres (120), et couplée à cette dernière, ce qui permet de remplir les vides (126). Les vides (126) donnent une aire de surface accrue et verrouillent mécaniquement la couche TBC (90), ce qui permet d'améliorer l'adhérence entre le substrat (88) céramique renforcé par des fibres et la couche TBC.
PCT/US2016/031607 2014-02-25 2016-05-10 Composant de turbine composite à matrice céramique ayant un substrat céramique renforcé par des fibres à gradient WO2017142572A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
EP16728764.8A EP3397840A1 (fr) 2014-02-25 2016-05-10 Composant de turbine composite à matrice céramique ayant un substrat céramique renforcé par des fibres à gradient
CN201680081909.1A CN108699916A (zh) 2014-02-25 2016-05-10 具有分级纤维增强陶瓷基底的陶瓷基复合材料涡轮机部件
US16/076,922 US20190048730A1 (en) 2014-02-25 2016-05-10 Ceramic matrix composite turbine component with graded fiber-reinforced ceramic substrate

Applications Claiming Priority (2)

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PCT/US2016/018224 WO2016133990A1 (fr) 2015-02-18 2016-02-17 Composant de turbine composite à matrice en céramique avec caractéristiques de surface techniques conservant un revêtement formant barrière thermique
USPCT/US2016/018224 2016-02-17

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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2018164693A1 (fr) * 2017-03-10 2018-09-13 Siemens Aktiengesellschaft Impression en trois dimensions de structures composites de fibres céramiques pour une meilleure adhérence de revêtement de barrière thermique
WO2019040079A1 (fr) * 2017-08-25 2019-02-28 Siemens Aktiengesellschaft Impression tridimensionnelle d'un composite de fibres céramiques pour former une couche abradable de turbine
WO2019203826A1 (fr) * 2018-04-19 2019-10-24 Siemens Aktiengesellschaft Aubes de turbine et procédé de formation d'une aube de turbine
WO2020018815A1 (fr) * 2018-07-18 2020-01-23 Poly6 Technologies, Inc. Articles et procédés de fabrication
US20200256200A1 (en) * 2019-02-08 2020-08-13 United Technologies Corporation Article with ceramic barrier coating and layer of networked ceramic nanofibers

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20030175116A1 (en) * 2001-11-14 2003-09-18 Snecma Moteurs Abradable coating for gas turbine walls
US20090017260A1 (en) * 2001-08-02 2009-01-15 Kulkarni Anand A Segmented thermal barrier coating
EP2275645A2 (fr) * 2009-07-17 2011-01-19 Rolls-Royce Corporation Composant de turbine à gaz comprenant des caractéristiques de réduction de la fatigue
WO2011085376A1 (fr) * 2010-01-11 2011-07-14 Rolls-Royce Corporation Éléments d'atténuation d'une contrainte thermique ou mécanique sur un revêtement anticorrosion protégeant de l'environnement

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20090017260A1 (en) * 2001-08-02 2009-01-15 Kulkarni Anand A Segmented thermal barrier coating
US20030175116A1 (en) * 2001-11-14 2003-09-18 Snecma Moteurs Abradable coating for gas turbine walls
EP2275645A2 (fr) * 2009-07-17 2011-01-19 Rolls-Royce Corporation Composant de turbine à gaz comprenant des caractéristiques de réduction de la fatigue
WO2011085376A1 (fr) * 2010-01-11 2011-07-14 Rolls-Royce Corporation Éléments d'atténuation d'une contrainte thermique ou mécanique sur un revêtement anticorrosion protégeant de l'environnement

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2018164693A1 (fr) * 2017-03-10 2018-09-13 Siemens Aktiengesellschaft Impression en trois dimensions de structures composites de fibres céramiques pour une meilleure adhérence de revêtement de barrière thermique
WO2019040079A1 (fr) * 2017-08-25 2019-02-28 Siemens Aktiengesellschaft Impression tridimensionnelle d'un composite de fibres céramiques pour former une couche abradable de turbine
WO2019203826A1 (fr) * 2018-04-19 2019-10-24 Siemens Aktiengesellschaft Aubes de turbine et procédé de formation d'une aube de turbine
WO2020018815A1 (fr) * 2018-07-18 2020-01-23 Poly6 Technologies, Inc. Articles et procédés de fabrication
US20200256200A1 (en) * 2019-02-08 2020-08-13 United Technologies Corporation Article with ceramic barrier coating and layer of networked ceramic nanofibers
US11591918B2 (en) * 2019-02-08 2023-02-28 Raytheon Technologies Corporation Article with ceramic barrier coating and layer of networked ceramic nanofibers

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