TURBINE ENGINE WITH COMPOSITE BLADE

Abstract

A blade for a turbine engine comprising a composite core defining a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. The composite core is formed from two materials with different compositions.

Description

BACKGROUND OF THE INVENTION

Turbine engines include rotating components mounted to shafts and surrounded by shrouds and casings that provide structural support and air flow guidance through the machine. This disclosure relates generally to a turbine engine with a fan section and in particular one of the rotating components being a fan blade having a composite core. Fan blades can include a metal leading edge component mounted to the composite core.

The fan blades and other rotating components, including the compressors, and turbines, for example, rotate with a tip of the respective component passing very close to the shrouds or casings. During some events, for example, a bird ingestion into the gas turbine engine, the blades may contact the shrouds or casings. Also, during at least some such events, the blades may be released from gas turbine engines, i.e., “fan blade out” (FBO. Such events typically cause damage to the shrouds or casings of the blades and to other components of gas turbine engines. The damage may also cause the gas turbine engine to operate with a lesser capability, necessitating repair.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, the disclosure relates to an airfoil for a turbine engine, the airfoil comprising a composite core having a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction and having a span-length, a first portion of the composite core extending radially from the root less than 100% of the span-length and formed from a first material, a second portion of the composite core proximate the first portion and defining at least a portion of the tip and formed from a second material, having a lower strain capability than the first material and having a different composition than the first material.

In another aspect, the disclosure relates to a blade for a turbine engine, the blade comprising interior composite core having a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction and having a span-length, a first portion of the composite core extending radially from the root less than 100% of the span-length and formed from a first material, a second portion of the composite core proximate the first portion and defining at least a portion of the tip and formed from a second material having a lower strain capability than the first material and a different composition than the first material.

In yet another aspect, the disclosure relates to a method for forming a blade having composite core and extending radially between a root and a tip defining a span-wise direction and having a span-length, the method comprising forming a first and a second portion of the composite core, extending the first portion radially outward from the root less than 100% of the span-length, extending the second portion radially inward from the tip to meet the first portion, and joining the first portion to the second portion to form a fuse.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic view of a turbine engine assembly including a fan blade.

FIG. 2 is an enlarged perspective view of the fan blade from FIG. 1.

FIG. 3 is a cross-sectional view taken along line along a tip of the fan blade from FIG. 2.

FIG. 4 is a variation of the cross section of FIG. 3 according to another aspect of the disclosure herein.

FIG. 5 is another variation of the cross section of FIG. 3 according to yet another aspect of the disclosure herein.

FIG. 6 is a variation of the fan blade from FIG. 2 with a leading edge strip according to an aspect of the disclosure herein.

FIG. 7 is a leading edge strip for the fan blade of FIG. 6 according to another aspect of the disclosure herein.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

Aspects of the disclosure described herein are directed to a blade of a turbine engine, in particular a fan blade, where the fan blade includes a composite core made of at least two different materials. The two materials have differing compositions where the material closest to a tip of the fan blade is less compliant than the material from which the rest of the composite core is made. A frangible zone of the blade is defined as an area around where a fuse is formed between the two materials. The frangible zone enables the portion of the blade formed from the less compliant material to break easily from the remaining portion during, by way of non-limiting example, an FBO event.

For purposes of illustration, the present disclosure will be described with respect to a fan blade in a turbine engine. While illustrated as a fan blade it should be understood that the disclosure as described herein can have applicability to other blades in the engine as well. Furthermore, it will be understood, that aspects of the disclosure described herein are not so limited and that a blade as described herein can be implemented in engines, including but not limited to turbojet, turboprop, turboshaft, and turbofan engines. Aspects of the disclosure discussed herein may have general applicability within non-aircraft engines having a blade, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the outlet of the engine or being relatively closer to the engine outlet as compared to another component. Additionally, as used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. Furthermore, as used herein, the term “set” or a “set” of elements can be any number of elements, including only one.

All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader's understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of aspects of the disclosure described herein. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary.

FIG. 1 illustrates an exemplary turbine engine assembly 10 having a longitudinal axis defining an engine centerline 12. A fan assembly 14, a nacelle 16, and a turbine engine core 18 can be included in the turbine engine assembly 10.

The turbine engine core 18 includes a low pressure compressor 20, a high pressure compressor 22, a combustor assembly 24, a high pressure turbine 26, and a low pressure turbine 28 arranged in a serial, axial flow relationship.

The fan assembly 14 includes an array of fan blades 30 extending radially outward from a rotor disc 32. The turbine engine assembly 10 has an intake side 34 and an exhaust side 36. The fan blades 30 and low pressure turbine 28 are coupled together with a rotor shaft 38. The fan assembly 14 and engine core 18 are at least partially positioned within the nacelle 16.

In operation, air 40 flows through the fan assembly 14 and a first portion 42 of the airflow is channeled through compressors 20, 22 wherein the first portion 42 of the airflow is further compressed and delivered to the combustor assembly 24. Hot products of combustion (not shown) from the combustor assembly 24 are utilized to drive turbines 26, 28 and thus produce engine thrust. A second portion 44 of the airflow discharged from fan assembly 14 is utilized to bypass around the turbine engine core 18.

FIG. 2 is a schematic view of a fan blade 30 for the fan assembly 14 of the turbine engine 10. The fan assembly 14 can be positioned within the nacelle 16 to form a clearance 45 between each fan blade 30 and the nacelle 16. Each fan blade 30 can have an airfoil cross-sectional shape and extend axially between a leading edge 46 and a trailing edge 48 to define a chord-wise direction and radially between a root 50 and a tip 52 to define a span-wise direction. A span-length (L) is the greatest length measured from the root 50 to the tip 52 in the span-wise direction. The fan blade 30 can include a tip cap 54 defining at least a portion of the fan blade 30 at the tip 52. By way of non-limiting example the tip cap 54 is formed from titanium sheet metal. Alternatively, the tip cap 54 can be formed from any material that facilitates operation of the fan assembly 14 as described herein.

A leading edge strip 56 can define at least a portion of the leading edge 46 of the fan blade 30. A trailing edge strip 58 can define at least a portion of the trailing edge 48 of the fan blade 30. The leading edge strip 56 and the trailing edge strip 58 can be formed from any metallic material that facilitates operation of the fan assembly 14 as described herein, including, but not being limited to, titanium alloys, steel alloys, or nickel alloys. During operation damage due to contact with foreign objects or extreme temperature differentiations can occur. The leading edge strip 56 and the trailing edge strip 58 are provided to protect the leading and trailing edges 46, 48 of the fan blade 30 from such damage during operation.

The fan blade 30 can further include a dovetail pin 60 that facilitates coupling the fan blade 30 to the rotor disc 32. A composite core 62 can extend from the dovetail pin 60 and provide a light weight inner structure to the fan blade 30. The entirety of the composite core 62 is illustrated in phantom and can extend axially between the leading edge 46 and the trailing edge 48 and radially between the root 50 and tip 52. It is contemplated that the composite core 62 defines a large portion of the fan blade 30 and can have a hollow interior or a solid interior. It is further contemplated that the composite core defines the entire fan blade 30.

The composite core 62 includes a first portion 64 formed from a first material and a second portion 66 formed from a second material. The first portion 64 extends from the root 50 towards the tip 52 less than 100% of the span-length (L). A fuse 68 is formed where the first portion 64 meets the second portion 66. The second portion 66 extends radially from the fuse 68 toward the tip 52 and can define at least a portion of the tip 52. It is further contemplated that the second portion 66 can extend in the span-wise direction up to 5% of the span-length (L).

The first and second materials can be composite materials, by way of non-limiting example glass or carbon fiber composites. The first and second materials can be other suitable composite materials with different compositions so long as the first material has a higher strain capability than the second material. Strain capability is in reference to the amount that the material is able to bend or deform in relationship to the materials original structure when impacted or otherwise undergoing an outside force. Strain capability is a ratio of a new length upon undergoing the outside force to an original length and is therefore unit-less. A higher strain capability equates to a higher amount of deformation in an object with respect to the original form of the object. According to an aspect of the disclosure herein, a higher strain capability equates to an object being more compliant, or having more flexibility, than an object with a lower strain capability. Therefore an object with a lower strain capability is more frangible and through deformation tends to break up into fragments rather than deforming. In an aspect of the disclosure herein the strain capability of the second material is less than 85% of the strain capability of the first material. By way of non-limiting example the first material can have a strain capability of 1.6% while the second material has a strain capability of less than 1.3%. It is further contemplated that the strain capability for the first material is greater than 1% while the strain capability of the second material is less than or equal to 1%. In either scenario the second material from which the second portion 66 is formed, has a lower strain capability than the first material from which the first portion 64 is formed.

It is also contemplated that the first material has a higher strain capability and a lower modulus of elasticity than the second material. The modulus of elasticity quantifies the tendency of the first or second portions 64, 66 to deform along an axis when opposing forces are applied along that axis. In other words the modulus of elasticity can quantify the stiffness of a material. An object with a high strain capability and a low modulus of elasticity is a compliant and relatively flexible material.

It is also contemplated, however, that the first material can have an equal modulus of elasticity as the second material and still maintain a higher strain capability. In such a case it would take more force to break the first material than it would to break the second material with a similar modulus of elasticity. By way of non-limiting example, some ceramic fibers can have a similar modulus of elasticity but have lower strain capability than carbon fibers. In one such case the first material would be made of a carbon fiber while the second material would be made from a ceramic fiber. In another non-limiting example the first material can be an intermediate modulus fiber, IM7 or T800 carbon fiber for example, while the second material can be formed from a high modulus of elasticity fiber, for example HM63 or M55J carbon fiber. According to aspects of the disclosure herein the first material from which the first portion 64 is formed is more compliant, or bendable, relative to the second material. More specifically, the second material from which the second portion 66 is formed is more frangible, or brittle, relative to the first material.

During operation contact with foreign objects can cause orbiting of the fan assembly 14 about the rotor shaft 38 which can induce an FBO as described previously herein. The second portion 66, defining at least a portion of the tip 52, can break off due to the discontinuity of composition between the tip 52 and the composite core 62. The fuse 68 and surrounding area define a frangible zone 70 where the break can occur, by way of non-limiting example at the fuse 68 or anywhere immediately surrounding the fuse 68. The difference in strain capability is to ensure an enabled frangible zone 70. A compositional difference between the first and second materials can contribute to enabling the frangible zone 70. This controlled break ensures that small pieces of the fan blade 30, rather than the entire fan blade 30, are ingested by the engine 10 during an FBO. Therefore, in the case of an FBO safe shut down of the engine can occur.

FIG. 3 is an enlarged cross-sectional view of the frangible zone 70 shown along line from FIG. 2. In one aspect of the disclosure herein a first material 74, as described herein, forming the first portion 64 of the composite core 62 is more compliant than a second material 76, as described herein. The first material can be a fibrous composite material with resin. The first material 74 can have a modulus of elasticity between 5 Msi and 20 Msi (34-140 Gpa).

In an aspect of the disclosure discussed herein the second material 76 forming the second portion 66 is a non-compliant material. Non-compliant meaning a stiff material with very little to no elasticity, maintaining a strain capability that is less than the first material 74. The non-compliant material can be a composite fiber, carbon fiber, an epoxy fiber, or a combination of both wherein the second portion 66 is a layered composite. It is contemplated that the modulus of elasticity for the second material 76 is greater than 10 Msi (65 GPa), and preferably greater than 20 Msi (140 Gpa).

In one non-limiting example the first material 74 can be a composite with a modulus of elasticity of 10 Msi (65 Gpa), a relatively low-range stiffness composite, and the second material 76 can be another composite having a modulus of elasticity of 20 Msi (200 GPa), a highly stiff composite. The discontinuity of composition between the tip 52 and the composite core 62 as described herein can be defined as two composites with differing moduli of elasticity.

The fuse 68 as described herein is shown having a generally flat profile, however it is understood that the fuse 68 can have a curved, fragmented, or varying height profile. In this manner, the extent (X) to which the second portion 66 extends in the span-wise direction can vary in and out of the page in the chord-wise direction. As described previously herein, the extent (X) is a length of up to 5% of the span-length (L) (FIG. 2). It is further contemplated that the fuse 68 as described herein can vary in width (W), a dimension perpendicular to the extent (X), depending on the methods used to connect the first portion 64 to the second portion 66. Methods used to connect the first and second portions 64, 66 can include, but are not limited to, brazing, welding, and adhesive bonding.

The tip cap 54 as described herein can be formed on both sides of the composite core 62 as illustrated. It is contemplated that the tip cap 54 is one continuous piece formed to envelop 100% of the composite core 62 at the tip 52, or two separate pieces bonded to the composite core in any suitable method, by way of non-limiting example brazing, welding, and adhesive bonding. The tip cap 54 can be formed on both a pressure side 78 and suction side 80 of the fan blade 30 as illustrated. It is further contemplated that the tip cap 54 extends along at least 50% of the tip 52 in the chord-wise direction.

A method for forming a blade, by way of non-limiting example the fan blade 30 with the composite core 62 as described herein can include forming the first and a second portions 64, 66 of the composite core 62. In forming the first portion 64, extending the first portion 64 radially outward from the root 50 (FIG. 1) up less than 100% of the span-length (L). In forming the second portion 66, extending the second portion 66 radially inward from the tip 52 to meet the first portion 64. Forming the second portion can further include forming the second portion 66 from a stiff non-compliant material as described herein. The method can be consistent with, by way of non-limiting example, commercial manufacturing methods for blades in which a fibrous composite material with resin is utilized for the first portion 64. The manufacturing method can include, by way of non-limiting example autoclave, press mold or RTM (resin transfer molding). The second portion 66 can be by way of non-limiting example, ceramic fiber or carbon fiber, with a lower strain capability than the first portion 64 as described herein.

The first portion 64 is joined to the second portion 66 to form the fuse 68. Joining can include one of a scarfing, interweaving, abutting, or lapping the first portion 64 with the second portion 66. In one non-limiting example the method can include layup and curing the second portion 66, layup and curing the first portion 64, then bonding them together either before or in the same step as bonding the tip cap 54. It is further contemplated that curing can occur in the same step. In another aspect of the disclosure herein, the first and second materials 74, 76 can be weaved as a preformed structure transitioning from a first material 74 to a second material 76 and then using injected resin cure with the tip cap 54. It should be understood that these process are not all inclusive and that other process of bonding the parts as described herein to form the fuse 68 are contemplated. The fuse 68 can therefore be, by way of non-limiting example, one of a scarfed, interwoven, butt, or lap joint. The method can further include extending the tip cap 54 as described herein around at least a portion of the tip 52.

FIG. 4 is an enlarged view of a frangible zone 170 for a composite core 162 according to another aspect of the disclosure discussed herein. The frangible zone 170 is substantially similar to the frangible zone 70, therefore, like parts will be identified with like numerals increased by 100, with it being understood that the description of the like parts of the frangible zone 70 applies to the frangible zone 170 unless otherwise noted. A tip cap 154 can be formed along one of a pressure side 178 or suction side 180, by way of non-limiting example along only the suction side 180 as illustrated.

FIG. 5 is an enlarged view of a frangible zone 270 for a composite core 262 according to another aspect of the disclosure discussed herein. The frangible zone 270 is substantially similar to the frangible zone 70, therefore, like parts will be identified with like numerals increased by 200, with it being understood that the description of the like parts of the frangible zone 70 applies to the frangible zone 270 unless otherwise noted. The composite core 262 defines both a pressure side 278 and a suction side 280. In this non-limiting example, a tip cap does not envelop any portion of the fan blade 30 and the composite core 262 defines the entire fan blade 30.

FIG. 6 is a schematic view of a fan blade 130 for the fan assembly 14 of the turbine engine 10 according to another aspect of the disclosure herein. The fan blade 130 is substantially similar to the fan blade 30. Therefore, like parts will be identified with like numerals increased by 100, with it being understood that the description of the like parts of the fan blade 30 applies to the fan blade 130 unless otherwise noted.

A leading edge strip 156 can be provided along a leading edge 146 of the fan blade 130. The leading edge strip 156 can be mounted to a composite core 162 of the fan blade 130. The leading edge strip 156 can include a first and second portion 156a, 156b wherein the first portion 156a extends in the span-wise direction from an area along the leading edge 146 proximate a root 150 of the fan blade 130 toward the tip 152. The first portion 156a of the leading edge strip 156 can be formed from any metallic material that facilitates operation of the fan assembly 14 as described herein, including, but not being limited to, titanium alloys, such as Ti64, steel alloys, or nickel alloys. The first portion 156a of the leading edge strip 156 can be formed from a material being a metallic material having a modulus of elasticity ranging from 15 Msi to 40 Msi. In some aspects of the disclosure herein the modulus of elasticity can range from 15 to 30 Msi. The modulus of elasticity quantifies the tendency of the first or second portions 156a, 156b to deform along an axis when opposing forces are applied along that axis. In other words the modulus of elasticity can quantify the stiffness of a material. An object with a relatively lower modulus of elasticity is a more flexible or pliable material.

The second portion 156b extends radially inward in the span-wise direction from the tip 152 toward first portion 156a. A leading edge fuse 163 is formed where the first portion 156a meets the second portion 156b. A leading edge fuse 163 can be formed, by way of non-limiting example with a butt joint or a lap joint. The first and second portions 156a, 156b can be bonded to each other, by way of non-limiting example using brazing, welding, or adhesive bonding. The first portion 156a can be coupled to the second portion 156b via methods, for example but not limited to, brazing, welding, and adhesive bonding. The second portion 156b of the leading edge strip 156 can be formed from any metallic material that facilitates operation of the fan assembly 14 as described herein and has a modulus of elasticity that is less than the material from which the first portion 156a of the leading edge strip 156 is formed. The second portion 156b can be formed from a material different than the material from which the first portion 156a is formed from. The second portion 156b can be a metal, including, but not being limited to, titanium alloys, such as CPTi, and aluminum alloys. The second portion 156b of the leading edge strip 156 can be formed from a metallic material having a modulus of elasticity ranging from 5 Msi to 15 Msi. In another aspect of the disclosure herein the modulus of elasticity of the material from which the second portion 156b is formed can be less than 80% of the value of the material from which the first portion 156a is formed.

In an aspect of the disclosure herein, the leading edge strip 156 and the composite core 162 both include frangible zones 165 and 170. Both the leading edge frangible zone 165 and the frangible zone 170 of the composite core 162 define portions of the fan blade 130 that are bendable or breakable when the blade comes in contact with, by way of non-limiting example, the nacelle 116 or a foreign object. Contact can occur during, by way of non-limiting example, an FBO as previously described herein. The frangible zones 165, 170 as described herein ensure that upon contact with the nacelle 16 or a foreign object, pieces break up that are small enough to pass through the engine safely. Consequently a safe shutdown of the engine 10 after an FBO is possible.

The leading edge strip 156, trailing edge strip 158, and blade tip cap 154 can be coupled to the composite core 162 via methods, for example but not limited to, brazing, welding, and adhesive bonding. The leading edge strip 156 can extend radially along a portion or substantially all as illustrated of a radial length (L) of the fan blade 130. The leading edge strip 156 can extend axially in the chord-wise direction along a portion as illustrated or substantially all of an axial width (W) of the fan blade 30.

FIG. 7 is a leading edge strip 256 according to another aspect of the disclosure discussed herein. The leading edge strip 256 is substantially similar to the leading edge strip 156. Therefore, like parts will be identified with like numerals increased by 100, with it being understood that the description of the like parts of the leading edge strip 156 applies to the leading edge strip 256 unless otherwise noted.

The leading edge strip 256 can include a third portion 256c. A first portion 256a of the leading edge strip 256 extends in the span-wise direction between a first end or root portion and a second end 290, 292 defining a majority of a leading edge 246 for the blade 30 as described herein. A second portion 256b is made of a material different than the first portion 256a and the third portion 256c. More specifically the portion has a modulus of elasticity that is lower than the modulus of elasticity of the material from which the first portion 256a or the third portion 256c is made. The second portion 256b extends radially outward from the second end 292 in the span-wise direction between a third and fourth end 294, 296.

The third portion 256c extends radially outward from the fourth end 296 in the span-wise direction. The third portion 256c can define at least a portion of a tip 252 for the blade 30 as described herein. The third portion 256c can be made of a different material than the first or second portions. It is also contemplated that the third portion 256c is made of the same material as the first portion 256a. The first, second, and third portions 256a, 256b, 256c can be bonded together by way of non-limiting example using brazing, welding, or adhesive bonding methods. The second portion 256b can define a leading edge fuse 263 formed between the first and third portions 256a, 256c where a leading edge frangible zone 265 can define an area around the leading edge fuse 263 for the blade 30 as described herein. It is contemplated that the leading edge fuse 263 is formed from two butt joints as illustrated or multiple lap joints.

The method for forming the blade 30 as described herein can further include forming the leading edge strip 156, 256 with a first portion 156a, 256a made from a material different than a material from which a second portion 156b, 256b is made and where the second portion 156b, 256b has a lower modulus of elasticity than the first portion 156a, 256a.

Benefits associated with forming a fan blade having a non-uniform composite material core include a controlled fan blade fragmentation during extreme events. Changing the material at the tip region of the blade to a different modulus of elasticity than the rest of the blade material enables the controlled fan blade fragmentation. By changing the modulus of elasticity of the material locally at the tip of the blade the fuse as described herein can be generated which would allow the blade tip to fail under heavy rub loads.

Selectively failing the blade during operation, by way of non-limiting example during an FBO, can contribute to cost savings in other areas of the engine. Removal of current known trench filler systems located on a containment case inner diameter where the blade conventionally nests can occur. This will reduce overall weight and lead to cost savings.

To the extent not already described, the different features and structures of the various embodiments can be used in combination with each other as desired. That one feature is not illustrated in all of the embodiments is not meant to be construed that it cannot be, but is done for brevity of description. Thus, the various features of the different embodiments can be mixed and matched as desired to form new embodiments, whether or not the new embodiments are expressly described. All combinations or permutations of features described herein are covered by this disclosure.

This written description uses examples 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 can 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 have 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.

Claims

1. An airfoil for a turbine engine, the airfoil comprising:

a composite core having a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction and having a span-length;

a first portion of the composite core extending radially from the root less than 100% of the span-length and formed from a first material;

a second portion of the composite core proximate the first portion and defining at least a portion of the tip and formed from a second material, having a lower strain capability than the first material and having a different composition than the first material.

2. The airfoil of claim 1 wherein the first material is a compliant material having a strain capability greater than or equal to 1% and the second material is a less compliant material having a strain capability less than or equal to 1%.

3. The airfoil of claim 1 wherein the second material has a strain capability that is less than 85% of the strain capability of the first material.

4. The airfoil of claim 1 wherein the second portion extends radially from the first portion up to 5% of the span-length.

5. The airfoil of claim 1 wherein the second material has a modulus of elasticity greater than 10 Msi.

6. The airfoil of claim 1 wherein the first material has a modulus of elasticity between 5 Msi and 20 Msi.

7. The airfoil of claim 1 further comprising a fuse between the first portion and the second portion to define a frangible zone.

8. The airfoil of claim 7 wherein the fuse is one of a scarfed, interwoven, butt, or lap joint.

9. The airfoil of claim 1 further comprising a tip cap extending along at least a portion of the tip.

10. The airfoil of claim 9 wherein the tip cap extends along at least 50% of the tip in the chord-wise direction.

11. The airfoil of claim 10 wherein the tip cap extends along 100% of the tip.

12. The airfoil of claim 1 wherein at least one of the first or second materials are composite materials.

13. The airfoil of claim 12 wherein the second material is an epoxy fiber material.

14. The airfoil of claim 1 further comprising a leading edge strip formed from at least two different materials.

15. The airfoil of claim 14 wherein the leading edge strip comprises a leading edge frangible zone.

16. A blade for a turbine engine, the blade comprising:

interior composite core having a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction and having a span-length;

a first portion of the composite core extending radially from the root less than 100% of the span-length and formed from a first material;

a second portion of the composite core proximate the first portion and defining at least a portion of the tip and formed from a second material having a lower strain capability than the first material and a different composition than the first material.

17. The blade of claim 16 wherein the second material has a strain capability that is less than 85% of the strain capability of the first material.

18. The blade of claim 16 wherein the second material has a modulus of elasticity greater than 10 Msi.

19. The blade of claim 16 wherein the first material has a modulus of elasticity between 5 Msi and 20 Msi.

20. The blade of claim 16 further comprising a fuse between the first portion and the second portion to define a frangible zone.

21. The blade of claim 20 wherein the fuse is one of a scarfed, interwoven, butt, or lap joint.

22. The blade of claim 16 further comprising a tip cap extending along at least a portion of the tip.

23. The blade of claim 16 wherein at least one of the first or second materials are composite materials.

24. The airfoil of claim 16 further comprising a leading edge strip formed from at least two different materials.

25. The airfoil of claim 24 wherein the leading edge strip comprises a leading edge frangible zone.

26. A method for forming a blade having composite core and extending radially between a root and a tip defining a span-wise direction and having a span-length, the method comprising:

forming a first and a second portion of the composite core;

extending the first portion radially outward from the root less than 100% of the span-length;

extending the second portion radially inward from the tip to meet the first portion; and

joining the first portion to the second portion to form a fuse.

27. The method of claim 26 wherein forming the fuse includes one of a scarfing, interweaving, abutting, or lapping the first portion to the second portion.

28. The method of claim 26 further comprising extending a tip cap around at least a portion of the tip.

29. The method of claim 26 wherein forming the second portion includes forming the second portion from a stiff non-compliant material.

30. The method of claim 26 further comprising forming a leading edge strip having a first portion made and a second portion made from a material different than the material from which the first portion is made and having a lower modulus of elasticity than the first portion.