Patent Publication Number: US-2011052405-A1

Title: Composite airfoil with locally reinforced tip region

Description:
BACKGROUND 
     Composite materials offer potential design improvements in gas turbine engines. For example, in recent years composite materials have been replacing metals in gas turbine engine fan blades because of their high strength and low weight. Most metal gas turbine engine fan blades are titanium. The ductility of titanium fan blades enables the fan to ingest a bird and remain operable or be safely shut down. The same requirements are present for composite fan blades. 
     A composite airfoil for a turbine engine fan blade can have a sandwich construction with a carbon fiber woven core at the center and two-dimensional filament reinforced plies or laminations on either side. To form the composite airfoil, individual two-dimensional plies are cut and stacked in a mold with the woven core. The mold is injected with a resin using a resin transfer molding process and cured. The plies vary in length and shape. The carbon fiber woven core is designed to accommodate ply drops so that multiple plies do not end at the same location. 
     Each ply comprises a plurality of oriented elongated fibers. For example, a ply can comprise a woven material or a uniweave material. With a woven material, half of the woven fibers are oriented in a first direction and half the fibers are oriented in a direction 90° from the first direction. A uniweave material, on the other hand, has about 98% of its fibers oriented in a first direction and a small number of fibers extending in a direction 90° from the first direction to stitch the uniweave material together. 
     Previous composite blades have been configured to improve the impact strength of the composite airfoils so they can withstand bird strikes. During use, foreign objects ranging from large birds to hail may be entrained in the inlet of the gas turbine engine. Impact of large foreign objects can rupture or pierce the blades and cause secondary damage downstream of the blades. There are design drivers in addition to the ability to withstand bird strikes which will improve composite blades. 
     SUMMARY 
     A composite airfoil has a root, a tip, a root region and a tip region. The composite airfoil further includes a woven core, a first filament reinforced airfoil ply, a second filament reinforced airfoil ply and a local reinforcement laminate section. The woven core extends from the root to the tip of the composite airfoil. The first filament reinforced airfoil ply is stacked on the woven core and the second filament reinforced airfoil ply is stacked adjacent to the first filament reinforced airfoil ply on the woven core. The local reinforcement laminate section is at the tip region of the composite airfoil and comprises a first reinforcement ply that does not extend to the root region. The local reinforcement laminate section increases a chordwise flexural stiffness of the tip region. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a cross-sectional view of a gas turbine engine. 
         FIG. 2  is a front view of a pressure side of a composite fan blade having a composite airfoil with a locally reinforced tip region. 
         FIG. 3  is a cross-sectional view of the composite airfoil of  FIG. 2  taken along line  3 - 3 . 
         FIG. 4  is an exploded schematic view of a lay-up for the pressure side of the composite airfoil of  FIGS. 2 and 3  having the locally reinforced tip region. 
         FIG. 5  is an exploded schematic view of an alternative lay-up for the pressure side of the composite airfoil of  FIGS. 2 and 3  having the locally reinforced tip region. 
         FIG. 6  is an enlarged cross-sectional view of the composite airfoil of  FIGS. 2 and 3  having a core with a recess. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  is a cross-sectional view of gas turbine engine  10 , which includes turbofan  12 , compressor section  14 , combustion section  16  and turbine section  18 . Compressor section  14  includes low-pressure compressor  20  and high-pressure compressor  22 . Air is taken in through fan  12 . Fan  12  spins and takes in a large amount of inlet air. A portion of the inlet air is directed to compressor section  14  where it is compressed by a series of rotating blades and vanes. The compressed air is mixed with fuel, and then ignited in combustor section  16 . The combustion exhaust is directed to turbine section  18 . Blades and vanes in turbine section  18  extract kinetic energy from the exhaust to turn shaft  24  and provide power output for engine  10 . 
     The portion of inlet air which is taken in through fan  12  and not directed through compressor section  14  is bypass air. Bypass air is directed through bypass duct  26  by guide vanes  28 . Then the bypass air flows through opening  30  to cool combustor section  16 , high pressure combustor  22  and turbine section  18 . 
     Turbofan  12  comprises a plurality of composite blades, such as composite blade  32  shown in  FIG. 2 . Composite blade  32  includes composite airfoil  34  (having leading edge  36 , trailing edge  38 , suction side  40  (not shown), pressure side  42 , tip region  44 , intermediate region  46 , root region  48 , local reinforcement laminate region  50 , root  52  and tip  54 ), protective tip  56 , protective leading edge  58  and longitudinal axis  60 . Root  52  is illustrated as a dovetail root. However, root  52  can have any configuration. Longitudinal axis  60  extends from root region  48  to tip region  44 . 
     Composite airfoil  34  extends from root  52 . The span of composite airfoil  34  is generally defined along longitudinal axis  60 . Root region  48  of composite airfoil  34  is proximate root  52 , tip region  44  is proximate tip  54  and opposite root region  48 , and intermediate region  46  is between root region  48  and tip region  44 . In one example, tip region  44  extends between about 80% of the span-wise extension of composite blade  32  (as measured from root  52  to tip  54 ) and tip  54 , such that tip region  44  has a length equal to about 20% of the span-wise extension of blade  32 . 
     Local reinforcement laminate region  50  is located at tip region  44  of composite airfoil  34 . Local reinforcement laminate section  50  locally reinforces tip region  44  of composite airfoil  34 . Local reinforcement laminate section  50  is limited to tip region  44  and does not extend to root region  48 . In one example, local reinforcement laminate section  50  extends less than or equal to about 20% of the span-wise extension of airfoil  34 . 
     Local reinforcement laminate region  50  comprises at least one filament reinforced ply configured to increase the chordwise stiffness of tip region  44 . For example, the composition or the fiber orientation of the ply of local reinforcement laminate region  50  can be configured to increase the chordwise stiffness of tip region  44 . As described further below, local reinforcement laminate region  50  reduces or eliminates blade flutter. 
     Protective tip  56  is located along tip  54  and protective leading edge  58  is located along leading edge  36  of composite airfoil  34 . Protective tip  56  and protective leading edge  58  protect composite airfoil  34  from damage caused by, for example, bird strikes. Protective tip  56  and protective leading edge  58  also protect composite airfoil  34  from erosion caused by sand, pebbles and other abrasive materials ingested by the turbine during operation. In one example, protective tip  56  and protective leading edge  58  are formed of titanium. Typically, protective tip  56  and protective leading edge  58  are attached to composite airfoil  34  after composite airfoil  34  has been cured and shaped. 
       FIG. 3  is a cross-sectional view of composite airfoil  34  taken along line  3 - 3  of  FIG. 2 . As illustrated in  FIG. 3 , composite airfoil  34  has a sandwich configuration and includes woven core  62  and filament reinforced airfoil laminations or plies  64 . Woven core  62  is located at the center of composite airfoil  34  and extends along longitudinal axis  60  between root region  48  to tip region  44 . Woven core  62  is a three-dimensional woven core containing, for example, carbon fiber. 
     Airfoil plies  64  are located on either side of woven core  62 . Airfoil plies  64  are two-dimensional fabric skins. Elongated fibers extend through airfoil plies  64  at specified orientations and give airfoil plies  64  strength. Airfoil plies  64  vary in shape, size and fiber orientation as described further below. Airfoil plies  64  can be a dry fabric that is combined with a resin in a suitable mold and cured to form composite airfoil  34 . Alternatively, airfoil plies  64  can be preimpregnated uncured composites, (i.e. “prepregs”) in which fibers and a resin are combined with a suitable curing. 
     Turbofan blade designs are primarily driven by three factors: efficiency, protection against bird strike impacts and reducing blade flutter. As described above, turbofan  12  can ingest foreign objects ranging in size from a large bird to hail. Such objects can cause foreign object damage (FOD). Composite fan blades are designed to protect against bird strike impacts and prevent damage to engine  10 . In composite airfoil  34 , woven core  62  absorbs damage due to bird strikes, and airfoil plies  64  provide additional in-plane strength, particularly at root region  48 . Composite airfoil  34  is designed to have reduced or eliminated blade flutter. Blade flutter is characterized by the flapping or vibrating of tip region  44  of composite fan blade  32 . Blade flutter is an aerodynamic phenomenon that is dependent on both the aerodynamic and the structural characteristics of the composite fan blade  32 . Locally reinforcing tip region  44  of composite airfoil  34  with local reinforcement laminate region  50  enables composite fan blade  32  to be tuned. By adjusting the stiffness of composite airfoil  34  along the chordwise axis (i.e. the chordwise stiffness) using local reinforcement laminate region  50 , blade flutter can be reduced or eliminated. The chordwise axis is perpendicular to longitudinal or spanwise axis  60 . The chordwise axis spans between leading edge  36  and trailing edge  38 . 
     Composite airfoil  34  is formed by stacking airfoil plies  64  on woven core  62 . Airfoil plies  64  are stacked in a mold on either side of woven core  62  according to a ply lay-up. Typically the ply lay-up on the pressure side of woven core  62  is a mirror image of the ply lay-up on the suction side of woven core  62 . Once all airfoil plies  64  are properly stacked, the mold is closed, resin is added and the resin is cured to produce composite airfoil  34 . After curing, material can be removed from root region  48  of composite airfoil  34  to further shape root region  48 , and protective tip  56  and protective leading edge  58  (shown in  FIG. 2 ) can be attached to composite airfoil  34 . In an alternative example, airfoil plies  64  contain resin so that resin is not directly added to airfoil plies  64  after stacking them in the mold. 
       FIG. 4  is an exploded schematic view of ply lay-up  68  having locally reinforced region  50  formed by replacing tip region  44  of select airfoil plies  64  with reinforcement plies. Ply lay-up  68  is for pressure side  42  of composite airfoil  34  and comprises filament reinforced airfoil plies  64 A- 64 O and filament reinforced root plies  70 A- 70 O. Airfoil plies  64 A- 64 O (referred to generally as airfoil plies  64 ) form pressure side of composite airfoil  34  of  FIGS. 2 and 3 . Airfoil ply  64 A is the outermost ply on pressure side  42 . Airfoil ply  64 O is the innermost ply and is adjacent woven core  62  (not shown in  FIG. 4 ). Ply lay-up  68  is the lay-up for pressure side plies  64  located between woven core  62  and pressure side  42  of composite airfoil  34 . The lay-up for plies  64  located between woven core  62  and suction side  40  is a minor image of ply lay-up  68 . 
     Airfoil ply  64 B is a locally reinforced ply that comprises two pieces: primary ply  72 B and reinforcement ply  74 B. Primary ply  72 B extends between root region  48  and a location within or proximate to tip region  44 . Reinforcement ply  74 B is aligned with and extends from the end of primary ply  72 B. Reinforcement ply  74 B extends along the longitudinal axis between the end of primary ply  72 B and a location within tip region  44 . Reinforcement ply  74 B may not extend to tip  54 . 
     Reinforcement ply  74 B has a different composition, a different fiber orientation or a different composition and a different fiber orientation than primary ply  72 B. For example, reinforcement ply  74 B can have a 90° fiber orientation and primary ply  72 B can have a 0° fiber orientation. Reinforcement ply  74 B is configured to increase the chordwise stiffness of tip region  44  of composite airfoil  34 . In one example, reinforcement ply  74 B and primary ply  72 B have approximately the same thickness so that when stacked in ply lay-up  68 , no tooling changes are required and composite airfoil  32  has the same geometry as a composite airfoil without reinforcement ply  74 B. When reinforcement ply  74 B has the same thickness as primary ply  72 B, reinforcement ply  74 B does not add additional thickness and an existing mold can be used to produce composite airfoil  34  having an increased chordwise stiffness. Alternatively, woven core  62  can be configured to compensate for a difference in thickness between reinforcement ply  74 B and primary ply  72 B. For example, as described further below, woven core  62  can be formed with a recess at tip region  44  having the same shape and size as additional thickness created by local reinforcement laminate region  50 . 
     Plies  64 D,  64 G and  64 I have configurations similar to ply  64 B. Plies  64 B,  64 D,  64 G and  64 I are locally reinforced plies formed from primary plies and reinforcement plies. Together reinforcement plies  74 B,  74 D,  74 G and  74 I form local reinforced region  50  at tip region  44  of composite airfoil  34 . 
     Root plies  70 A- 70 O (referred to generally as root plies  70 ) are inserted between sections of airfoil plies  64  and form a portion of root region  48  of composite airfoil  34 . Root plies  70  extend between root region  48  and intermediate region  46 . Root plies  70  do not extend into tip region  44 . Root plies  70  provide strength and bending stiffness at root region  48  which enables composite blade  32  to withstand aerodynamic loads and loads generated by bird strikes. 
     Airfoil plies  64  and root plies  70  can be formed from the same material or from different materials. For example, airfoil plies  64  can be formed from a woven fabric or a uniweave material, and root plies  70  can be formed from a uniweave material. In a woven fabric, half of the fibers are orientated in a first direction and the other half of the fibers are oriented 90° to the first direction. For example, half of the fibers of a 0/90° woven fabric are oriented along the longitudinal axis and the other half of the fibers are oriented along the chordwise axis, perpendicular to the longitudinal axis. Similarly, half of the fibers of a +/−45° woven fabric are oriented at +45° from the longitudinal axis and the other half of the fibers are oriented at −45° from the longitudinal axis. The woven fabric can be a carbon woven fabric, such as a carbon woven fabric containing IM7 fibers, to which resin is added to form a composite. In one example, the woven fabric is a 5 hardness satin (5HS) material. Alternatively, the woven fabric can be a prepreg. In a prepreg material, the fibers, resin, and a suitable curing agent are combined. Further, the prepreg material can be a hybrid prepreg which contains two different types of fibers and an epoxy. Example prepreg hybrids include hybrids containing an epoxy and two different types of carbon fibers, such as low modulus carbon fibers (modulus of elasticity below about 200 giga-Pascals (GPa)), standard modulus carbon fibers (modulus of elasticity between about 200 GPa and about 250 GPa), intermediate modulus carbon fibers (modulus of elasticity between about 250 GPa and about 325 GPa) and high modulus carbon fibers (modulus of elasticity greater than about 325 GPa). In one example, the prepreg hybrid is a standard modulus carbon fiber/high modulus carbon fiber/epoxy hybrid. Example prepreg hybrids also include carbon fibers/boron fibers/epoxy hybrid prepregs. 
     In contrast to woven materials, a uniweave material has about 98% of its fibers oriented along the longitudinal axis of airfoil  34 . A small number of fibers extend perpendicular to the longitudinal axis and stitch the uniweave material together. 
     The fiber orientation affects the strength of the material. For example, a composite formed of a 0/90° 5HS woven fabric has a modulus of approximately 75 giga-Pascals (GPa) (11 million pounds per square inch (msi)) in both the 0° and 90° directions, where 0° represents the represents the longitudinal axis (span direction) of airfoil  34 . In comparison, a composite formed of a 0° uniweave material comprising the same fibers has a modulus of approximately 165 GPa (24 msi) in the 0° direction and approximately 9.6 GPa (1.4 msi) in the 90° direction. 
     In  FIG. 4 , tip region  44  of four pressure side airfoil plies, airfoil plies  64 B,  64 D,  64 G and  64 I, include reinforcement plies  74 B,  74 D,  74 G and  74 I to reinforce tip region  44 . Airfoil plies  64 B,  64 D,  64 G and  64 I are locally reinforced plies while airfoil plies  64 A,  64 C,  64 E,  64 F,  64 H and  64 J- 64 O are non-locally reinforced plies. In one example, airfoil plies  64 A,  64 F,  64 J,  64 L and  64 N are 0/90° 5HS woven material; airfoil plies  64 C,  64 K and  64 M and root plies  70  are 0° uniweave material; and airfoil plies  64 E,  64 H and  64 O are +/−45° 5HS woven material. Airfoil plies  64 B,  64 D,  64 G and  64 I comprise primary plies  72 B,  72 D,  72 G and  72 I, respectively, at root region  48  and reinforcement plies  74 B,  74 D,  74 G and  74 I, respectively, at tip region  44 . Airfoil plies  64 B,  64 D,  64 G and  64 I have a different material at root region  48  than at tip region  44 . Primary plies  72 B,  72 D,  72 G and  72 I are formed from 0/90° 5HS woven material, and reinforcement plies  74 B,  74 D,  74 G and  74 I are formed from 90° uniweave. Root plies  70  are formed of 0° uniweave material to provide stiffness along the longitudinal axis at root region  48 . At mid-chord at tip  54 , airfoil plies  64  each have a thickness of about 0.26 millimeters (0.01 inches) and woven core  62  (not shown) has a thickness of about 2.31 millimeters (0.09 inches). The plies on the concave or suction side of woven core  62  have a similar configuration, and airfoil  34  has a total thickness of about 10.2 millimeters (0.4 inches). The flexural stiffness of composite airfoil  34  along longitudinal axis  60  (the spanwise stiffness) is about 64.1 GPa (9.3 msi) and the flexural stiffness of composite airfoil  34  in the direction perpendicular to longitudinal axis  60  (the chordwise stiffness) is about 92.3 GPa (13.4 msi), where the flexural stiffness is the flexural stiffness at mid-chord of the tip region and was determined using finite element modeling software. 
     In comparison, a composite airfoil having a layup similar to layup  68  of  FIG. 4  except having single piece airfoil plies  64 B,  64 D,  64 G and  64 I, such that airfoil plies  64 B,  64 D,  64 G and  64 I are formed entirely from 0/90° 5HS woven material, has a spanwise flexural stiffness of about 88.3 GPa (12.8 msi) and a chordwise flexural stiffness of about 61.0 GPa (8.9 msi), where the flexural stiffness is the flexural stiffness at mid-chord of the tip region and was determined using finite element modeling software. Locally reinforcing tip region  44  by replacing a portion airfoil plies  64 B,  64 D,  64 G and  64 I with local reinforcement plies  74 B,  74 D,  74 G and  74 I results in a 27% decrease in the spanwise flexural stiffness of airfoil  34  and a 51% increase in the chordwise flexural stiffness. That is, local reinforcement laminate region  50  increases the chordwise flexural stiffness of composite airfoil  34  compared to a composite airfoil not having local reinforcement lamination region  50  and having airfoil plies  64  having a uniform composition from root to tip. 
     Previous fan blades were formed from a metal, such as titanium. Metals are typically isotropic in nature so that the stiffness properties are generally the same in every direction. In contrast, the stiffness properties of a composite material can differ greatly depending on the orientation of the fibers. The anisotropic nature of composites allows airfoil  34  to be designed with different flexural stiffnesses in different directions based on the fiber orientation, quantity of plies, stacking sequence of plies and fiber stiffness. The tensile stiffness of airfoil  34  can also be controlled. Tensile strength depends on the fiber orientation, quantity of plies and fiber stiffness. Tensile stiffness is not affected by the stacking sequence. 
     Locally reinforcing tip region  44  with reinforcement plies  74 B,  74 D,  74 G and  74 I enables the chordwise stiffness of tip region  44  to be increased to reduce blade flutter while the spanwise stiffness of root region  48  is maintained to reduce damage from bird strikes. Further, by replacing a portion of plies  64 B,  64 D,  64 G and  64 I with reinforcement plies  74 B,  74 D,  74 G and  74 I having about the same thickness as primary plies  72 B,  72 D,  72 G and  72 I, the geometry of composite airfoil  34  is unchanged and the same mold for stacking and curing can be used without a tooling change to produce composite airfoil  34  with reinforced region  50  and a composite airfoil without reinforced region  50 . 
     Adjustments of the stiffness of tip region  44  to reduce blade flutter can be based on finite element analysis of composite airfoil  34 . With a given blade geometry, blade flutter is dependent on the stiffness and density of composite blade  32 . Finite element analysis is used to determine the tip region stiffness that reduces blade flutter at specific frequency and mode ranges. Based on this stiffness, the number, composition and position of reinforcement plies  74  are determined. Local reinforcement of tip region  44  using reinforcement plies  74 B,  74 D,  74 G and  74 I provides an additional factor that can be adjusted to tune composite blade  32  and reduce or eliminate blade flutter. 
     Reinforcement plies  74  and primary plies  72  are separate plies that have different compositions, different fiber orientations or different compositions and different fiber orientations. In one example, reinforcement plies  74  are formed from a 90° uniweave boron/carbon hybrid material, and primary plies  72  are formed from a 0° uniweave carbon material. In  FIG. 4 , reinforcement plies  74  extend from primary plies  72  and are only located in tip region  44 . Together primary plies  72  and reinforcement plies  74  form a locally reinforced airfoil ply. During production of composite airfoil  34 , airfoil plies  64  and root plies  70  are stacked in a mold on either side of woven core  64  in an order specified in a lay-up schematic. Ply lay-up  68  shows the lay-up for airfoil plies  64  on pressure side  42  of composite airfoil  34 . The lay-up for airfoil plies  64  on suction side  40  is a minor image about the centerplane of ply lay-up  68 . After airfoil plies  64 , primary plies  72 , reinforcement plies  74  and root plies  70  are aligned in the lay-up, the mold is closed, resin is added if necessary and composite airfoil  34  is cured according to manufacture&#39;s instructions. For airfoil plies  64  comprising reinforcement plies  74  and primary plies  72 , reinforcement plies  74  and primary plies  72  can be stacked as separate plies and the resin of composite airfoil  34  will bind the plies together to form composite airfoil  34 . 
       FIG. 5  is an exploded schematic view of an alternative example ply lay-up  76  having locally reinforced laminate region  50  formed by adding reinforcement plies  74  at tip region  44  of select airfoil plies  64 .  FIG. 5  is similar to ply lay-up  68  of  FIG. 4 , except that tip region  44  of select airfoil plies  64  are not removed and replaced with reinforcement plies  74 . In lay-up  76 , all airfoil plies  64 A- 64 O (referred to generally as airfoil plies  64 ) extend to tip region  44 , and reinforcement plies  74 B,  74 D,  74 G and  74 I (referred to generally as reinforcement plies  74 ) are positioned at tip region  44  between select airfoil plies  64 . Reinforcement plies  74  each have leading edge  73  and trailing edge  75 . In one example, leading edge  73  and trailing edge  75  of reinforcement ply  74 B have about the same shape as leading edge  36  and trailing edge  38  of either airfoil ply  64 A or  64 B, which reinforcement ply  74 B is positioned between. In another example, leading edge  73  and trailing edge  75  of reinforcement ply  74 B have the same shape as leading edge  36  and trailing edge  38  of airfoil ply  64 B. 
     In layup  76 , woven core  62  (shown in  FIG. 6 ) is formed with a recess at tip region  44  corresponding to the size and shape of reinforcement plies  74 . The recess in woven core  62  accommodates the additional thickness of reinforcement plies  74  so that composite airfoil  34  has the same geometry as an airfoil without reinforcement plies  74  and no tooling change is necessary. In one example, airfoil plies  64 A,  64 B,  64 D,  64 F,  64 G,  64 I,  64 J,  64 L and  64 N are formed of 5 HS 0/90° woven fabric; airfoil plies  64 E,  64 H and  64 O are formed of 5HS +/−45° woven fabric; airfoil plies  64 C,  64 K and  64 M and root plies  70 A- 70 O are formed of 0° uniweave material; and reinforcement plies  74 B,  74 D,  74 G and  74 I are formed of 90° uniweave material. 
       FIG. 6  is an enlarged cross-sectional view of composite airfoil  34   b  having recessed core  62   b  taken along the longitudinal axis of composite airfoil  34   b.  Airfoil plies  64  are positioned on either side of recessed core  62   b.  For clarity, each individual airfoil ply  64  is not shown. Woven core  62   b  includes recess  80 , tip region  82 , intermediate region  84 , pressure side  86  and suction side  88 . Woven core  62   b  is a three-dimensional woven structure. In one example, woven core  62   b  is formed of woven carbon fibers. Tip region  82  of woven core  62   b  is proximate tip region  44  of airfoil  34   b  and intermediate region  84  of woven core  62   b  is proximate intermediate region  46  of airfoil  34   b.  Recess  80  is formed at tip region  84  of core  62   b  on pressure side  86  and suction side  88 . 
     Airfoil plies  64  are stacked on pressure side  86  of woven core  62   b  to form pressure side  42  of composite airfoil  34   b,  and airfoil plies  64  are stacked on suction side  86  of woven core  62   b  to form suction side  40  of composite airfoil  34   b.  As described above, reinforcement plies  74  can be inserted at tip region  44  between two adjacent airfoil plies  64  (see  FIG. 5 ). Without recess  80 , inserting reinforcement plies  74  in layup  76  would increase the thickness of composite airfoil  34   b  at tip region  44  and would require retooling of the composite blade mold. To eliminate the necessity to retool, recess  80  is configured to compensate for the additional thickness of airfoil  34   b  caused by reinforcement plies  74 . Recess  80  also enables composite airfoil  34  having locally reinforced laminate region  50  to have the same geometry as a composite airfoil without locally reinforced laminate region  50 . 
     Recess  80  is a void formed in tip region  82  of woven core  62   b.  Recess  80  can be a stair-stepped configuration such that multiple reinforcement plies  74  do not end at the same spanwise location. In one example, recess  80  is formed in woven core  62   b  when woven core  62   b  is fabricated or woven. When reinforcement plies  74  are positioned in lay-up  76 , reinforcement plies  74  align with recess  80 . Recess  80  is configured to have the same height, width and thickness as reinforcement plies  74 . In this way, the additional thickness created by reinforcement plies  74  extends into woven core  62   b  and does not extend from the outer surface of airfoil  34   b.  Recess  80  enables reinforcement plies  74  to be added to airfoil  34   b  without changing the profile of the resulting composite airfoil  34   b.    
     Recess  80  can be used in a similar manner to compensate for additional thickness due to reinforcement plies  74  in any type of ply lay-up. For example, woven core  62   b  having recess  80  can also be used in lay-up  68  when reinforcement plies  74  are thicker than airfoil plies  64 . In such a case, recess  80  is sized to compensate for the difference in thickness between reinforcement plies  74  and airfoil plies  64  so that the addition of reinforcement plies  74  does not change the profile of composite airfoil  34   b.    
     The vibration effects of blade flutter are driven by the stiffness and geometry of composite fan blade  32 . By locally changing the lay-up of composite fan blade  32  at tip region  44 , flutter can be reduced or eliminated. In the lay-ups presented in  FIGS. 4 and 5 , reinforcement plies  74  locally reinforce tip region  44  and form local reinforcement laminate region  50 . Reinforcement plies  74  are used to adjust the chordwise stiffness of tip region  44 . As discussed above, chordwise stiffness is affected by the orientation of the fibers, the quantity of plies, the stacking sequence of the plies and the fiber stiffness. Reinforcement plies  74  provide an additional factor that can be adjusted to optimize composite fan blade  32 . 
     Reinforcement plies  74  also allow tip region  44  to be tuned while not affecting the stiffness of root portion  48 . This allows previous optimizations made to root portion  48 , such as improved protection against bird strike impacts, to be maintained. Further, the methods of locally reinforcing tip region  44  presented in  FIGS. 4 and 5  maintain the geometry of composite blade  32  so that tool changes are not necessary in order to add reinforcement plies  74  to the layup. 
     While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. For example, although four reinforcement plies  74  were used in local reinforcement laminate region  50 , local reinforcement laminate region  50  can comprise any number of reinforcement plies  74  such that local reinforcement laminate region  50  increases the chordwise flexural stiffness and chordwise flexural modulus of composite airfoil  34  compared to an airfoil not containing local reinforcement lamination region  50  and having plies  64  with uniform compositions from root region  48  to tip region  44 . Additionally, reinforcement plies  74  can be positioned at any location in reinforcement tip region  44  and are not limited to the locations disclosed. Further, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.