Patent Description:
Gas turbine engine blades can be made out of composite material obtained from fiber reinforcement made by three-dimensional weaving and densified with a matrix. Three-dimensional (3D) or multilayer weaving serves to give the composite material blade mechanical strength. Nevertheless, the good mechanical strength imparted by three-dimensional or multilayer weaving is more difficult to obtain in portions of the blade that present small thicknesses, of the order of <NUM> millimeter (mm) to <NUM>, such as the leading and/or trailing edges of a blade. The leading edge and the trailing edge need to be capable of withstanding or limiting damage when they are subjected to various stresses such as bird strikes or repeated cycles in flight (erosion/lifetime). <CIT> discloses an airfoil including a core with a first Young's Modulus; and an outer section at least partially surrounding the core with a second Young's Modulus, wherein the first Young's Modulus is higher than the second Young's Modulus. <CIT> discloses a CMC airfoil (<NUM>) and a method of manufacturing (<NUM>-<NUM>) a CMC airfoil (<NUM>), where the airfoil (<NUM>) comprises a core (<NUM>) and a shell (<NUM>). The core (<NUM>) comprises core ceramic fibers (<NUM>) extending along a span (<NUM>) of the airfoil (<NUM>). The shell (<NUM>) surrounds the core (<NUM>) and includes shell ceramic fibers (<NUM>,<NUM>). Substantially all of the core ceramic fibers (<NUM>) are arranged in a radial direction (<NUM>). The shell ceramic fibers (<NUM>,<NUM>) may be woven or braided. The braided fibers may be arranged in a braid angle (<NUM>), preferably being between <NUM>° and <NUM>° with respect to the radial axis (<NUM>). The core ceramic fibers (<NUM>) are preferably made of a ceramic material having a higher creep resistance than the material forming the shell ceramic fibers (<NUM>,<NUM>). <CIT> discloses a ceramic matrix composite blade for use in a gas turbine engine is disclosed. The ceramic matrix composite blade includes a root, an airfoil, and a platform located between the root and the airfoil.

What is needed is a composite material blade that exhibits mechanical strength responsive to radial loading as well as being capable of withstanding damage proximate the leading and trailing edges.

In accordance with the present disclosure, there is provided a ceramic matrix composite blade comprising a central core surrounded by an outer profile, the central core comprising layers of unidirectional layup having two dimension and three dimension fiber weaves; the outer profile comprises a biased weave layup that radiates toward a leading edge and a trailing edge of the blade; wherein the biased weave layup includes fibers that extend proximate an attachment region of the blade radially and axially along an airfoil portion of the blade toward the leading edge and trailing edge.

Particular embodiments may include at least one of the following optional features. These embodiments may include one of these optional features, or a plurality of these optional features in combination, unless specified otherwise.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the unidirectional layup comprises unidirectional fibers with <NUM>-<NUM>% fibrous reinforcement.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the biased weave layup fibers are attached from a concave side of the blade around the leading edge to a convex side of the blade.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the biased weave layup fibers are attached from a convex side of the blade around the leading edge to a concave side of the blade.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the biased weave layup fibers are attached to extend from an attachment region radially and axially over an airfoil of the blade to a trailing edge of the blade.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the biased weave layup fibers are angled from a zero degree direction.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the angle from the zero degree direction ranges from about <NUM> degrees to about <NUM> degrees.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the biased weave layup includes an angle relative to the zero degree direction with respect to a camber line of the blade from the leading edge to the trailing edge.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the ceramic matrix composite blade further comprising a first edge portion of the outer profile from about <NUM> percent of the camber line proximate the leading edge comprises an angle that ranges from about <NUM> degrees to about <NUM> degrees from the zero degree direction.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the ceramic matrix composite blade further comprising a second edge portion of the outer profile from about <NUM> percent of the camber line proximate the trailing edge comprises an angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the ceramic matrix composite blade further comprising a middle portion of the outer profile located between the first edge portion and the second edge portion comprises about <NUM> percent to <NUM> percent of the camber line and includes an angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include an arrangement of the biased weave layup fibers are angled to obtain a predetermined load profile in each of the first edge region, second edge region and middle region.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the outer profile includes a structure in the biased weave layup that extends radially and spirals around contours of the airfoil at a predetermined angle, allowing a predetermined leading edge and trailing edge radii with predetermined fiber bend radii.

In accordance with the present disclosure, there is provided a process for reducing damage to a ceramic matrix composite blade comprising forming a central core, the central core comprising layers of unidirectional layup having two dimension and three dimension fiber weaves; forming an outer profile around the central core; and forming within the outer profile a biased weave layup that radiates toward a leading edge and a trailing edge of the blade; wherein the biased weave layup includes fibers that extend proximate an attachment region of the blade radially and axially along an airfoil portion of the blade toward the leading edge and trailing edge.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the process further comprising attaching the biased weave layup fibers from a concave side of the blade around the leading edge to a convex side of the blade; attaching the biased weave layup fibers from the convex side of the blade around the leading edge to the concave side of the blade; and attaching the biased weave layup fibers to extend from an attachment region radially and axially over an airfoil of the blade to a trailing edge of the blade.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively includethe process further comprising forming a first edge portion of the outer profile from about <NUM> percent of the camber line proximate the leading edge comprising the angle that ranges from about <NUM> degrees to about <NUM> degrees from the zero degree direction.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the process further comprising forming a second edge portion of the outer profile from about <NUM> percent of the camber line proximate the trailing edge comprising the angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction.

A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the process further comprising forming a middle portion of the outer profile located between the first edge portion and the second edge portion comprises about <NUM> percent to <NUM> percent of the camber line and includes an angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction.

Other details of the composite blade are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements.

The fan section <NUM> may include a single-stage fan <NUM> having a plurality of fan blades <NUM>. The fan blades <NUM> may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan <NUM> drives air along a bypass flow path B in a bypass duct <NUM> defined within a housing <NUM> such as a fan case or nacelle, and also drives air along a core flow path C for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. A splitter <NUM> aft of the fan <NUM> divides the air between the bypass flow path B and the core flow path C. The housing <NUM> may surround the fan <NUM> to establish an outer diameter of the bypass duct <NUM>. The splitter <NUM> may establish an inner diameter of the bypass duct <NUM>.

The inner shaft <NUM> is connected to the fan <NUM> through a speed change mechanism, which in the exemplary gas turbine engine <NUM> is illustrated as a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The inner shaft <NUM> may interconnect the low pressure compressor <NUM> and low pressure turbine <NUM> such that the low pressure compressor <NUM> and low pressure turbine <NUM> are rotatable at a common speed and in a common direction. In other embodiments, the low pressure turbine <NUM> drives both the fan <NUM> and low pressure compressor <NUM> through the geared architecture <NUM> such that the fan <NUM> and low pressure compressor <NUM> are rotatable at a common speed. Although this application discloses geared architecture <NUM>, its teaching may benefit direct drive engines having no geared architecture.

The low pressure compressor <NUM>, high pressure compressor <NUM>, high pressure turbine <NUM> and low pressure turbine <NUM> each include one or more stages having a row of rotatable airfoils. Each stage may include a row of static vanes adjacent the rotatable airfoils. The rotatable airfoils and vanes are schematically indicated at <NUM> and <NUM>.

The engine <NUM> may be a high-bypass geared aircraft engine. The bypass ratio can be greater than or equal to <NUM> and less than or equal to about <NUM>, or more narrowly can be less than or equal to <NUM>. The geared architecture <NUM> may be an epicyclic gear train, such as a planetary gear system or a star gear system. The epicyclic gear train may include a sun gear, a ring gear, a plurality of intermediate gears meshing with the sun gear and ring gear, and a carrier that supports the intermediate gears. The sun gear may provide an input to the gear train. The ring gear (e.g., star gear system) or carrier (e.g., planetary gear system) may provide an output of the gear train to drive the fan <NUM>. A gear reduction ratio may be greater than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, and in some embodiments the gear reduction ratio is greater than or equal to <NUM>. The fan diameter is significantly larger than that of the low pressure compressor <NUM>. The low pressure turbine <NUM> can have a pressure ratio that is greater than or equal to <NUM> and in some embodiments is greater than or equal to <NUM>. All of these parameters are measured at the cruise condition described below.

The flight condition of <NUM> Mach and <NUM>,<NUM> feet (<NUM>,<NUM> meters), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The engine parameters described above, and those in the next paragraph are measured at this condition unless otherwise specified.

"Low fan pressure ratio" is the pressure ratio across the fan blade <NUM> alone, without a Fan Exit Guide Vane ("FEGV") system. A distance is established in a radial direction between the inner and outer diameters of the bypass duct <NUM> at an axial position corresponding to a leading edge of the splitter <NUM> relative to the engine central longitudinal axis A. The low fan pressure ratio is a spanwise average of the pressure ratios measured across the fan blade <NUM> alone over radial positions corresponding to the distance. The low fan pressure ratio can be less than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, such as between <NUM> and <NUM>. The "low corrected fan tip speed" can be less than or equal to <NUM> ft / second (<NUM> meters/second), and greater than or equal to <NUM> ft / second (<NUM> meters/second).

Referring also to <FIG>, an exemplary composite blade is shown. It is noted that the exemplary composite blade is described, a vane can also include the material composition. Although the blade <NUM> is illustrated, the disclosure applies to all types of components made from a composite material and is not limited to the composite blade <NUM> illustrated. The blade <NUM> can be a ceramic matrix composite (CMC) gas turbine engine blade. The blade <NUM> includes an attachment or root portion <NUM> with a neck <NUM>. The neck <NUM> is not located in the hot working fluid flow path. The blade <NUM> is coupled with a turbine disk (not shown) proximate the root <NUM>. The blade includes a platform region, specifically a platform <NUM> disposed along an upper portion of the neck <NUM>. The platform region <NUM> defines the flowpath that separates the hot working fluid from the cooler working fluid proximate the neck <NUM>. The blade <NUM> further includes an airfoil <NUM> located opposite the root <NUM> relative to the platform <NUM> and extending radially outward from the platform <NUM>. The airfoil <NUM> includes a concave side <NUM> and an oppositely facing convex side <NUM>. The blade <NUM> includes an axially leading edge <NUM>, an axially trailing edge <NUM>, a radially outer side <NUM> and a radially inner side <NUM>, and an axially central portion <NUM>. The blade tip <NUM> is opposite the blade root <NUM>.

Referring also to <FIG>, <FIG> and <FIG>, the schematic cross sectional plan view of an exemplary composite blade <NUM> is shown at <FIG>. The cross section is cut through B-B of <FIG>. The blade <NUM> includes a central core <NUM> surrounded by an outer profile <NUM>.

The central core <NUM> can be made up of layers of unidirectional layup <NUM>, two dimension, and three dimension weaves. The unidirectional layup <NUM> is strongest when loaded in the zero direction (<NUM> degree direction), and the weakest when loaded in the ninety degree direction (<NUM> degree direction). The zero degree direction <NUM> is along the radial direction (coming out of the page) shown in Fig..

<NUM> that can include a range from -<NUM> degrees to <NUM> degrees and/or <NUM> degrees to <NUM> degrees in the radial direction. The <NUM> degree direction <NUM> is along the direction orthogonal and/or substantially orthogonal to the zero degree direction <NUM>, as shown coming out of the page. The <NUM> degree direction can range from <NUM> degrees to <NUM> degrees in a direction orthogonal/substantially orthogonal to the zero degree direction <NUM>. The axial direction <NUM> is generally indicated. The unidirectional layup <NUM> can comprise mostly unidirectional fibers <NUM> with <NUM>-<NUM>% fibrous reinforcement <NUM> "fiber fraction", as in the percent of the volume of a given region that is fiber. The remaining percent of a given volume consists of matrix + voids. Fibrous reinforcement <NUM> is achieved by three-dimensional weaving on a jacquard-type loom. During weaving, warp yarn bundles (or warp strands) are disposed in several layers of several hundred yarns each. Weft yarns (or weft strands) are interlaced with the warp yarns so as to bind the various layers of warp yarns together. The three-dimensional weaving is a weaving with an "interlock" pattern. By "interlock" is meant a weaving pattern in which each layer of weft yarns binds several layers of warp yarns with all the yarns of the same weft column having the same movement in the plane of the pattern. Other types of known three-dimensional weavings may be used.

The blade <NUM> includes a camber line <NUM> that defines the locus of the mid-points between the concave side <NUM> and convex side <NUM> surfaces when measured perpendicular to the camber line <NUM>. In this disclosure, the camber line <NUM> is being used to indicate general locations of various layups of fibers. The blade <NUM> also defines a stiffness neutral axis <NUM> and a geometric neutral axis <NUM> as shown at <FIG>.

The geometric neutral axis <NUM> describes the distribution of the airfoil section, relative to the radial "stacking line. " When airfoils are designed, the radial pull of the outermost sections must be supported by the successive inner sections. The opposite of a layer cake, you start at the tip <NUM>, and stack the airfoil <NUM> sections under the outer portions, as needed to carry the load, with the desired stress states. What happens when one uses a stiff core <NUM> and a softer shell outer profile <NUM>, is the location of the core <NUM>, needed to support whole airfoil <NUM> is no longer in the geometric center <NUM>. Plus, the core <NUM> shape is somewhat limited by the thickness of the shell <NUM>, and the resulting core section <NUM> is less airfoil-like and more like a curved beam. A different approach has to be taken to locate the central core <NUM>, such that the radial pull and the bending is balanced to create a stress state that does two things, sets the core <NUM> within allowable stress levels, and in the case of the shell <NUM>, intentionally drive the shell <NUM> into as much compression as possible. With the shell <NUM> leading edge <NUM> and trailing edge <NUM> in compression, and most of the shell <NUM> at low stress, the ability to be damaged, without overloading the core <NUM> is increased.

The descriptive method for locating the airfoil <NUM> is the section neutral axis <NUM>. But with the core <NUM> and shell <NUM> approach, the location of the geometric centroid and the structural centroid are no longer the same. The idea is to give guidance on where to locate the core <NUM> section, to get maximum benefit. Put as much load in the core <NUM>, with good stress distribution. Unload the shell <NUM>, but more importantly, make the stress state such that damage has a much lower risk of initiating a blade fracture.

The leading edge <NUM> and trailing edge <NUM> are located farther from the stiffness neutral axis <NUM> but have a lower stress due to a lower modulus.

The outer profile <NUM> can include biased weave layup <NUM> that radiates toward the leading edge <NUM> and trailing edge <NUM>. The outer profile <NUM> can include a low radial stiffness component. The outer profile <NUM>, in an exemplary embodiment can have a Young's modulus of <NUM>% to <NUM>% of the central core <NUM>.

As seen in <FIG>, the outer profile <NUM> includes the biased weave layup <NUM> fibers that extend proximate the attachment region <NUM> radially and axially along the airfoil portion <NUM> toward the leading edge <NUM> and trailing edge <NUM>. The biased weave layup <NUM> fibers can be wrapped from the convex side <NUM> around the leading edge <NUM> to the concave side <NUM>. The biased weave layup <NUM> fibers can be wrapped from the concave side <NUM> around the leading edge <NUM> to the convex side <NUM>. The biased weave layup <NUM> fibers can also be wrapped to extend from the attachment region <NUM> radially and axially over the airfoil <NUM> to the trailing edge <NUM>. The biased weave layup <NUM> can be arranged to extend from the radially inner side <NUM> to the radially outer side <NUM>. The load carried by the outer profile <NUM> can be angled radially into the attachment region <NUM>. Fibers of the biased weave layup <NUM> in the center of the camber line support the tip <NUM>. Fibers of the biased weave layup <NUM> near the forward region <NUM> and after region <NUM> of the attachment region <NUM> support the neck <NUM>.

The unique biased weave layup <NUM> in the outer profile <NUM> provides damage protection to the leading edge <NUM> and trailing edge <NUM> portions of the blade <NUM>. The impact damage <NUM> from an object (not shown) striking the leading edge <NUM> or trailing edge <NUM> can be mitigated due to the unique biased weave layup <NUM>. The outer profile <NUM> includes a structure in the outer weave profile <NUM> that flows radially and spirals around the contours of the airfoil at a steep angle, allowing a predetermined leading edge <NUM> and trailing edge <NUM> radii <NUM> with predetermined fiber bend radii <NUM>. The radii <NUM> can be less than <NUM>" trailing edge radii. Ceramic fiber bundles reach a limit around <NUM>-. <NUM>" radius, where bending any smaller results in large scale breakage. With a spiral wrap in the fiber, it will bend far less than the planer section radius. Impact damage <NUM> proximate the platform region <NUM> does not affect fibers of the biased weave layup <NUM> supporting airfoil portions proximate the tip <NUM>, because they are closer to the mid-chord of the blade <NUM>.

Referring again to <FIG>, the biased weave layup <NUM> fibers can be angled from the zero degree direction <NUM>. In an exemplary embodiment, the angle from the zero degree direction <NUM> can range from about <NUM> degrees to about <NUM> degrees. In an exemplary embodiment, the angle from the zero degree direction <NUM> can range from about <NUM> degrees. The biased weave layup <NUM> can vary the angle relative to the zero degree direction <NUM> relative to the camber line <NUM> from leading edge <NUM> to trailing edge <NUM>. In an exemplary embodiment, a first edge portion <NUM> of the outer profile <NUM> from about <NUM> percent of the camber line <NUM> proximate the leading edge <NUM> can include an angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction <NUM>. A second edge portion <NUM> of the outer profile <NUM> from about <NUM> percent of the camber line <NUM> proximate the trailing edge <NUM> can include an angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction <NUM>. A middle portion <NUM> of the camber line <NUM> located between the first edge portion <NUM> and second edge portion <NUM> and including about <NUM> percent to <NUM> percent of the camber line <NUM> can include an angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction <NUM>. The arrangement of the biased weave layup <NUM> fibers can be angled to obtain a predetermined load profile <NUM> in the various first edge region <NUM>, second edge region <NUM> and middle region <NUM>.

Referring to <FIG> a process diagram <NUM>. The first step <NUM> in the process <NUM> includes forming a central core <NUM>. The central core <NUM> comprises layers of unidirectional layup <NUM> having two dimension and three dimension fiber weaves. The next step <NUM> in the process includes forming an outer profile <NUM> over the central core <NUM>. The next step <NUM> includes installing biased weave layup <NUM> within the outer profile <NUM>. The biased weave layup <NUM> radiates toward a leading edge <NUM> and a trailing edge <NUM> of the blade <NUM>. The biased weave layup <NUM> includes fibers that extend proximate an attachment region <NUM> of the blade <NUM> radially and axially along an airfoil portion <NUM> of the blade <NUM> toward the leading edge <NUM> and trailing edge <NUM>. The process <NUM> can further include attaching the biased weave layup fibers <NUM> from a concave side <NUM> of the blade <NUM> around the leading edge <NUM> to a convex side <NUM> of the blade <NUM>. The process <NUM> can further include attaching the biased weave layup fibers <NUM> from the convex side <NUM> of the blade <NUM> around the leading edge <NUM> to the concave side <NUM> of the blade <NUM>. The process can further include attaching the biased weave layup fibers <NUM> to extend from an attachment region <NUM> radially and axially over an airfoil <NUM> of the blade <NUM> to a trailing edge <NUM> of the blade <NUM>. The biased weave layup <NUM> includes an angle relative to the zero degree direction <NUM> with respect to a camber line <NUM> of the blade from the leading edge <NUM> to the trailing edge <NUM>. The process <NUM> can further include forming a first edge portion <NUM> of the outer profile <NUM> from about <NUM> percent of the camber line <NUM> proximate the leading edge <NUM> comprising said angle that ranges from about <NUM> degrees to about <NUM> degrees from the zero degree direction <NUM>. The process <NUM> can further include forming a second edge portion <NUM> of the outer profile <NUM> from about <NUM> percent of the camber line <NUM> proximate the trailing edge <NUM> comprising the angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction <NUM>. The process <NUM> can further include forming a middle portion <NUM> of the outer profile <NUM> located between the first edge portion <NUM> and the second edge portion <NUM> that comprises about <NUM> percent to <NUM> percent of the camber line <NUM> and includes an angle that ranges from about <NUM> degrees to about <NUM> degrees angle from the zero degree direction <NUM>.

A technical advantage of the disclosed composite blade layup can include leading and trailing edges that are more tolerant to damage, because the stress state at the edges is low enough that stress concentrations created by damage do not overload the edges.

A technical advantage of the disclosed composite blade layup can include an airfoil composite weave layup and structural loading scheme which creates low stresses in the critical leading edge and trailing edge regions of the blade.

A technical advantage of the disclosed composite blade layup can include an airfoil composite layup the focuses stiffness distribution and balance, thus unloading the leading and trailing edges.

A technical advantage of the disclosed composite blade layup can include an airfoil composite layup that can tolerate a defect at the leading edge or trailing edge.

A technical advantage of the disclosed composite blade layup can include the capacity to keep the stress below a proportional limit by use of stiffness and the use of center of gravity to reduce the stress at the leading edge and the trailing edge.

Claim 1:
A ceramic matrix composite blade (<NUM>) comprising:
a central core (<NUM>) surrounded by an outer profile (<NUM>), said central core (<NUM>) comprising layers of unidirectional layup (<NUM>) having two dimension and three dimension fiber weaves; said outer profile (<NUM>) comprises a biased weave layup (<NUM>) that radiates toward a leading edge (<NUM>) and a trailing edge (<NUM>) of the blade; characterized in that said biased weave layup (<NUM>) includes fibers that extend proximate an attachment region (<NUM>) of the blade radially and axially along an airfoil portion (<NUM>) of the blade toward the leading edge (<NUM>) and trailing edge (<NUM>).