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 composite materials including ceramic matrix composite (CMC) materials, are being used for various components within gas turbine engines. For example, because CMC materials can withstand relatively extreme temperatures, there is particular interest in replacing components within the flow path of the combustion gases with CMC materials. However, typical methods of constructing components from plies of a material, such as plies of a CMC material, often result in components having local variations in thickness and/or unsmooth surfaces. Components that have a non-uniform thickness can negatively impact tolerances such that the components do not line up with or are difficult to line up with one or more interfacing components. Additionally, components with non-uniform thickness within the component present manufacturing difficulties. For example, local variations in component thickness can result in inconsistent machining of holes or apertures in the component. <CIT> relates to a manufacturing process for heat shield covering consisting of assembling a series of components with at least one layer of fibers. <CIT> relates to a method of forming a composite component with off-set sub-layers of plies.

Accordingly, a method for forming a component that reduces or eliminates variations in component thickness would be desirable. In particular, a method for forming an axisymmetric component that reduces or eliminates variations in component thickness while providing a component having hoop strength would be advantageous. Axisymmetric components having a substantially uniform thickness throughout the component also would be useful.

In one exemplary aspect of the present subject matter, a method for forming a component according to claim <NUM> is provided.

In another exemplary aspect of the present subject matter, a component according to claim <NUM> is provided.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> is a schematic cross-sectional view of a turbomachine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of <FIG>, the turbomachine is configured as a gas turbine engine, or rather as 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), a radial direction R, and a circumferential direction C (<FIG>), i.e., a direction extending about the axial direction A. 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 and the core turbine engine <NUM> includes, 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>. Accordingly, the LP shaft <NUM> and HP shaft <NUM> are each rotary components, rotating about the axial direction A during operation of the turbofan engine <NUM>.

Referring still to the embodiment of <FIG>, the fan section <NUM> includes a 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. The fan blades <NUM> and disk <NUM> are together rotatable about the longitudinal axis <NUM> by LP shaft <NUM>. More particularly, the fan section includes a fan shaft rotatable by the LP shaft <NUM>. Accordingly, the fan shaft may also be considered a rotary component and is similarly supported by one or more bearings.

Referring still to the exemplary embodiment of <FIG>, the disk <NUM> is covered by a front hub <NUM> aerodynamically contoured to promote an airflow through the plurality of fan blades <NUM>. Front hub <NUM> may rotate with fan <NUM> about longitudinal centerline <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>. The exemplary nacelle <NUM> is 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> extends 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 <NUM> of the air <NUM> is directed or routed into the bypass airflow passage <NUM> and a second portion <NUM> of the air <NUM> is directed or routed into the core air flowpath <NUM>, or more specifically into the LP compressor <NUM>. The ratio between the first portion <NUM> of air and the second portion <NUM> of air is commonly known as a bypass ratio. The pressure of the second portion <NUM> of air 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 <NUM> of air <NUM> is substantially increased as the first portion <NUM> of air 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>.

It should be appreciated that the exemplary turbofan engine <NUM> depicted in <FIG> is by way of example only. In other exemplary embodiments, the turbofan engine <NUM> may have any other suitable configuration.

Referring now to <FIG>, a schematic, cross-sectional view is provided of a combustor assembly <NUM> according to an exemplary embodiment of the present subject matter. More particularly, <FIG> provides a side, cross-sectional view of an exemplary combustor assembly <NUM>, which may, for example, be positioned in the combustion section <NUM> of the exemplary turbofan engine <NUM> of <FIG>.

Combustor assembly <NUM> depicted in <FIG> generally includes a combustion chamber <NUM> defined by an inner liner <NUM> and an outer liner <NUM>, e.g., combustion liners <NUM>, <NUM> together at least partially define combustion chamber <NUM> therebetween. Combustion liners <NUM>, <NUM>, or other components of combustor assembly <NUM>, may be made from a ceramic matrix composite (CMC) material as further described below. Combustor assembly <NUM> extends generally along the axial direction A from a forward end <NUM> to an aft end <NUM>. Inner liner <NUM> generally defines a hot side <NUM> exposed to and defining in part a portion of the core air flowpath <NUM> extending through the combustion chamber <NUM>. Inner liner <NUM> further defines a cold side <NUM> opposite hot side <NUM>. Similarly, outer liner <NUM> also defines a hot side <NUM> exposed to and defining in part a portion of the core air flowpath <NUM> extending through the combustion chamber <NUM>, and outer liner <NUM> further defines a cold side <NUM> opposite hot side <NUM>.

The inner and outer liners <NUM>, <NUM> are each attached to an annular dome <NUM> at the forward end <NUM> of combustor assembly <NUM>. More particularly, annular dome <NUM> includes an inner dome section <NUM> attached to inner liner <NUM> and an outer dome section <NUM> attached to outer liner <NUM>. The inner and outer dome sections <NUM>, <NUM> may each extend along the circumferential direction C to define an annular shape. Inner and outer dome sections <NUM>, <NUM> each also define a slot <NUM> for receipt of inner liner <NUM> and outer liner <NUM>, respectively.

The combustor assembly <NUM> further includes a plurality of fuel air mixers <NUM> spaced along the circumferential direction C and positioned at least partially within the annular dome <NUM>. More particularly, the plurality of fuel air mixers <NUM> are disposed at least partially between outer dome section <NUM> and inner dome section <NUM> along the radial direction R. Compressed air from the compressor section of the turbofan engine <NUM> flows into or through the fuel air mixers <NUM>, where the compressed air is mixed with fuel and ignited to create the combustion gases <NUM> within the combustion chamber <NUM>. The inner and outer dome sections <NUM>, <NUM> are configured to assist in providing the flow of compressed air from the compressor section into or through the fuel air mixers <NUM>. For example, inner dome section <NUM> includes an inner cowl <NUM>, and outer dome section <NUM> similarly includes an outer cowl <NUM>. The inner and outer cowls <NUM>, <NUM> may assist in directing the flow of compressed air from the compressor section into or through one or more of the fuel air mixers <NUM>.

In certain exemplary embodiments, the inner dome section <NUM> with inner cowl <NUM> may be formed integrally as a single annular component, and similarly, the outer dome section <NUM> with outer cowl <NUM> also may be formed integrally as a single annular component. It should be appreciated, however, that in other exemplary embodiments, the inner dome section <NUM> and/or the outer dome section <NUM> alternatively may be formed by one or more components being joined in any suitable manner. For example, with reference to the outer dome section <NUM>, in certain exemplary embodiments, outer cowl <NUM> may be formed separately from outer dome section <NUM> and attached to outer dome section <NUM> using, e.g., a welding process. Additionally or alternatively, the inner dome section <NUM> may have a similar configuration.

Referring still to <FIG>, the exemplary combustor assembly <NUM> further includes a heat shield <NUM> positioned around the fuel air mixer <NUM> as depicted. The exemplary heat shield <NUM>, for the embodiment depicted, is attached to and extends between inner and outer dome sections <NUM>, <NUM>. The heat shield <NUM> is configured to protect certain components of the turbofan engine <NUM> from the relatively extreme temperatures of the combustion chamber <NUM>.

Keeping with <FIG>, combustor assembly <NUM> at the aft end <NUM> includes an inner piston ring seal <NUM> at inner liner <NUM> and an outer piston ring seal <NUM> at outer liner <NUM>. The inner piston ring seal <NUM> is attached to an inner piston ring holder <NUM> extending from and attached to an inner casing <NUM>. Similarly, the outer piston ring seal <NUM> is attached to an outer piston ring holder <NUM> extending from and attached to an outer casing <NUM>. Inner piston ring holder <NUM> and outer piston ring holder <NUM> are configured to accommodate an expansion of the inner liner <NUM> and the outer liner <NUM> generally along the axial direction A, as well as generally along the radial direction R. To allow for a relative thermal expansion between the outer liner <NUM> and the outer dome section <NUM>, as well as between the inner liner <NUM> and the inner dome section <NUM>, a plurality of mounting assemblies <NUM> are used to attach outer liner <NUM> to outer dome section <NUM> and inner liner <NUM> to inner dome section <NUM>. More particularly, the mounting assemblies <NUM> attach the forward end of outer liner <NUM> to outer dome section <NUM> within the slot <NUM> of outer dome section <NUM> and the forward end of inner liner <NUM> to inner dome section <NUM> within the slot <NUM> of inner dome section <NUM>. As further described herein, when formed from CMC materials, inner and outer liners <NUM>, <NUM> may be formed to hold dimensional tolerances and/or to minimize variations in the thickness of the liners, e.g., to ensure proper fitment with inner and outer piston ring seals <NUM>, <NUM> and slots <NUM> of dome sections <NUM>, <NUM> such that inner and outer liners <NUM>, <NUM> may thermally expand without damaging the liners and/or any adjacent components.

Further, as is discussed above, the combustion gases <NUM> flow from the combustion chamber <NUM> into and through the turbine section of the turbofan engine <NUM>, where a portion of thermal and/or kinetic energy from the combustion gases <NUM> is extracted via sequential stages of turbine stator vanes and turbine rotor blades. A stage one (<NUM>) stator vane <NUM> is depicted schematically in <FIG>, aft of the combustor assembly <NUM>.

In some embodiments, components of turbofan engine <NUM>, particularly components within hot gas path <NUM>, may comprise a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. For the depicted embodiment, inner liner <NUM> and outer liner <NUM> of combustor <NUM> are each formed of a CMC material. Exemplary CMC materials utilized for such components 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 Dow Corning'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, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). As further examples, the CMC materials may also include silicon carbide (SiC) or carbon fiber cloth. Of course, other appropriate materials, including other composite materials, may be used to form the components of engine <NUM>, including the components within hot gas path <NUM> such as inner and outer combustion liners <NUM>, <NUM>.

Turning to <FIG>, a schematic illustration is provided of an exemplary CMC component preform <NUM>, e.g., a preform for forming a CMC component of turbofan engine <NUM> such as outer liner <NUM> of combustor <NUM>. Although the schematic illustration of <FIG> depicts the CMC component preform <NUM> as generally cylindrical, it will be appreciated that preform <NUM> may have any appropriate shape for forming the resultant CMC component. In particular embodiments, the resultant CMC component may be axisymmetric, i.e., symmetric about an axis A-A, which may be longitudinal axis <NUM> of engine <NUM> in embodiments in which the resultant CMC component is a component of engine <NUM>. Accordingly, to form an axisymmetric CMC component, the component preform <NUM> may be axisymmetric, e.g., as shown in <FIG>, preform <NUM> is symmetric about axis A-A. However, in other embodiments, the CMC component may have other shapes or configurations, and preform <NUM> has a shape or contour appropriate for forming the CMC component. As further shown in <FIG>, the preform <NUM> defines a radial direction R' that is orthogonal to axis A-A, as well as a circumferential direction C' that extends about axis A-A. In some embodiments where preform <NUM> results in a component of engine <NUM>, the radial direction R' may be the radial direction R and the circumferential C' may be the circumferential direction C defined by engine <NUM>.

Referring now to <FIG>, CMC component preform <NUM> may be made from a plurality of CMC plies, in particular, a plurality of first CMC plies <NUM> and a plurality of second CMC plies <NUM>. In the exemplary embodiment shown in <FIG>, the plurality of CMC plies are arranged such that second CMC plies <NUM> overlap first CMC plies <NUM>. More specifically, each of first CMC plies <NUM> has ends 302a, 302b and each of second CMC plies <NUM> has ends 304a, 304b. Ends 304a, 304b of second CMC plies <NUM> overlap ends 302a, 302b of first CMC plies <NUM> such that each end 304a, 304b of each second CMC ply <NUM> overlies an end 304a or 304b of a first CMC ply <NUM>. By overlapping ends 302a, 302b with ends 304a, 304b, first CMC plies <NUM> are in contact with second CMC plies <NUM> along the ends of the plies <NUM>, and second plies <NUM> are in contact with first plies <NUM> along the ends of the plies <NUM>. The overlapping ends of first and second CMC plies <NUM>, <NUM> define a plurality of overlap regions <NUM> that are spaced apart from one another along the circumferential direction C'. Further, the first and second CMC plies <NUM>, <NUM> are arranged such that the ceramic fibers within each ply <NUM>, <NUM> are oriented along the circumferential direction C'. That is, when laid up to form CMC component preform <NUM>, the ceramic fibers within first CMC plies <NUM> are oriented in the same direction as the ceramic fibers within second CMC plies <NUM>, and in the exemplary preform <NUM>, the fibers of first and second plies <NUM>, <NUM> are oriented along the circumferential direction C'. Overlapping first and second plies <NUM>, <NUM> such that their ceramic fibers are overlapped in the circumferential or hoop direction can provide the structure needed to carry mechanical hoop strength in the resultant CMC component because the overlapped plies help ensure there are no breaks or cuts in the fibers extending along the circumferential direction C'.

First and second CMC plies <NUM>, <NUM> may be, e.g., plies pre-impregnated (prepreg) with matrix material and may be formed from prepreg tapes or the like. For example, the CMC plies may be formed from a prepreg tape comprising a desired ceramic fiber reinforcement material, one or more precursors of the CMC matrix material, and organic resin binders. According to conventional practice, prepreg tapes can be formed by impregnating the reinforcement material with a slurry that contains the ceramic precursor(s) and binders. The slurry also may contain solvents for the binders that promote the fluidity of the slurry to enable impregnation of the fiber reinforcement material, as well as one or more particulate fillers intended to be present in the ceramic matrix of the CMC component, e.g., silicon and/or SiC powders in the case of a Si-SiC matrix. Preferred materials for the precursor will depend on the particular composition desired for the ceramic matrix of the CMC component. For example, the precursor material may be SiC powder and/or one or more carbon-containing materials if the desired matrix material is SiC; notable carbon-containing materials include carbon black, phenolic resins, and furanic resins, including furfuryl alcohol (C<NUM>H<NUM>OCH<NUM>OH). Of course, the plurality of CMC plies <NUM>, <NUM> may formed in other ways as well.

Referring still to <FIG>, end 304b of second plies <NUM> overlies or overlaps end 302a of first plies <NUM> such that ends 304b, 302a are in contact for a contact length <NUM>. End 304a of second plies <NUM> overlies or overlaps end 302b of first plies <NUM> such that ends 304a, 302b are in contact for contact length <NUM>. More particularly, each first CMC ply <NUM> has an inner surface <NUM> and an outer surface <NUM>, and each second CMC ply <NUM> has an inner surface <NUM> and an outer surface <NUM>. As illustrated in <FIG>, inner surface <NUM> of second ply <NUM> contacts outer surface <NUM> of first ply <NUM> along the contact length <NUM> at ends 304b, 302a and 304a, 302b such that the ends 304a, 304b of second ply <NUM> overlap the ends 302a, 302b of first ply <NUM>. Each of the overlapping ends defines an overlap region <NUM>, which has a width equal to the contact length <NUM> over which the ends 304b, 302a and 304a, 302b are in contact.

Further, as depicted in <FIG>, first CMC plies <NUM> and second CMC plies <NUM> are overlapped such that the plurality of CMC plies <NUM>, <NUM> extends along the circumferential direction C'. By overlapping the ends of adjacent plies <NUM>, <NUM>, when CMC component preform <NUM> undergoes processing, e.g., to debulk and densify the CMC material as described below, a continuous ring of ceramic fibers may be oriented along the circumferential direction C', which helps to impart hoop strength to the resultant CMC component. Moreover, although shown as been substantially equal, each contact length <NUM>, and accordingly, the width of each overlap region <NUM>, need not be equal. However, each contact length <NUM> should be a sufficient length to ensure first plies <NUM> securely join with second plies <NUM> to form the continual ring of fibers along the circumferential direction C'.

Overlapped first and second CMC plies <NUM>, <NUM> define a layer <NUM> of CMC plies, and a plurality of layers <NUM> may be used to form CMC component preform <NUM>. For example, as shown in <FIG>, the plurality of CMC plies may be arranged in a first layer 318a, a second layer 318b, a third layer 318c, a fourth layer 318d, and a fifth layer 318e. A layer of axial CMC plies <NUM> is disposed between each layer <NUM> of first and second plies <NUM>, <NUM>. Axial plies <NUM> have ceramic fibers that extend generally axially within CMC component preform <NUM>, i.e., the fibers extend generally parallel to axis A-A, about which preform <NUM> is symmetric. In embodiments in which component preform <NUM> is an outer liner preform for forming outer liner <NUM> of engine <NUM>, axis A-A may be longitudinal centerline <NUM> of engine <NUM> or axis A-A may be parallel to longitudinal centerline <NUM>. Further, axial plies <NUM> may be laid up such that butt joints are formed between adjacent axial plies <NUM>. That is, adjacent axial plies <NUM> may be butted up against one another along their edges such that, unlike first and second CMC plies <NUM>, <NUM>, an axial CMC ply <NUM> does not overlap an adjacent axial ply <NUM>, although adjacent axial plies <NUM> may be in contact with one another.

Further, as illustrated in <FIG>, each layer <NUM> of first and second CMC plies <NUM>, <NUM> defines a plurality of overlap regions <NUM>. More particularly, first layer 318a of CMC plies defines first overlap regions 306a, second layer 318b of CMC plies defines second overlap regions 306b, third layer 318c of CMC plies defines third overlap regions 306c, fourth layer 318d of CMC plies defines fourth overlap regions 306d, and fifth layer 318e of CMC plies defines fifth overlap regions 306e. Each layer <NUM> is positioned such that no two overlap regions <NUM> are radially aligned. For example, second layer 318b is positioned with respect to first layer 318a such that second overlap regions 306b are offset from first overlap regions 306a along the circumferential direction C'. That is, none of the plurality of second overlap regions 306b is radially aligned with a first overlap region 306a. Similarly, second layer 318b is positioned with respect to third, fourth, and fifth layers 318c, 318d, 318e such that second overlap regions 306b also are offset from third overlap regions 306c, fourth overlap regions 306d, and fifth overlap regions 306e along the circumferential direction C'. More specifically, each layer <NUM> is positioned with respect to the other layers <NUM> such that no overlap region <NUM> is radially aligned with another overlap region <NUM>. Accordingly, any radial line <NUM> drawn through layers 318a, 318b, 318c, 318d, 318e and the layers of axial plies <NUM>, which form CMC component preform <NUM>, passes through the same number of plies <NUM>, <NUM>, <NUM> and does not pass through more than one overlap region <NUM>.

It will be appreciated that <FIG> is provided by way of example only. In other embodiments, layers <NUM> may be circumferentially and/or radially arranged in any order, including but not limited to the order shown in <FIG>. However, regardless of the order in which the layers are arranged, layers <NUM> are positioned such that no overlap region <NUM> is radially aligned with another overlap region <NUM>. Further, although illustrated in <FIG> as including five layers <NUM> with one layer of axial plies <NUM> between layers <NUM>, in other embodiments the CMC component preform <NUM> may include any suitable number of layers <NUM> with any suitable number of axial plies <NUM> between the layers <NUM>.

Referring still to <FIG>, it will be appreciated that the plurality of CMC plies may be laid up on a layup tool <NUM> to form CMC component preform <NUM>. Although shown as generally linear in the schematic illustration of <FIG>, tool <NUM> may be generally cylindrical in shape with an inner surface <NUM> and an outer surface <NUM>. In such embodiments, the CMC plies may be laid up on or against the generally cylindrical inner surface <NUM> of tool <NUM> to form a generally cylindrical, axisymmetric CMC component preform <NUM>. As such, each successive layer <NUM> of first and second CMC plies <NUM>, <NUM> may be positioned radially inward with respect to the preceding layer <NUM>. In other embodiments, tool <NUM> may have any appropriate shape or configuration for laying up the plurality of CMC plies to define the shape or contour of CMC component preform <NUM>, and the plurality of CMC plies <NUM>, <NUM> may be laid up on inner surface <NUM> or outer surface <NUM> of tool <NUM> as appropriate.

Turning now to <FIG>, a schematic cross-section view of outer liner <NUM> of <FIG> is provided. In the depicted exemplary embodiment of the present subject matter, outer liner <NUM> is a CMC body <NUM> that is symmetric about longitudinal centerline <NUM> of engine <NUM>, i.e., centerline <NUM> is the axis of symmetry for CMC body <NUM>. The CMC body <NUM> defines the circumferential direction C that extends about the centerline <NUM>, as well as radial direction R, which is orthogonal to centerline <NUM>. The CMC body <NUM> is formed from a plurality of CMC plies <NUM>, <NUM>, <NUM> as described above. More particularly, the CMC plies are laid up on a layup tool to define a generally smooth outer surface <NUM> of outer liner <NUM>/CMC body <NUM>. Further, the ends of a plurality of first CMC plies <NUM> and second CMC plies <NUM> are overlapped to define a plurality of overlap regions <NUM>. Each overlap region <NUM> formed between contacting, overlapped plies <NUM>, <NUM> is offset along the circumferential direction C such that no two overlap regions <NUM> are aligned along any radial line <NUM> drawn through longitudinal centerline <NUM>. That is, any radial line <NUM> drawn through centerline <NUM> passes through only one overlap region <NUM>. Further, as described above, each of first and second CMC plies <NUM>, <NUM> comprises a plurality of ceramic fibers that are oriented substantially along one direction. The CMC plies <NUM>, <NUM> forming CMC body <NUM> are positioned such that the ceramic fibers of first and second CMC plies <NUM>, <NUM> are oriented along the circumferential direction C. Thus, overlapping ends of the plurality of first and second CMC plies <NUM>, <NUM>, which may be arranged in a plurality of layers <NUM> with axial plies <NUM> disposed between each layer <NUM> as previously discussed, forms a continuous ring of ceramic fibers in the circumferential direction C.

Accordingly, by offsetting overlap regions <NUM>, the inner surface <NUM> of outer liner <NUM> generally is smooth and the CMC body <NUM> defining outer liner <NUM> has a substantially uniform thickness. More particularly, overlapping plies are not built up along the radial direction R, which could result in a liner <NUM> having one or more sections that are thicker than other sections of the liner such that the liner does not have a uniform thickness. Rather, the overlap regions <NUM> are distributed throughout the component such that liner <NUM> has a substantially uniform thickness and a generally smooth surface opposite the tooled surface, which is also a generally smooth surface if the tool on which the plies were laid up is substantially smooth. The generally smooth surfaces and substantially uniform thickness can help outer liner <NUM> properly interface with interfacing components such as outer dome section <NUM> and/or outer piston ring seal <NUM> as described above, as well as provide a smooth flow path for combustion gases <NUM>. Further, while <FIG> depicts outer liner <NUM>, it should be appreciated that inner liner <NUM> or any other appropriate component of engine <NUM> may be formed in substantially the same manner as the depicted outer liner <NUM> such that outer liner <NUM>, inner liner <NUM>, or other CMC component of engine <NUM> has a substantially uniform thickness. Additionally or alternatively, other components that are made using CMC materials, particularly axisymmetric CMC components, may be formed as described with respect to outer liner <NUM>.

<FIG> provides a flow diagram illustrating a method <NUM> for forming a CMC component according to an exemplary embodiment of the present subject matter. For example, the CMC component may be outer liner <NUM> for combustor <NUM> of gas turbine engine <NUM>. As shown at <NUM> in <FIG>, a plurality of plies of a CMC material for forming the CMC component may be laid up to form CMC component preform <NUM> having a desired shape or contour. The desired shape of CMC component preform <NUM> may be a desired shape or contour of the resultant CMC component. As an example, the plies may be laid up to define a shape of CMC component preform <NUM> that is the shape of outer liner <NUM>. As described above, the plurality of CMC plies <NUM>, <NUM>, <NUM> forming preform <NUM> may be laid up on layup tool <NUM> or may be laid up on another appropriate device for supporting the plies and/or for defining the desired shape.

Further, as described above, laying up the plurality of first and second CMC plies <NUM>, <NUM> may comprise orienting the ceramic fibers of the CMC plies <NUM>, <NUM> along the circumferential direction C', as shown at 802a in <FIG>. Further, as shown at 802b, laying up first and second CMC plies <NUM>, <NUM> may include overlapping ends of plies <NUM>, <NUM> to form a plurality of overlap regions <NUM>. Overlapping ends of the plurality of CMC plies includes positioning ends 302a, 302b of the plurality of first CMC plies <NUM> in contact with ends 304a, 304b of the plurality of second CMC plies <NUM>. More particularly, each end of each second CMC ply <NUM> may be positioned to contact an end of a first CMC ply <NUM> over contact length <NUM>. Each contact length <NUM> defines the width of the corresponding overlap region <NUM>. By overlapping ends of plies <NUM>, <NUM> with the ceramic fibers of the plies oriented along the circumferential direction C', a continuous ring of ceramic fibers may be formed along the circumferential direction C'.

Moreover, plies <NUM>, <NUM> may be laid up in a plurality of layers <NUM>, each layer having a plurality of overlapped first and second CMC plies <NUM>, <NUM> and, thus, a plurality of overlap regions <NUM>. Referring still to <FIG>, as shown at 802c, laying up plies <NUM>, <NUM> in layers <NUM> further may comprise offsetting overlap regions <NUM> of each layer <NUM> such that any radial line <NUM> drawn through axis A-A of CMC component preform <NUM> does not pass through more than one overlap region <NUM>. Additionally, laying up the plurality of CMC plies may comprise laying up axial plies <NUM> between layers <NUM> of first and second plies <NUM>, <NUM>. That is, layers <NUM> of first and second plies <NUM>, <NUM> may be alternated with layers of axial plies <NUM>. More particularly, layers <NUM> of plies <NUM>, <NUM> are laid up to alternate with layers of axial plies <NUM> and to offset overlap regions <NUM> of layers <NUM> along the circumferential direction C' such that any radial line <NUM> drawn through axis A-A and the layup passes through the same number of plies <NUM>, <NUM>, <NUM> and passes through only one overlap region <NUM>.

After the plurality of plies <NUM>, <NUM>, <NUM> are laid up, the plies may be processed, e.g., compacted and cured in an autoclave, as shown at <NUM> in <FIG>. After processing, the plies form a green state CMC component, e.g., a green state CMC outer liner <NUM>. The green state CMC component is a single piece component, i.e., curing plies <NUM>, <NUM>, <NUM> joins the plies to produce a CMC component formed from a continuous piece of CMC material. The green state component then may undergo firing (or burn-off) and densification, illustrated at <NUM> and <NUM> in <FIG>, to produce a final CMC component. In an exemplary embodiment of method <NUM>, the green state component is placed in a furnace with silicon to burn off any mandrel-forming materials and/or solvents used in forming the CMC plies <NUM>, <NUM>, <NUM>, to decompose binders in the solvents, and to convert a ceramic matrix precursor of the plies into the ceramic material of the matrix of the CMC component. The silicon melts and infiltrates any porosity created with the matrix as a result of the decomposition of the binder during burn-off/firing; the melt infiltration of the CMC component with silicon densifies the CMC component. However, densification may be performed using any known densification technique including, but not limited to, Silcomp, melt-infiltration (MI), chemical vapor infiltration (CVI), polymer infiltration and pyrolysis (PIP), and oxide/oxide processes. In one embodiment, densification and firing may be conducted in a vacuum furnace or an inert atmosphere having an established atmosphere at temperatures above <NUM>° C to allow silicon or another appropriate material or materials to melt-infiltrate into the component <NUM>. Optionally, as shown at <NUM> in <FIG>, after firing and densification the CMC component may be finish machined, if and as needed, and/or coated with an environmental barrier coating (EBC).

Claim 1:
A method for forming a component, the method comprising:
laying up a plurality of ceramic matrix composite, CMC, plies to form a CMC component preform (<NUM>), wherein the plurality of CMC plies comprises a plurality of first CMC plies and a plurality of second CMC plies, each of the first and second CMC plies comprising a plurality of ceramic fibers, the CMC component preform defining an axis of symmetry and a circumferential direction, wherein laying up the plurality of CMC plies includes
overlapping ends of the plurality of CMC plies, each overlapped end defining an overlap region (<NUM>), and characterized by
offsetting the overlap regions along the circumferential direction such that any radial line drawn from the axis of symmetry through the plurality of CMC plies passes through only one overlap region,
wherein the first and the second CMC plies are arranged such that the ceramic fibers within each ply are oriented along the circumferential direction.