Patent Description:
A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.

More commonly, non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are being used for various components within gas turbine engines. For example, given an ability for CMC materials to withstand relatively extreme temperatures, there is particular interest in replacing components within the flow path of the combustion gases with CMC materials. More particularly, there is interest in replacing rotor blades of the turbine section of the gas turbine engine with blades formed of CMC materials.

CMC turbine rotor blades generally are formed from a plurality of plies of CMC material. The plies may be divided into segments, with each segment corresponding to a portion of the rotor blade. For example, one segment of plies may correspond to an airfoil portion of the blade, one segment of plies may correspond to a dovetail portion of the blade, and so forth for different portions of the turbine rotor blade. The segments of plies may be processed in an autoclave to compact and cure the plies to form the turbine rotor blade.

However, typical rotor blades have plies in three dimensions, e.g., plies in some segments of a blade have a first ply direction and plies in other segments of the blade have a second ply direction, e.g., normal to the first ply direction. Compaction of such a blade having plies in three dimensions can be difficult, as the plurality of plies of the blade are not oriented in the same ply direction.

Accordingly, a method for fabricating a gas turbine engine component utilizing multiple processing steps would be useful. Further, a method for forming a component of a gas turbine engine having plies in three dimensions would be beneficial. More particularly, a method for fabricating a turbine rotor blade of a gas turbine engine, the turbine rotor blade having CMC plies in three dimensions, would be particularly advantageous. <CIT> discloses a method for fabricating a ceramic matrix composite component of a gas turbine engine.

In claim <NUM>, a method for fabricating a ceramic matrix composite component of a gas turbine engine is provided.

As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. For example, "upstream" refers to the direction from which the fluid flows and "downstream" refers to the direction to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of <FIG>, the gas turbine engine is a high-bypass turbofan jet engine <NUM>, referred to herein as "turbofan engine <NUM>. " As shown in <FIG>, the turbofan engine <NUM> defines an axial direction A (extending parallel to a longitudinal centerline <NUM> provided for reference) and a radial direction R. In general, the turbofan <NUM> includes a fan section <NUM> and a core turbine engine <NUM> disposed downstream from the fan section <NUM>.

The exemplary core turbine engine <NUM> depicted generally includes a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. The outer casing <NUM> encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor <NUM> and a high pressure (HP) compressor <NUM>; a combustion section <NUM>; a turbine section including a high pressure (HP) turbine <NUM> and a low pressure (LP) turbine <NUM>; and a jet exhaust nozzle section <NUM>. A high pressure (HP) shaft or spool <NUM> drivingly connects the HP turbine <NUM> to the HP compressor <NUM>. A low pressure (LP) shaft or spool <NUM> drivingly connects the LP turbine <NUM> to the LP compressor <NUM>.

For the embodiment depicted, the fan section <NUM> includes a variable pitch fan <NUM> having a plurality of fan blades <NUM> coupled to a disk <NUM> in a spaced apart manner. As depicted, the fan blades <NUM> extend outwardly from disk <NUM> generally along the radial direction R. Each fan blade <NUM> is rotatable relative to the disk <NUM> about a pitch axis P by virtue of the fan blades <NUM> being operatively coupled to a suitable actuation member <NUM> configured to collectively vary the pitch of the fan blades <NUM> in unison. The fan blades <NUM>, disk <NUM>, and actuation member <NUM> are together rotatable about the longitudinal axis <NUM> by LP shaft <NUM> across a power gear box <NUM>. The power gear box <NUM> includes a plurality of gears for stepping down the rotational speed of the LP shaft <NUM> to a more efficient rotational fan speed.

Referring still to the exemplary embodiment of <FIG>, the disk <NUM> is covered by rotatable front nacelle <NUM> aerodynamically contoured to promote an airflow through the plurality of fan blades <NUM>. Additionally, the exemplary fan section <NUM> includes an annular fan casing or outer nacelle <NUM> that circumferentially surrounds the fan <NUM> and/or at least a portion of the core turbine engine <NUM>. It should be appreciated that the nacelle <NUM> may be configured to be supported relative to the core turbine engine <NUM> by a plurality of circumferentially-spaced outlet guide vanes <NUM>. Moreover, a downstream section <NUM> of the nacelle <NUM> may extend over an outer portion of the core turbine engine <NUM> so as to define a bypass airflow passage <NUM> therebetween.

During operation of the turbofan engine <NUM>, a volume of air <NUM> enters the turbofan <NUM> through an associated inlet <NUM> of the nacelle <NUM> and/or fan section <NUM>. As the volume of air <NUM> passes across the fan blades <NUM>, a first portion of the air <NUM> as indicated by arrows <NUM> is directed or routed into the bypass airflow passage <NUM> and a second portion of the air <NUM> as indicated by arrow <NUM> is directed or routed into the LP compressor <NUM>. The ratio between the first portion of air <NUM> and the second portion of air <NUM> is commonly known as a bypass ratio. The pressure of the second portion of air <NUM> is then increased as it is routed through the high pressure (HP) compressor <NUM> and into the combustion section <NUM>, where it is mixed with fuel and burned to provide combustion gases <NUM>.

The combustion gases <NUM> are routed through the HP turbine <NUM> where a portion of thermal and/or kinetic energy from the combustion gases <NUM> is extracted via sequential stages of HP turbine stator vanes <NUM> that are coupled to the outer casing <NUM> and HP turbine rotor blades <NUM> that are coupled to the HP shaft or spool <NUM>, thus causing the HP shaft or spool <NUM> to rotate, thereby supporting operation of the HP compressor <NUM>. The combustion gases <NUM> are then routed through the LP turbine <NUM> where a second portion of thermal and kinetic energy is extracted from the combustion gases <NUM> via sequential stages of LP turbine stator vanes <NUM> that are coupled to the outer casing <NUM> and LP turbine rotor blades <NUM> that are coupled to the LP shaft or spool <NUM>, thus causing the LP shaft or spool <NUM> to rotate, thereby supporting operation of the LP compressor <NUM> and/or rotation of the fan <NUM>.

The combustion gases <NUM> are subsequently routed through the jet exhaust nozzle section <NUM> of the core turbine engine <NUM> to provide propulsive thrust. Simultaneously, the pressure of the first portion of air <NUM> is substantially increased as the first portion of air <NUM> is routed through the bypass airflow passage <NUM> before it is exhausted from a fan nozzle exhaust section <NUM> of the turbofan <NUM>, also providing propulsive thrust. The HP turbine <NUM>, the LP turbine <NUM>, and the jet exhaust nozzle section <NUM> at least partially define a hot gas path <NUM> for routing the combustion gases <NUM> through the core turbine engine <NUM>.

Referring now to <FIG>, an exemplary LP turbine rotor blade <NUM> is illustrated, the blade <NUM> having a forward side <NUM> and an aft side <NUM> and including an airfoil <NUM>, a dovetail <NUM>, a platform <NUM>, opposing angel wings <NUM>, and opposing flowpath portions <NUM> (only one angel wing <NUM> and flowpath portion <NUM> is shown in <FIG>). More particularly, dovetail <NUM> extends radially inwardly from the substantially planar platform <NUM>, which defines the radially inner boundary of the hot gases of combustion flowing through the LP turbine <NUM> of the turbofan engine <NUM>. Turbine rotor blade <NUM> also includes airfoil <NUM> extending radially outwardly from platform <NUM>. Additionally, blade <NUM> includes angel wings <NUM> configured to provide radial sealing between the rotating components coupled to the rotor disk (not shown), e.g., turbine rotor blade <NUM>, and the stationary components (not shown) disposed forward and aft of such rotating components so as to prevent hot gas ingestion within the wheel space (not shown) adjacent to the rotor disk. Further, opposing flowpath portions <NUM> extend outward from platform <NUM>, e.g., one flowpath portion <NUM> may be positioned on the forward side <NUM> of blade <NUM>, extending generally forward from platform <NUM>, and one flowpath portion <NUM> may be positioned on the aft side <NUM>, extending generally aft from platform <NUM>. As shown in <FIG>, angel wings <NUM> may extend radially inward from the opposing flowpath portions <NUM> such that angel wings <NUM> are opposing portions of blade <NUM>. Accordingly, <FIG> illustrates rotor blade <NUM> consists of various three dimensional portions such that blade <NUM> is defined in three dimensions and has a three-dimensional shape.

For the embodiment depicted, turbine rotor blade <NUM> is comprised of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Exemplary CMC materials utilized for such rotor blades <NUM> may include silicon carbide, silicon, silica, or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron's SCS-<NUM>), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and ATK-COI's SYLRAMIC®), alumina silicates (e.g., Nextel's <NUM> and <NUM>), and chopped whiskers and fibers (e.g., Nextel's <NUM> and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si, A1, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). As further examples, blade <NUM> may be formed from a CMC material such as silicon carbide (SiC) or carbon fiber cloth. In some embodiments, each portion of rotor blade <NUM>, i.e., airfoil <NUM>, dovetail <NUM>, platform <NUM>, angel wings <NUM>, and flowpath portions <NUM>, may be made from a CMC material. In other embodiments, some portions of blade <NUM> may be made from a CMC material and other portions of blade <NUM> may be made from a different material, e.g., a metal, metal alloy, or the like.

Referring now to the schematic illustration of <FIG>, in an exemplary embodiment, rotor blade <NUM> comprises an airfoil portion <NUM> fabricated from a plurality of plies <NUM> of a CMC material and a dovetail portion <NUM> fabricated from a plurality of plies <NUM> of a CMC material. In one exemplary embodiment, the plurality of dovetail plies <NUM> may comprise about <NUM> plies, but other numbers of dovetail plies <NUM> may be used as well. As shown, the plurality of airfoil plies <NUM> and dovetail plies <NUM> extend generally along the radial direction R. To fabricate blade <NUM>, the plurality of plies <NUM>, <NUM> for forming airfoil <NUM> and dovetail <NUM> may be laid up in a first layup tool <NUM> (<FIG>). Then, referring to the schematic illustration of <FIG>, a plurality of plies <NUM> of a CMC material for fabricating the platform portion <NUM> of rotor blade <NUM> may be laid up in first layup tool <NUM>. The plurality of platform plies <NUM> also extends along the radial direction R.

After laying up in first layup tool <NUM> plies <NUM>, <NUM>, <NUM> corresponding to airfoil <NUM>, dovetail <NUM>, and platform <NUM>, the plies are ready for processing, e.g., compaction and curing in an autoclave. In alternative embodiments, airfoil <NUM>, dovetail <NUM>, and platform <NUM> may be molded as shown in <FIG> to form a molded assembly AM, e.g., a preform assembly comprising airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM>, and then molded assembly AM may be laid up in first layup tool <NUM> for processing in an autoclave. In an exemplary processing cycle, three caul sheets may be used - e.g., one on airfoil <NUM>, one on the suction side of dovetail <NUM>, and one on the pressure side of dovetail <NUM> - and a standard autoclave cycle may be used to compact and cure the airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> or plies <NUM>, <NUM>, <NUM> of rotor blade <NUM>. In such a cycle, the airfoil, dovetail, and platform portions or plies are compacted generally along a first direction D<NUM>, e.g., generally along the pressure and suction sides of airfoil, dovetail, and/or platform portions <NUM>, <NUM>, <NUM>. In another exemplary cycle, first layup tool <NUM> includes has pressure side tooling including a caul sheet or a metal tool, and the pressure side tooling provides the compact force during the processing of airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM>.

Other processing cycles, e.g., utilizing a different number and/or configuration of caul sheets and the like, other known methods or techniques for compacting and/or curing CMC plies, or other configurations of first layup tool <NUM>, may be used as well. As an example, first layup tool <NUM> could be configured such that the plies are laid up on a pressure side tool and compacted from the suction side or that the plies are compacted from both the pressure side and the suction side. As a further example, the plies may be processed using a melt infiltration process, a chemical vapor infiltration process, use of a matrix of pre-ceramic polymer fired to obtain a ceramic matrix, or any combinations of these or other known processes.

As illustrated schematically in <FIG>, after airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> have been processed, these portions of rotor blade <NUM> form a first assembly A<NUM>. In one embodiment, pre-formed angel wings <NUM> and flowpath portions <NUM> may be laid up with first assembly A<NUM> in a second layup tool <NUM> (<FIG>); then, the angel wing <NUM> and flowpath portion <NUM> preforms may be processed, e.g., in an autoclave, with first assembly A<NUM> to form a second assembly A<NUM>. In another embodiment, angel wing plies <NUM> and flowpath plies <NUM> may be laid up with first assembly A<NUM> in second layup tool <NUM>, and then angel wing plies <NUM> and flowpath plies <NUM> may be processed, e.g., in an autoclave, with first assembly A<NUM> to form second assembly A<NUM>. In exemplary embodiments, a standard autoclave cycle may be used to compact and cure first assembly A<NUM> and angel wing and flowpath preforms or plies. In such a standard cycle, the angel wing and flowpath preform portions <NUM>, <NUM> or plies <NUM>, <NUM> are compacted generally along a second direction D<NUM> and may also be compacted along a third direction D<NUM>, as shown in <FIG>. That is, the compaction force may be applied along second direction D<NUM> as well as third direction D<NUM>. Second direction D<NUM> is orthogonal, i.e., perpendicular, to third direction D<NUM>, and each of second direction D<NUM> and third direction D<NUM> are orthogonal to first direction D<NUM>. After processing, i.e., following the second processing cycle to bond angel wings <NUM> and flowpath portions <NUM> to the airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> of blade <NUM>, second assembly A<NUM> may be machine finished to form a finished turbine rotor blade <NUM>.

As shown in <FIG>, first assembly A<NUM> may define a first plane that includes the radial direction R and a blade direction B. It should be understood that airfoil, dovetail, and platform plies <NUM>, <NUM>, <NUM> extend parallel to the first plane. Further, as depicted schematically in <FIG>, angel wing plies <NUM> and/or flowpath plies <NUM> may define a second plane including the radial direction R and a flow direction F, and each ply <NUM>, <NUM> generally may extend within or parallel to the second plane. Second plane may be substantially normal to the first plane defined by first assembly A<NUM>. Thus, rotor blade <NUM> may be formed from plies in three dimensions, i.e., airfoil, dovetail, and platform plies <NUM>, <NUM>, <NUM> extending within or parallel to the first plane and angel wing and flowpath plies <NUM>, <NUM> extending within or parallel to the second plane, normal to the first plane.

By processing blade <NUM> in at least two steps, compaction of the plies can be improved. For example, processing airfoil, dovetail, and platform plies <NUM>, <NUM>, <NUM> in first layup tool <NUM> to form first assembly A<NUM> may allow optimal or improved compaction of these plies within or parallel to the first plane. Then, processing first assembly A<NUM> with angel wing plies <NUM> and flowpath plies <NUM> may allow optimal or improved compaction of these plies within or parallel to the second plane, as well as blade <NUM> overall. As one particular example, because the majority of the plies forming blade <NUM> are the airfoil and dovetail plies <NUM>, <NUM>, processing these portions of blade <NUM> first can function as an intermediate debulking step, such that the overall processing time and required compaction can be reduced. Further, separating the processing of airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> from the processing of angel wing and flowpath portions <NUM>, <NUM> can simplify the tooling required for processing blade <NUM>. That is, the tooling required to process blade <NUM> in one cycle can be more complicated than the tooling used to process blade <NUM> in two or more steps, e.g., first layup tool <NUM> and second layup tool <NUM>.

Additionally or alternatively, because the airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> are processed before the angel wings <NUM> and flowpath portions <NUM>, the airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> are stiffer during the second processing cycle than they would be if the entire blade was processed in a single cycle. That is, the airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> are stiffer when angel wings <NUM> and flowpath portions <NUM> are bonded to the other portions of blade <NUM> through application of a compaction force normal to the airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM>. Thus, the airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> are less susceptible to undesirable deformation during the second processing cycle.

<FIG> illustrate the stages of fabrication of an exemplary rotor blade <NUM>. <FIG> illustrates molded assembly AM, comprising airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> of blade <NUM>. <FIG> depicts first assembly A<NUM>, i.e., molded assembly AM post-processing. <FIG> illustrates second assembly A<NUM>, comprising airfoil, dovetail, platform, angel wing, and flowpath portions <NUM>, <NUM>, <NUM>, <NUM>, <NUM> of blade <NUM> following a second processing cycle. <FIG> depicts a finished blade <NUM>, i.e., blade <NUM> post-machining.

<FIG> illustrates an exemplary method <NUM> for fabricating a CMC component of a gas turbine engine, such as LP turbine rotor blades <NUM>. At step <NUM>, a material, such as a CMC material as described above, is molded to form a first portion of the gas turbine engine component. In embodiments in which the gas turbine engine component formed using method <NUM> is a turbine rotor blade such as blade <NUM>, the first portion may be molded assembly AM, comprising airfoil, dovetail, and platform portions <NUM>, <NUM>, <NUM> as described above. An exemplary molded assembly AM is shown in <FIG>. In other embodiments, the first portion may be any portion of the gas turbine engine component, but typically, the first portion is only a portion of the gas turbine engine component rather than the entire component.

At step <NUM>, the first portion is prepared for processing. For example, the first portion may be positioned on a first layup tool to prepare the first portion for processing. Continuing with the above example, where the first portion is molded assembly AM, in one embodiment the first portion may be laid up on first layup tool <NUM> to prepare the first portion for processing. The first layup tool, e.g., first layup tool <NUM>, may have any appropriate shape and configuration for supporting the first portion and/or aiding in the processing of the first portion. Other techniques or methods also may be used to prepare the first portion for processing.

After the first portion is prepared for processing, as shown at <NUM>, the first portion is processed to form a first assembly. As an example, the first portion may be processed to form the first assembly by curing in an autoclave. Alternatively or additionally, the first portion may be compacted in an autoclave as part of processing the first portion to form the first assembly. Continuing with the foregoing example, as previously described, molded assembly AM may be processed using a standard cycle in an autoclave, including compaction along the first direction D<NUM> and curing at an appropriate temperature and pressure. Processing molded assembly AM forms first assembly A1, comprising airfoil <NUM>, dovetail <NUM>, and platform <NUM> of turbine rotor blade <NUM>. An exemplary first assembly A<NUM> is illustrated in <FIG>.

At step <NUM>, the first assembly and a second portion of the gas turbine engine component are prepared for processing. In one embodiment, the first assembly and the second portion may be positioned on a second layup tool to prepare the first assembly and the second portion for processing. In the above example, where the first assembly is first assembly A<NUM>, the second portion may be angel wings <NUM> and flowpath portions <NUM>. In such embodiments, positioning the first assembly and second portion on a second layup tool may comprise laying up first assembly A<NUM> and pre-formed angel wings <NUM> and flowpath portions <NUM> on second layup tool <NUM>. Alternatively, preparing the first assembly and second portion for processing may comprise laying up first assembly A<NUM>, angel wing plies <NUM>, and flowpath plies <NUM> on second layup tool <NUM>. The second layup tool may have any appropriate shape and configuration for supporting the first assembly and second portion and/or aiding in the processing of the first assembly and second portion.

In some embodiments, first assembly A<NUM> may define a first plane and the second portion, e.g., angel wing <NUM> or flowpath <NUM> preforms or plies <NUM>, <NUM>, may define a second plane. Preparing first assembly A<NUM> and the second portion for processing may include positioning the second portion adjacent the first assembly A<NUM> on a second layup tool such that the second plane extends perpendicular to the first plane. The positions of first assembly A<NUM> and the second portion are depicted in <FIG> and <FIG>. Other techniques or methods also may be used to prepare the first assembly and the second portion for processing.

After being prepared for processing, as shown at <NUM>, the first assembly and second portion are processed to join or bond the first assembly and second portion and thereby form a second assembly. For example, the first assembly and the second portion may be cured in an autoclave to form the second assembly. Alternatively or additionally, the first assembly and second portion may be compacted in an autoclave as part of processing to form the second assembly. In one embodiment, the first assembly and second portion may be processed in an autoclave using a standard cycle, including applying a compaction force to the second portion along a second direction D<NUM> and/or a third direction D<NUM> and curing the assembly at an appropriate temperature and pressure. Continuing with the above example, the first direction D<NUM>, second direction D<NUM>, and third direction D<NUM> may be orthogonal to each other, i.e., the first direction D<NUM> may be perpendicular to the second direction D<NUM>, the second direction D<NUM> may be perpendicular to the third direction D<NUM>, and the third direction D<NUM> may be perpendicular to the first direction D<NUM>. The second portion, which may be angel wing <NUM> and flowpath portion <NUM> preforms or angel wing plies <NUM> and flowpath plies <NUM>, such that processing first assembly A<NUM> and the second portion forms second assembly A<NUM>, comprising airfoil <NUM>, dovetail <NUM>, platform <NUM>, angel wings <NUM>, and flowpath portions <NUM> of turbine rotor blade <NUM> as previously described. That is, the pre-formed angel wings <NUM> and flowpath portions <NUM>, or angel wing plies <NUM> and flowpath plies <NUM>, may be bonded to first assembly A<NUM> to form second assembly A<NUM>. An exemplary second assembly A<NUM> is shown in <FIG>.

At step <NUM>, the second assembly may be machined to produce the finished gas turbine engine component. Continuing with the foregoing example, second assembly A<NUM> may be machined to form a finished turbine rotor blade <NUM>. An exemplary machine finished turbine rotor blade <NUM> is depicted in <FIG>.

While described above with respect to fabricating LP turbine rotor blade <NUM>, it should be readily understood that method <NUM> also may be used to fabricate other gas turbine engine components. As one example, method <NUM> may be used to form HP turbine rotor blades <NUM>, as well as other components of turbofan engine <NUM>. Other gas turbine engine components may comprise CMC plies extending within or parallel more than two planes or requiring compaction along a plurality of directions. Accordingly, the layup tools or other tooling used to process the component may have an appropriate configuration for supporting the component during processing, and method <NUM> may be adjusted to have any appropriate or desirable number of processing steps.

<FIG> illustrates another exemplary method <NUM> for fabricating a component of a gas turbine engine, such as LP turbine rotor blades <NUM>. At step <NUM>, a first plurality of plies is prepared for processing. The first plurality of plies may comprise plies of a CMC material, such as the CMC materials described above. Further, as previously described, the first plurality of plies may comprise airfoil plies <NUM> and dovetail plies <NUM>, as well as platform plies <NUM>, for forming turbine rotor blade <NUM>. In an exemplary embodiment, the first plurality of plies are positioned on a first layup tool to prepare the first plurality of plies for processing. For example, airfoil plies <NUM> and dovetail plies <NUM> may be laid up on first layup tool <NUM>, and then platform plies <NUM> may be laid up on first layup tool <NUM>. As previously described, the first layup tool, e.g., first layup tool <NUM>, may have any appropriate shape and configuration for supporting the first plurality of plies and/or aiding in the processing of the first plurality of plies. Further, in other embodiments, the first plurality of plies may be plies for forming another component of a gas turbine engine. Other ways of laying up the plies and otherwise preparing the first plurality of plies for processing may be used as well.

At step <NUM>, the first plurality of plies is processed to form a first assembly. As an example, the first plurality of plies may be cured in an autoclave to form the first assembly. Alternatively or additionally, the first plurality of plies may be compacted in an autoclave as part of processing the first plurality of plies to form the first assembly. Continuing with the above example, where the first plurality of plies includes airfoil plies <NUM>, dovetail plies <NUM>, and platform plies <NUM> of turbine rotor blade <NUM>, the plies may be processed using the previously described standard autoclave cycle, including applying a compaction force along the first direction D<NUM> and curing at an appropriate temperature and pressure. Processing plies <NUM>, <NUM>, <NUM> forms first assembly A<NUM>, comprising airfoil <NUM>, dovetail <NUM>, and platform <NUM> of blade <NUM>. An exemplary first assembly A<NUM> is illustrated in <FIG>.

At step <NUM>, the first assembly and a second plurality of plies are prepared for processing. In one embodiment, the first assembly and the second plurality of plies may be positioned on a second layup tool to prepare the first assembly and the second plurality of plies for processing. Continuing with the foregoing example, where the first assembly is first assembly A<NUM>, the second plurality of plies may be angel wing plies <NUM> and flowpath plies <NUM>. In such embodiments, preparing the first assembly and second plurality of plies for processing may comprise laying up first assembly A<NUM>, angel wing plies <NUM>, and flowpath plies <NUM> on second layup tool <NUM>. First assembly A<NUM> may define a first plane and the second plurality of plies, e.g., angel wing and flowpath plies <NUM>, <NUM>, may define a second plane. Preparing first assembly A<NUM> and the second plurality of plies for processing may include positioning plies <NUM>, <NUM> adjacent first assembly A<NUM> on second layup tool <NUM> such that the second plane extends perpendicular to the first plane. The positions of first assembly A<NUM> and the second plurality of plies are depicted in <FIG> and <FIG>. Further, as described with respect to method <NUM>, the second layup tool, e.g., second layup tool <NUM>, may have any appropriate shape and configuration for supporting the first assembly and the second plurality of plies and/or aiding in the processing of the first assembly and the second plurality of plies. Other ways or methods also may be used to prepare the first assembly and the second plurality of plies for processing.

After being prepared for processing, as shown at <NUM>, the first assembly and second plurality of plies are processed to join or bond the first assembly and second plurality and thereby form a second assembly. For example, the first assembly and the second plurality of plies may be cured in an autoclave to form the second assembly. Alternatively or additionally, the first assembly and second plurality of plies may be compacted in an autoclave as part of processing to form the second assembly. In one embodiment, the first assembly and second plurality of plies may be processed in an autoclave using a standard cycle, including compacting the second plurality of plies along a second direction D<NUM> and/or a third direction D<NUM> and curing the assembly at an appropriate temperature and pressure. Where processing the first assembly and second plurality of plies follows processing the first plurality of plies by applying a compaction force along the first direction D<NUM>, the first direction D<NUM>, second direction D<NUM>, and third direction D<NUM> may be orthogonal to each other. That is, the first direction D<NUM> may be perpendicular to the second direction D<NUM>, the second direction D<NUM> may be perpendicular to the third direction D<NUM>, and the third direction D<NUM> may be perpendicular to the first direction D<NUM>.

In the above example, where the first assembly is first assembly A<NUM> and the second plurality of plies includes angel wing plies <NUM> and flowpath plies <NUM>, processing first assembly A<NUM> and the second plurality of plies forms second assembly A<NUM>, comprising airfoil <NUM>, dovetail <NUM>, platform <NUM>, angel wings <NUM>, and flowpath portions <NUM> of turbine rotor blade <NUM> as previously described. That is, the angel wing plies <NUM> and flowpath plies <NUM> may be joined or bonded to first assembly A<NUM> to form second assembly A<NUM>. An exemplary second assembly A<NUM> is shown in <FIG>.

Although described above with respect to fabricating LP turbine rotor blade <NUM>, it should be readily understood that method <NUM> also may be used to fabricate other gas turbine engine components. As one example, method <NUM> may be used to form HP turbine rotor blades <NUM>, as well as other components of turbofan engine <NUM>. Further, as previously described, other gas turbine engine components may comprise CMC plies extending within or parallel more than two planes or requiring compaction along a plurality of directions. Accordingly, the layup tools or other tooling used to process the component may have an appropriate configuration for supporting the component during processing, and method <NUM> may be adjusted to have any appropriate or desirable number of processing steps.

Claim 1:
A method for fabricating a ceramic matrix composite component of a gas turbine engine, the method comprising:
molding (<NUM>) a plurality of plies of a ceramic matrix composite material to form a first portion of the gas turbine engine component, wherein the first portion of the gas turbine engine component comprises an airfoil (<NUM>), a dovetail (<NUM>), and a platform (<NUM>) of a turbine rotor blade (<NUM>), and wherein the plurality of plies in the airfoil, the dovetail, and the platform extend parallel to a first plane;
processing (<NUM>) the first portion of the gas turbine engine component to form a first assembly (A<NUM>);
preparing (<NUM>) the first assembly (A<NUM>) and a second portion of the gas turbine engine component for processing;
processing (<NUM>) the first assembly (A<NUM>) and second portion of the gas turbine engine component to join the first assembly (A<NUM>) and the second portion and thereby form a second assembly (A<NUM>); and
wherein the first assembly (A<NUM>) defines the first plane and the second portion defines a second plane, and wherein preparing (<NUM>) the first assembly (A<NUM>) and the second portion for processing comprises positioning the second portion adjacent the first assembly (A<NUM>) on a second layup tool (<NUM>) such that the second plane extends perpendicular to the first plane.