Patent Publication Number: US-10760428-B2

Title: Frangible gas turbine engine airfoil

Description:
FIELD 
     The present subject matter relates generally to airfoils, and more particularly, to frangible airfoils for gas turbine engines. 
     BACKGROUND 
     Airfoils used in aircraft engines, such as fan blades of a gas turbine engine, can be susceptible to extreme loading events. For instance, a fan blade might strike a bird that is ingested into the engine, or a blade-out occurrence may arise wherein one of the fan blades is severed from a rotor disk. If the impact is large enough, a fan blade may break apart into one or more shards before traveling downstream through the engine. 
     Gas turbine engines, such as turbofans, generally include fan cases surrounding a fan assembly including the fan blades. The fan cases are generally configured to withstand an impact of the fan blades due to adverse engine conditions resulting in a failure mode, such as foreign object damage, hard rubs due to excessive or extreme unbalance or fan rotor oscillations, or fan blade liberation. However, such airfoil configurations generally increase the weight of the fan case, thereby increasing the weight of the engine and aircraft and reducing performance and efficiency. 
     Known fan cases generally include frangible structures, such as honeycombs or trench-filler material, configured to mitigate load transfer to and through the fan case. However, this approach is generally costly. Furthermore, this approach may result in larger, heavier, less efficient fan cases. Still further, this approach may not address issues relating to fan rotor unbalance following deformation or liberation of one or several airfoils such as fan blades. 
     As such, there is a need for an airfoil that enables a controlled and consistent failure mode of the airfoil that may enable reducing a cost, weight, and load transfer to a surrounding casing. 
     BRIEF DESCRIPTION 
     Aspects and advantages will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     In one aspect, the present subject matter is directed to an airfoil defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge. The airfoil includes a leading edge sheath coupled to the leading edge of the airfoil. The leading edge sheath includes a flange extending from a frangible line toward the tip along the span and defining a first width along the chord at each point along the span. The leading edge sheath further including a base extending from the frangible line at least partially along the span to the root and defining a second width along the chord at a point along the span of the frangible line such that the second width is greater than the first width. 
     In one embodiment, the airfoil may define a frangible airfoil portion extending between the frangible line and the tip along the span. In one particular embodiment, the airfoil may be a fan blade of a gas turbine engine. In one embodiment, the leading edge sheath and airfoil each may include at least one of a metal, metal alloy, or composite. 
     In another embodiment, the leading edge sheath may extend along the entire span. In one embodiment, the flange may extend from the frangible line along the span toward the tip to a point along the span within 5% of the span from the tip. In a further embodiment, the flange may extend along at least 5% of the span but less than 25% of the span. In other embodiments, the flange may extend along at least 10% of the span but less than 20% of the span. In a further embodiment, the first width may extend along 10% or less of the chord at each point along the span. In such an embodiment, the second width may extend along at least 10% of the chord but less than 90% of the chord at the point along the span of the frangible line. 
     In one embodiment, the airfoil may further define a pressure side and a suction side. The base may extend along the second width on both the pressure and suctions side equally. In such an embodiment, the flange may extend along the first width on both the pressure and suction side equally. In another embodiment, the base may extend along the second width on at least one of the pressure or suction sides and less than the second width on the other of the pressure or suction side. In such an embodiment, the flange may extend along the first width on at least one of the pressure or suction sides and less than the first width on the other of the pressure or suction side. 
     In another aspect, the present subject matter is directed to a gas turbine engine defining a central axis. The gas turbine engine includes an engine shaft extending along the central axis, a compressor attached to the engine shaft and extending radially about the central axis, a combustor positioned downstream of the compressor to receive a compressed fluid therefrom, a turbine mounted on the engine shaft downstream of the combustor to provide a rotational force to the compressor, and a plurality of airfoils operably connected to the engine shaft. Each of the plurality of airfoils defines a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge. Each airfoil includes a leading edge sheath coupled to the leading edge of the airfoil. The leading edge sheath includes a flange extending between from a frangible line toward the tip along the span and defining a first width along the chord at each point along the span. The leading edge sheath further includes a base extending from the frangible line at least partially along the span to the root and defining a second width along the chord at a point along the span of the frangible line such that the second width is greater than the first width. 
     In one embodiment, the gas turbine engine may further include a fan section including the plurality of airfoils configured as fan blades. It should be further understood that the gas turbine engine may further include any of the additional features as described herein. 
     These and other features, aspects and advantages will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain certain principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended FIGS., in which: 
         FIG. 1  illustrates a cross-sectional view of one embodiment of a gas turbine engine that may be utilized within an aircraft in accordance with aspects of the present subject matter, particularly illustrating the gas turbine engine configured as a high-bypass turbofan jet engine; 
         FIG. 2  illustrates a cross-sectional view of the fan section of  FIG. 1  in accordance with aspects of the present subject matter, particularly illustrating a fan blade of the fan section; 
         FIG. 3  illustrates a fan blade of the fan section of  FIGS. 1 and 2  in accordance with aspects of the present subject matter, particularly illustrating a leading edge sheath; 
         FIG. 4  illustrates one embodiment of the leading edge sheath in accordance with aspects of the present subject matter, particularly illustrating a base and flange of the leading edge sheath; and 
         FIG. 5  illustrates another embodiment of the leading edge sheath in accordance with aspects of the present subject matter, particularly illustrating a non-symmetrical leading edge sheath. 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention. 
     DETAILED DESCRIPTION 
     Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. 
     As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. 
     The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. 
     The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein. 
     The terms “communicate,” “communicating,” “communicative,” and the like refer to both direct communication as well as indirect communication such as through a memory system or another intermediary system. 
     A frangible airfoil for gas turbine engines is generally provided. The airfoil may include a leading edge sheath coupled to a leading edge of the airfoil. The leading edge sheath may include a flange positioned toward a tip of the airfoil and a base positioned toward a root of the airfoil. The base may extend farther along a chord of the airfoil than the flange. Further, the flange and base may define a notch and frangible line therebetween. A frangible airfoil portion of the airfoil may be positioned radially outward from the frangible line and include a reduced bending stiffness such that the frangible airfoil portion may break-off or bend during a failure mode of the airfoil. For example, the embodiments generally shown and described herein may enable a controlled and consistent failure of the airfoil, such as a fan blade, following a failure event, such as a hard rub against a surrounding fan case. The embodiments generally described herein enable the airfoil to deform or detach at a desired span of the airfoil to mitigate load transfer to a surrounding casing. The embodiments generally provided herein may further enable the airfoil to deform or detach such that excessive or extreme unbalance of the fan rotor may be reduced following a failure event, such as airfoil liberation, foreign object damage (e.g., bird strikes, icing, etc.), or loss of lube or damper to a bearing assembly. 
     Referring now to the drawings,  FIG. 1  illustrates a cross-sectional view of one embodiment of a gas turbine engine  10  that may be utilized within an aircraft in accordance with aspects of the present subject matter. More particularly, for the embodiment of  FIG. 1 , the gas turbine engine is a high-bypass turbofan jet engine, with the gas turbine engine  10  being shown having a longitudinal or axial centerline axis  12  extending therethrough along an axial direction A for reference purposes. The gas turbine engine  10  further defines a radial direction R extended from the centerline  12 . Although an exemplary turbofan embodiment is shown, it is anticipated that the present disclosure can be equally applicable to turbomachinery in general, such as an open rotor, a turboshaft, turbojet, or a turboprop configuration, including marine and industrial turbine engines and auxiliary power units. 
     In general, the gas turbine engine  10  includes a core gas turbine engine (indicated generally by reference character  14 ) and a fan section  16  positioned upstream thereof. The core engine  14  generally includes a substantially tubular outer casing  18  that defines an annular inlet  20 . In addition, the outer casing  18  may further enclose and support a low pressure (LP) compressor  22  for increasing the pressure of the air that enters the core engine  14  to a first pressure level. A multi-stage, axial-flow high pressure (HP) compressor  24  may then receive the pressurized air from the LP compressor  22  and further increase the pressure of such air. The pressurized air exiting the HP compressor  24  may then flow to a combustor  26  within which fuel is injected into the flow of pressurized air, with the resulting mixture being combusted within the combustor  26 . The high energy combustion products are directed from the combustor  26  along the hot gas path of the gas turbine engine  10  to a high pressure (HP) turbine  28  for driving the HP compressor  24  via a high pressure (HP) shaft or spool  30 , and then to a low pressure (LP) turbine  32  for driving the LP compressor  22  and fan section  16  via a low pressure (LP) drive shaft or spool  34  that is generally coaxial with HP shaft  30 . After driving each of turbines  28  and  32 , the combustion products may be expelled from the core engine  14  via an exhaust nozzle  36  to provide propulsive jet thrust. 
     Additionally, as shown in  FIGS. 1 and 2 , the fan section  16  of the gas turbine engine  10  generally includes a rotatable, axial-flow fan rotor  38  that configured to be surrounded by an annular fan casing  40 . In particular embodiments, the LP shaft  34  may be connected directly to the fan rotor  38  or rotor disk  39 , such as in a direct-drive configuration. In alternative configurations, the LP shaft  34  may be connected to the fan rotor  38  via a speed reduction device  37  such as a reduction gear gearbox in an indirect-drive or geared-drive configuration. Such speed reduction devices may be included between any suitable shafts/spools within the gas turbine engine  10  as desired or required. 
     It should be appreciated by those of ordinary skill in the art that the fan casing  40  may be configured to be supported relative to the core engine  14  by a plurality of substantially radially-extending, circumferentially-spaced outlet guide vanes  42 . As such, the fan casing  40  may enclose the fan rotor  38  and its corresponding fan rotor blades (fan blades  44 ). Moreover, a downstream section  46  of the fan casing  40  may extend over an outer portion of the core engine  14  so as to define a secondary, or by-pass, airflow conduit  48  that provides additional propulsive jet thrust. 
     During operation of the gas turbine engine  10 , it should be appreciated that an initial airflow (indicated by arrow  50 ) may enter the gas turbine engine  10  through an associated inlet  52  of the fan casing  40 . The air flow  50  then passes through the fan blades  44  and splits into a first compressed air flow (indicated by arrow  54 ) that moves through the by-pass conduit  48  and a second compressed air flow (indicated by arrow  56 ) which enters the LP compressor  22 . The pressure of the second compressed air flow  56  is then increased and enters the HP compressor  24  (as indicated by arrow  58 ). After mixing with fuel and being combusted within the combustor  26 , the combustion products  60  exit the combustor  26  and flow through the HP turbine  28 . Thereafter, the combustion products  60  flow through the LP turbine  32  and exit the exhaust nozzle  36  to provide thrust for the gas turbine engine  10 . 
     Referring to  FIGS. 2 and 3 , exemplary airfoil  62  embodiments are provided in the context of a fan blade  44 . Although the illustrated airfoils  62  are shown as part of a fan blade  44 , it is understood that the following discussion of an airfoil  62  may be equally applied to another airfoil embodiment, e.g., a stator vane or rotor blade of a compressor  22 ,  24  and/or turbine  28 ,  32  (see  FIG. 1 ). As shown, each fan blade  44  extends radially outwardly along a span S from an airfoil root  64  to an airfoil tip  66 . A pressure side  68  and a suction side  70  of the airfoil  62  extend from the airfoil&#39;s leading edge  72  to a trailing edge  74  and between the airfoil root  64  and airfoil tip  66  along the span S. Further, it should be recognized that airfoil may define a chord C at each point along the span S between the airfoil root  64  and the airfoil tip  66 . Further, the chord C may vary along the span of the airfoil  62 . For instance, in the depicted embodiment, the chord C increases along the span S toward the airfoil tip  66 . Though, in other embodiments, the chord C may be approximately constant throughout the span S or may decrease from the airfoil root  64  to the airfoil tip  66 . 
     Optionally, each fan blade  44  includes an integral component having an axial dovetail  76  with a pair of opposed pressure faces  78  leading to a transition section  80 . When mounted within the gas turbine engine  10 , as illustrated in  FIG. 2 , the dovetail  76  is disposed in a dovetail slot of the fan rotor disk  39 , thereby attaching the fan blades  44  to the fan rotor  38 . 
     In the depicted embodiment, the airfoil  62  may include a leading edge sheath  82  including a base  84  and flange  86  coupled to the leading edge  72  of the airfoil  62 . For example, as illustrated, the leading edge sheath  82  may extend from the airfoil tip  66  along at least a portion of the span S of the airfoil  62 . For instance, the leading edge sheath  82  may be bonded to and provide protection for the leading edge  72  of the airfoil  62 . It should be recognized that the leading edge sheath  82  may be coupled to the leading edge  72  suing any suitable means, such as by adhesives, tape, welding, and/or mechanical fasteners (e.g., bolts, screws, and rivets). The leading edge sheath  82  may generally strengthen the airfoil  62 , minimize danger to the airfoil  62  (e.g., the fan blade  44 ) during a fan blade out event, and protect the airfoil  62  from foreign object damage. 
     In one embodiment, the leading edge sheath  82  may extend along the entire span S. As such, the leading edge sheath  82  may protect the entire leading edge  72  of the airfoil  62  from the airfoil root  64  to the airfoil tip  66 . In other embodiments, the leading edge sheath  82  may protect only a portion of the leading edge  72  of the airfoil  62 , such as a portion of the leading edge  72  toward the airfoil tip  66 . 
     The leading edge sheath  82  may include the base  84  and the flange  86  meeting at a frangible line  88 . As illustrated, the base  84  and the flange  86  may defining a notch  90  therebetween. More particularly, the position where the flange  86  and the base  84  meet at the notch  90  may define the frangible line  88 . For example, the notch  90  may be defined relative to a difference in a first width  96  of the flange  86  and a second width  100  of the base  84  as described in more detail in regards to  FIGS. 4 and 5 . Further, the frangible line  88  may generally extend along the chord C toward the trailing edge  74 . It should be recognized that the frangible line  88  may generally extend along the chord C at approximately the same point along the span S. In other embodiments, the frangible line  88  may at least partially extend radially inward or outward (e.g., along the span S) as the frangible line  88  extends axially along the chord C toward the trailing edge  74 . In one embodiment, the airfoil  62  may define a residual airfoil portion  92  extending from the root  64  to the frangible line  88  along the span S of the airfoil  62 . In such an embodiment, the airfoil  62  may further define a frangible airfoil portion  94  extending from the airfoil tip  66  to the frangible line  88  along the span S of the airfoil  62 . The frangible airfoil portion  94  may meet the residual airfoil portion  90  at the frangible line  88 . Additionally, the notch  90  may at least partially define the frangible airfoil portion  92 . The frangible airfoil portion  94  may have a reduced overall bending stiffness compared to the residual airfoil portion  92 , as described in more detail in regards to  FIGS. 4 and 5 . 
     Still referring to the exemplary airfoil  62  of  FIG. 3 , the airfoil  62  may be configured to fracture, break, or liberate at approximately the frangible line  88  up to the airfoil tip  66  (e.g., the frangible airfoil portion) following a loading or impact upon the airfoil  62 . For example, the airfoil  62  configured as the fan blade  44  within the fan casing  40  or nacelle of the gas turbine engine  10  ( FIG. 1 ) may be configured to detach, decouple, deform, break, or liberate the frangible airfoil portion  94  of the airfoil  62  above the frangible line  88 . In one non-limiting example, the frangible airfoil portion  94  of the airfoil  62  may be defined as the difference in spanwise dimensions of the frangible airfoil portion  94  and the residual airfoil  92 . For example, the frangible airfoil portion  94  may be defined within approximately 3% to approximately 15% of the total span S from the airfoil tip  66 . 
     During operation of the gas turbine engine  10 , such as following an event generating substantial imbalance in the fan rotor  38  or LP shaft  34 , the frangible airfoil portion  94 , e.g., of the fan blade  44 , as shown and described in various embodiments in  FIGS. 3-5  may be configured to deform or partially or fully detach from the remainder of the airfoil  62 , e.g., along the frangible line  88 . Further, the frangible airfoil portion  94  may detach (e.g., along the frangible line  88 ) while leaving all of or at least a portion of the residual airfoil portion  92 . Events generating substantial unbalance in the fan rotor  38  and/or LP shaft  34  may include, but are not limited to, foreign object damage (e.g., bird strikes, ice ingestion, other debris, etc.) or fan blade  44  detachment. Detaching or decoupling the frangible airfoil portion  94  may reduce undesired unbalance or vibrations as the fan rotor  38  and/or LP shaft  36  continue to rotate. Furthermore, the embodiments of the airfoil  62  generally shown and described in regard to  FIGS. 3-5  may enable a lighter fan casing  44  or nacelle, such as reducing an amount of metal materials or abradable material of the fan casing  40  or nacelle. 
     Referring now to  FIG. 4 , one embodiment of the leading edge sheath  82  is illustrated according to aspects of the present subject matter. Particularly,  FIG. 4  illustrates the base  84  and the flange  86  of the leading edge sheath  82  with the remaining components of the airfoil  62  omitted for clarity. For example, the leading edge sheath  82  may include the flange  86  extending from the frangible line  88  toward the airfoil tip  66  along the span S. It should be appreciated that, in other embodiments, the flange  86  may not extend fully to the airfoil tip  66 , such as shown in  FIG. 5 . For instance, the flange  86  may extend from the frangible line  88  along the span S to a point along the span S within 5% of the span S from the airfoil tip  66 . Further, the flange  86  may define a first width  96  along the chord C at each point along the span S (e.g., with the flange  86 ). In a certain embodiments, the first width  96  may extend along 10% or less of the chord C at each point along the span S. In one embodiment, the flange  86  may define a flange height  98  extending along at least 5% of the span S but less than 25% of the span S. In other embodiments, the flange height  98  of the flange  96  may extend along at least 10% of the span S but less than 20% of the span S. As such, it should be recognize that the notch  90  and/or the frangible airfoil portion  94  may also extend along the flange height  98  between the airfoil tip  66  and the frangible line  88 . Or more particularly, the flange  96  may define both the frangible airfoil portion  94  and the notch  90  such that they each have a height that is the same as or approximately the same as the flange height  98 . 
     The leading edge sheath  82  may further include the base  84  extending from the frangible line  88  at least partially along the span S to the airfoil root  64  ( FIG. 3 ). In embodiments where the leading edge sheath  82  extends along the full span S of the leading edge  72  ( FIG. 3 ), the base  84  may extend the full distance between the frangible line  88  and the airfoil root  64 . Further, the base  84  may define a second width  100  along the chord C at the point along the span S of the frangible line  88  such that the second width  100  is greater than the first width  96 . As described herein, the second width  100  greater than the first width  96  refers to a first width  96  that extends along less of a percentage of the chord C at each point along the span S of the flange  86  than the percentage of the chord C that the second width  100  of the base  84  extends at the point along the span S of the frangible line  88 . In certain embodiments, the second width  100  may extend along at least 10% of the chord C at the point along the span of the frangible line  88  but less than 90% of such chord C. For instance, the base  84  may cover the pressure side  68  and suctions side  70  within proximity of the trailing edge  74 , such as a distance approximately 10% of the chord C away from the trailing edge  74  at the point along the span S of the frangible line  88 . As such, it should be recognized that, in the illustrated embodiment, the base  84  extends farther along the chord C at the point along the span S of the frangible line  88  than the flange  86  extends along the chord C at each point along the span S. 
     In certain embodiments, the base  84  may define a portion of the airfoil  62  with a first overall bending stiffness (e.g., the residual airfoil portion  92 ). Similarly, the flange  66  may define a portion of the airfoil  62  with a second overall bending stiffness (e.g., the frangible airfoil portion  94 ) less than the first overall bending stiffness of the residual airfoil portion  94 . For instance, the notch  90  may reduce an amount of the leading edge sheath  82  above the frangible line  88  along the span S and the chord C. As such, the frangible airfoil portion  94  may have a reduced stiffness allowing the frangible airfoil portion  94  to fracture, break, liberate, decouple, deform, deflect, etc. at the frangible line  88  as described in regards to  FIG. 3  above. 
     Still referring to the embodiment of  FIG. 4 , the leading edge sheath  82  may be symmetrical about the leading edge  72 . For example, in such an embodiment, the flange  86  may extend along the first width  96  on both the pressure and suction side  68 ,  70  ( FIG. 3 ) equally. For instance, the flange  86  may provide protection equally on both sides of the leading edge  72 . Further, the base  84  may extend along the second width  100  on both the pressure and suction sides  68 ,  70  equally. 
     Referring now to  FIG. 5 , another embodiment of the leading edge sheath  82  is illustrated according to aspects of the present subject matter. Particularly,  FIG. 5  illustrates a non-symmetrical leading edge sheath  82 . For instance, the flange  86  may extend along the first width  96  on at least one of the pressure or suction sides  68 ,  70  ( FIG. 3 ) and less than the first width  96  on the other of the pressure or suction side  68 ,  70 . More particularly, the flange  96  may extend along a third width  102  along one of the pressure or suctions sides  68 ,  70  less than the first width  96 . 
     As such, the flange  86  may be non-symmetrical around the leading edge  72  of the airfoil  62 . For instance, the flange  86  of the leading edge sheath  82  may extend along the chord C at each point along the span S farther on the pressure side  68  (e.g., 10%) than on the suction side  70  (e.g., 5%). For example, the leading edge sheath  82  may not extend to the full first width  96  toward direction the airfoil  62  is intended to bend and/or break. In another embodiment, the base  84  may extend along the second width  100  on at least one of the pressure or suction sides  68 ,  70  and less than the second width  100  on the other of the pressure or suction side  68 ,  70 . For instance, the base  84  may extend along a fourth width  104  on one of the sides  68 ,  70  at the point along the span S of the frangible line  88  less than the second width  100 . In certain embodiments, the fourth width  104  may be greater than the first width  96 . Though, in other embodiments, the fourth width  104  may be less than the first width  96  but greater than the third width  102 . As such, the base  84  may be non-symmetrical around the leading edge  72  of the airfoil  62 . 
     It should be recognized that, in certain embodiments, both the flange  86  and the base  84  may extend along the first width  96  and second width  100 , respectively, on one of the pressure or suction side  68 ,  70  and less than the first width  96  and second width  100  on the other side  68 ,  70 . Though, in still further embodiments, the flange  86  may extend less than the first width  96  (the third width  102 ) on one side, such as the suction side  70 , and the base  84  may extend less than the second width  100  (the fourth width  104 ) on the other side, such as the pressure side  68 . For instance, the base  84  may extend farther on the side the frangible airfoil portion  94  is intended to bend and/or break, thereby providing more protection to the airfoil  62  in the event of a failure mode. Similarly, the flange  86  may extend farther on the opposite side the frangible airfoil portion  94  is intended to bend and/or break, thereby providing more protection to the side of the airfoil  62  that may rub against the fan casing  40 , e.g., the pressure side  68 . 
     In one embodiment, the airfoil  62  and/or leading edge sheath  82  may include at least one of a metal, metal alloy, or composite. For instance, the airfoil  62  and/or leading edge sheath  82  may be formed at least partially from a ceramic matrix composite. More particularly, in certain embodiments, the airfoil  62  and leading edge sheath  82  may be formed from one or more ceramic matrix composite prepreg plies. For instance, such prepreg plies forming the leading edge sheath  82  may be wrapped around the leading edge  72  of the airfoil  62  and cured and processed to form the leading edge sheath  82 . In other embodiments, the airfoil  62  and/or leading edge sheath  82  may be formed at least partially from a metal, such as but not limited to, steel, titanium, aluminum, nickel, or alloys of each. For instance, in certain embodiments, the airfoil  62  and/or leading edge sheath  82  may be cast. In one particular embodiment, the airfoil  62  may be formed from a ceramic matric composite while the leading edge sheath  82  may be formed from a metal. Though, it should be recognized that the airfoil  62  and/or leading edge sheath  82  may be formed from multiple materials, such as a combination of metals, metal alloys, and/or composites. For example, the flange  86  may include one material while the base  84  includes another material. Further, in such an embodiment, the flange  86  and base  84  may be bonded at the frangible line  88 . Similarly, the residual airfoil portion  90  may include one material while the frangible airfoil portion  94  includes another material bonded with the residual airfoil portion  90  at the frangible line  88 . It should be recognized that the materials forming the flange  86  and/or frangible airfoil portion  94  may have a reduced stiffness comparted to the materials forming the base  84  and the residual airfoil portion  92 . 
     Composite materials may include, but are not limited to, metal matrix composites (MMCs), polymer matrix composites (PMCs), or ceramic matrix composites (CMCs). Composite materials, such as may be utilized in the airfoil  62  and/or leading edge sheath  82 , generally comprise a fibrous reinforcement material embedded in matrix material, such as polymer, ceramic, or metal material. The reinforcement material serves as a load-bearing constituent of the composite material, while the matrix of a composite material serves to bind the fibers together and act as the medium by which an externally applied stress is transmitted and distributed to the fibers. 
     Exemplary CMC materials may include silicon carbide (SiC), 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&#39;s SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon&#39;s NICALON®, Ube Industries&#39; TYRANNO®, and Dow Corning&#39;s SYLRAIVIIC®), alumina silicates (e.g., Nextel&#39;s 440 and 480), and chopped whiskers and fibers (e.g., Nextel&#39;s 440 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). For example, in certain embodiments, bundles of the fibers, which may include a ceramic refractory material coating, are formed as a reinforced tape, such as a unidirectional reinforced tape. A plurality of the tapes may be laid up together (e.g., as plies) to form a preform component. The bundles of fibers may be impregnated with a slurry composition prior to forming the preform or after formation of the preform. The preform may then undergo thermal processing, such as a cure or burn-out to yield a high char residue in the preform, and subsequent chemical processing, such as melt-infiltration with silicon, to arrive at a component formed of a CMC material having a desired chemical composition. In other embodiments, the CMC material may be formed as, e.g., a carbon fiber cloth rather than as a tape. 
     Similarly, in various embodiments, PMC materials may be fabricated by impregnating a fabric or unidirectional tape with a resin (prepreg), followed by curing. For example, multiple layers of prepreg may be stacked to the proper thickness and orientation for the part, and then the resin may be cured and solidified to render a fiber reinforced composite part. As another example, a die may be utilized to which the uncured layers of prepreg may be stacked to form at least a portion of the composite component. The die may be either a closed configuration (e.g., compression molding) or an open configuration that utilizes vacuum bag forming. For instance, in the open configuration, the die forms one side of the blade (e.g., the pressure side  68  or the suction side  70 ). The PMC material is placed inside of a bag and a vacuum is utilized to hold the PMC material against the die during curing. In still other embodiments, the airfoil  62  and/or leading edge sheath  82  may be at least partially formed via resin transfer molding (RTM), light resin transfer molding (LRTM), vacuum assisted resin transfer molding (VARTM), a forming process (e.g. thermoforming), or similar. 
     Prior to impregnation, the fabric may be referred to as a “dry” fabric and typically comprises a stack of two or more fiber layers (plies). The fiber layers may be formed of a variety of materials, non-limiting examples of which include carbon (e.g., graphite), glass (e.g., fiberglass), polymer (e.g., Kevlar®) fibers, and metal fibers. Fibrous reinforcement materials can be used in the form of relatively short chopped fibers, generally less than two inches in length, and more preferably less than one inch, or long continuous fibers, the latter of which are often used to produce a woven fabric or unidirectional tape. Other embodiments may include other textile forms such as plane weave, twill, or satin. 
     In one embodiment, PMC materials can be produced by dispersing dry fibers into a mold, and then flowing matrix material around the reinforcement fibers. Resins for PMC matrix materials can be generally classified as thermosets or thermoplastics. Thermoplastic resins are generally categorized as polymers that can be repeatedly softened and flowed when heated and hardened when sufficiently cooled due to physical rather than chemical changes. Notable example classes of thermosplastic resins include nylons, thermoplastic polyesters, polyaryletherketones, and polycarbonate resins. Specific examples of high performance thermoplastic resins that have been contemplated for use in aerospace applications include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), and polyphenylene sulfide (PPS). In contrast, once fully cured into a hard rigid solid, thermoset resins do not undergo significant softening when heated but, instead, thermally decompose when sufficiently heated. Notable examples of thermoset resins include epoxy, bismaleimide (BMI), and polyimide resins. 
     This written description uses exemplary embodiments to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.