Fragile airfoil with shape memory alloy

Abstract

An airfoil includes a plurality of composite material layers extending from a leading edge to a trailing edge and between a tip and a root. The airfoil further includes a frangible airfoil section at the tip that extends between the leading edge and the trailing edge and extends spanwise between the tip and a frangible line. The frangible airfoil section includes a plurality of composite material layers and one or more shape memory alloy inserts disposed between the plurality of composite material layers. Also disclosed is a gas turbine engine including a frangible airfoil and a method for forming a frangible airfoil.

Description

[0001] Priority Information

[0002] This application claims priority to Indian Patent Application No. 202111033700, filed on July 27, 2021. Technical Field

[0003] The present subject matter generally relates to airfoils, and more particularly, to frangible airfoils including shape memory alloy (SMA) materials. Background Art

[0004] Gas turbine engines, such as turbofans, generally include a fan case surrounding a fan assembly that includes fan blades. The fan case is typically configured to withstand impacts to the fan blades due to adverse engine conditions that result in failure modes such as foreign object damage, hard friction due to excessive or extreme imbalance or fan rotor oscillation or fan blade release. The fan case generally includes a frangible structure, such as honeycomb or trench fill material, that is configured to mitigate load transfer to and through the fan case. Incorporating such trench fill material results in a larger and / or heavier fan case. The inventors of the present invention have found a need for an airfoil that can achieve a controlled and consistent failure mode of the airfoil that can reduce cost, weight, and load transfer to the surrounding case. Brief Description of the Drawings

[0005] A complete and enabling disclosure to one of ordinary skill in the art is set forth in the specification, including its best mode, which refers to the drawings in which:

[0006] Figure 1 A cross-sectional view of one embodiment of a gas turbine engine utilized within an aircraft in accordance with aspects of the present subject matter is shown, specifically showing a gas turbine engine configured as a high bypass turbofan jet engine;

[0007] Figure 2 A cross-sectional view of a Figure 1 fan section in accordance with aspects of the present subject matter is shown, specifically showing the fan blades of the fan section;

[0008] Figure 3 A cross-sectional view of a Figure 1 and 2 fan blades within the fan section of

[0009] Figure 4 is shown in accordance with aspects of the present subject matter, specifically showing the frangible airfoil portion and the remaining airfoil portion;

[0010] Figure 5A perspective view of an SMA insert according to various aspects of the present subject matter is shown;

[0011] Figure 6 A cross-sectional view of one embodiment of a frangible airfoil section including an SMA insert according to an aspect of the present subject matter is shown;

[0012] Figure 7 A cross-sectional view of one embodiment of a frangible airfoil section including an SMA insert according to an aspect of the present subject matter is shown; and

[0013] Figure 8 A method of forming a frangible airfoil according to an aspect of the present subject matter is depicted.

[0014] The reuse of reference symbols in this specification and the drawings is intended to represent the same or similar features or elements of the invention. Detailed Description

[0015] Reference will now 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 to the invention without departing from the scope or spirit of the invention. For example, features shown or described as part of one embodiment can be used with another embodiment to yield yet another embodiment. Accordingly, the invention is intended to cover such modifications and variations that come within the scope of the appended claims and their equivalents.

[0016] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to denote the position or importance of a single component.

[0017] The terms “upstream” and “downstream” refer to the relative direction with respect to the flow of fluid in a fluid path. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

[0018] The terms “coupled,” “fixed,” “attached,” etc. refer to 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.

[0019] The terms “communicate,” “in communication,” “communicating,” etc. refer to direct communication as well as indirect communication such as through a memory system or another intermediate system.

[0020] As used herein, the term "exemplary" is used to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless otherwise specifically noted, all embodiments described herein are to be considered exemplary.

[0021] Unless the context clearly dictates otherwise, the singular forms "a," "an," and "the" include plural references.

[0022] The approximating language used herein and in the claims is used to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by one or more terms, such as "about," "approximately," and "substantially," is not limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the component and / or system. For example, the approximating language may refer to being within a margin of 1, 2, 4, 10, 15, or 20%. These approximating margins may be applied to a single value defining an endpoint of a numerical range and / or to the margins of a range between endpoints.

[0023] Herein and throughout the specification and claims, range limitations are combined and interchanged, and these ranges are identified and include all the subranges subsumed therein, unless the context or language indicates otherwise. For example, all ranges disclosed herein include the endpoints, and the endpoints may be combined with each other independently.

[0024] Generally provided is a frangible airfoil for a gas turbine engine. The airfoil may define a frangible airfoil section that extends from a trailing edge to a leading edge and extends less than about 50% of the wingspan from a tip. The frangible airfoil section includes one or more composite material layers and one or more SMA inserts. The frangible airfoil allows the frangible airfoil section to deform without separating the frangible airfoil section from the remainder of the airfoil after an event that creates an imbalance. The embodiments generally shown and described herein may achieve a controlled and consistent failure of an airfoil (such as a fan blade) after a fault event such as a hard friction with a surrounding fan case. The embodiments generally described herein enable the airfoil to deform at a desired wingspan of the airfoil to reduce the load transfer to the surrounding case. The embodiments generally provided herein may further cause the airfoil to deform such that an excessive or extreme imbalance of a fan rotor may be reduced after a fault event (such as airfoil release, foreign object damage (such as bird strike, icing, etc.), or loss of inserts in a lubricating oil or bearing assembly). Using the described frangible airfoil may reduce the amount of trench fill material required when incorporated into a ducted gas turbine engine, resulting in a lighter and more efficient fan case.

[0025] Now referring to the drawings,Figure 1 FIG. Figure 1 shows a cross-sectional view of an embodiment of a gas turbine engine 10 that may be used within an aircraft according to aspects of the present subject matter. More specifically, for Figure 1 the embodiment of Figure 1 , the gas turbine engine is a high bypass turbofan jet engine, and shows that the gas turbine engine 10 has a longitudinal or axial centerline axis 12 that, for reference purposes, extends in the axial direction A through the gas turbine engine. The gas turbine engine 10 also defines a radial direction R extending from the axial centerline axis 12. Although an exemplary turbofan embodiment is shown, it is contemplated that the present disclosure may equally apply to turbomachinery, such as open rotors, turboshafts, turbojets, or turboprop configurations, including marine and industrial gas turbine engines and auxiliary power units.

[0026] Generally, the gas turbine engine 10 includes a turbine (generally denoted by reference numeral 14) and a fan section 16 located upstream thereof. The turbine 14 generally includes a substantially tubular outer casing 18 that defines an annular inlet 20. Additionally, the outer casing 18 further encloses and supports a low pressure (LP) compressor 22 for increasing the pressure of the air entering the turbine 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 the air. The pressurized air leaving the HP compressor 24 may then flow to a combustor 26 where fuel is injected into the pressurized air stream and the resulting mixture is 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 30, and then to a low pressure (LP) turbine 32 for driving the LP compressor 22 and the fan section 16 via a low pressure (LP) shaft 34, the low pressure (LP) shaft 34 generally being coaxial with the HP shaft 30. After driving each turbine 28 and 32, the combustion products may be exhausted from the turbine 14 via an exhaust nozzle 36 to provide a propulsive jet thrust.

[0027] Additionally, the fan section 16 of the gas turbine engine 10 generally includes a rotatable axial flow fan rotor 38 that is configured to be surrounded by an annular fan casing 40. In a particular embodiment, the LP shaft 34 may be directly connected to the fan rotor 38, such as in a direct drive configuration. In an alternative configuration, the LP shaft 34 may be connected to the fan rotor 38 via a reduction gear 37, such as a reduction gearbox in an indirect drive or gear drive configuration. Such a reduction gear may be included between any suitable shafts / spools within the gas turbine engine 10 as desired or required.

[0028] Those of ordinary skill in the art will understand that the fan casing 40 can be configured to be supported relative to the turbine 14 by a plurality of substantially radially extending, circumferentially spaced outlet guide vanes 42. Thus, the fan casing 40 can enclose the fan rotor 38 and its corresponding fan rotor blades (fan blades 44). In addition, a downstream section 46 of the fan casing 40 can extend over an outer portion of the turbine 14 to define a secondary or bypass airflow line 48 that provides additional propulsive jet thrust.

[0029] During operation of the gas turbine engine 10, it should be understood that an initial airflow (shown by arrow 50) can enter the gas turbine engine 10 through an associated inlet 52 of the fan casing 40. The airflow 50 then passes through the fan blades 44 and divides into a first compressed airflow (shown by arrow 54) and a second compressed airflow (shown by arrow 56), with the first compressed airflow moving through the bypass line 48 and the second compressed airflow entering the low-pressure compressor 22. The pressure of the second compressed airflow 56 is then increased and enters the HP compressor 24 (as shown by arrow 58). After being mixed with fuel and burned within the combustor 26, the combustion products 60 leave the combustor 26 and flow through the HP turbine 28. Thereafter, the combustion products 60 flow through the LP turbine 32 and exit through the exhaust nozzle 36 to provide thrust for the gas turbine engine 10.

[0030] Referring to Figure 2 , exemplary airfoil 62 embodiments are provided in the context of the fan blades 44 positioned in the fan section 16. Although the airfoil 62 shown is depicted as part of the fan blade 44, it should be understood that the following discussion of the airfoil 62 can equally apply to another airfoil embodiment, e.g., the stator vanes or rotor blades of the compressors 22, 24 and / or turbines 28, 32 (see Figure 1)。The tubular housing 18, the LP compressor 22, and the exit guide vanes 42 are also shown. It should be noted that during operation of the gas turbine engine 10, the initial air flow (shown by arrow 50) enters the gas turbine engine 10 through the associated inlet 52 of the fan housing 40. Then, the air flow 50 passes through the fan blades 44. Optionally, each fan blade 44 includes an integral member having an axial dovetail 76 leading to the transition section 80. When installed within the gas turbine engine 10, the axial dovetail 76 is disposed within a dovetail slot of the fan rotor disk 39, thereby attaching the fan blade 44 to the fan rotor 38. Each fan blade 44 extends radially outward from the airfoil root 64 to the airfoil tip 66. The airfoil includes a remaining airfoil portion 92 and a frangible airfoil portion 94 that meet at a frangible line 88. The remaining airfoil portion 92 and the frangible airfoil portion 94 each extend between the leading edge 72 and the trailing edge 74 of the airfoil and the airfoil root 64 and the airfoil tip 66. The frangible airfoil portion 94 includes one or more SMA inserts 100.

[0031] During operation of the gas turbine engine 10, after an event that creates a significant imbalance in, for example, the fan rotor 38 or the LP shaft 34, the frangible airfoil portion 94 of the fan blade 44 can be configured to deform, for example, along the frangible line 88 or radially outside the frangible line 88 or to partially separate from the remainder of the airfoil 62. Events that create a substantial imbalance in the fan rotor 38 and / or the LP shaft 34 can include, but are not limited to, foreign object damage (e.g., bird strike, ice ingestion, other debris, etc.) or detachment of the fan blade 44. As the fan rotor 38 and / or the LP shaft 34 continue to rotate, the deformation of the frangible airfoil portion 94 can reduce undesirable imbalance or vibration. Further, delamination of the frangible airfoil portion 94 can occur during light friction conditions (e.g., crosswind) or during moderate friction conditions (e.g., bird strike). Inclusion of the SMA inserts 100 described herein helps prevent material wear at the airfoil tip 66 under certain conditions and enables controllable fan friction. Additionally, during light friction events, inclusion of the SMA inserts 100 can reduce material wear along the airfoil tip. Additionally, embodiments of the airfoil 62 can enable a lighter fan housing 44 or nacelle, such as reducing the amount of metallic or wearable material in the fan housing 40 or nacelle.

[0032] Now referring to Figure 3, an exemplary airfoil 62 for a fan blade 44 is shown. Optionally, each fan blade 44 includes an integral component having an axial dovetail 76 that has a pair of opposing pressure faces 78 leading to a transition section 80. The fan blade 44 extends radially outward along a span that defines a spanwise direction 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 a leading edge 72 to a trailing edge 74 of the airfoil and extend along the span between the airfoil root 64 and the airfoil tip 66. Additionally, it should be appreciated that the airfoil 62 can define a chordwise direction C along the chord at each point along the span and extends between the leading edge 72 and the trailing edge 74. Further, the chord can vary along the span of the airfoil 62. For example, in the depicted embodiment, the chord increases along the span toward the airfoil tip 66. However, in other embodiments, the chord can be approximately constant throughout the span or can decrease from the airfoil root 64 to the airfoil tip 66.

[0033] As Figure 3 Specifically shown in, the airfoil 62 can define a thickness direction T along the thickness extending between the pressure side 68 and the suction side 70 at each point along the span. In some embodiments, the thickness of the entire span of the airfoil 62 can be approximately constant. In other embodiments, the airfoil 62 can define a variable thickness between the airfoil root 64 and the airfoil tip 66. For example, the thickness can generally decrease along the span toward the airfoil tip 66. Additionally, the airfoil 62 can define an approximately constant thickness along the chord at each point along the span. Alternatively, in other embodiments, at least one point along the span of the airfoil 62 can define a variable thickness along the chord. For example, the airfoil 62 can define a maximum thickness at the position along the chord at each point along the span.

[0034] Furthermore, the airfoil 62 can define a frangible line 88 that separates a frangible airfoil portion 94 and a remaining airfoil portion 92. The frangible airfoil portion 94 can be generally positioned toward the airfoil tip 66 and extends between the leading edge 72 and the trailing edge 74 and between the airfoil tip 66 and the frangible line 88. The remaining portion can extend along the spanwise direction S from the frangible line 88 to the airfoil root 64. Optionally, a metal leading edge shroud 71 can cover an axially extending portion of the airfoil 62, and the airfoil 62 includes at least a portion of the leading edge 72. The metal leading edge shroud 71 can also cover a portion of the tip 66 and the trailing edge 74 (not shown in the figure).

[0035] As Figure 3As further shown, the frangible line 88 may generally extend between the leading edge 72 and the trailing edge 74 in a chordwise direction C. It should be appreciated that the frangible line 88 may generally extend in the chordwise direction C at approximately the same points along the span. In other embodiments, as the frangible line 88 extends axially in the chordwise direction C toward the trailing edge 74, the frangible line 88 may extend at least partially radially inward or outward (e.g., in the spanwise direction S). For example, in certain embodiments, the frangible line 88 may be defined at a point along the chordwise direction C along the span. In other embodiments, the frangible line 88 may extend from the leading edge 72 to the trailing edge 74 along a variable percentage of the span.

[0036] In one embodiment, the frangible airfoil section 94 may extend from the airfoil tip 66 along at least 10% of the span, such as extending in the chordwise direction C at a point along the span. More specifically, as shown, the frangible airfoil section 94 may define a frangible height 84 that extends between the airfoil tip 66 and the frangible line 88. Thus, the frangible height 84 may extend between the leading edge 72 and the trailing edge 74 along at least 10% of the span. In another embodiment, the frangible airfoil section 94 and / or the frangible height 84 may extend between the leading edge 72 and the trailing edge 74 along at least 15%, but less than 50% of the span. In other embodiments, the frangible airfoil section 94 may extend from the airfoil tip 66 along less than 50% of the span. The frangible airfoil section 94 may have a reduced overall bending stiffness compared to the remaining airfoil section 92.

[0037] One or more SMA inserts 100 are disposed in the frangible airfoil section 94. The arrangement of the one or more SMA inserts 100 allows the frangible airfoil section 94 of the airfoil tip 66 to change shape and / or deform in response to an event that creates an imbalance. For example, the one or more SMA inserts 100 allow the frangible airfoil section 94 to deform after loading or impacting the airfoil 62. For example, the frangible airfoil section 94 including the SMA inserts may change shape, structure, or deform at or radially outside of the frangible line 88. Specifically, in an embodiment, the frangible airfoil section 94 may be configured to at least partially delaminate when exposed to an event that creates an imbalance, without being completely severed from the remaining airfoil section 92. For example, including the SMA inserts 100 within the frangible airfoil section 94 allows for partial delamination of the frangible airfoil section 94 without severing the entire frangible airfoil section 94 at or radially outside of the frangible line 88.

[0038] Figure 4 is shown Figure 3 a cross-sectional view of the airfoil 62 taken along the B-B axis as shown therein. In other words, Figure 4Shows a cross-section of an airfoil taken along the thickness direction T. The airfoil 62 includes a suction side 70 and a pressure side 68, and the pressure side 68 includes a metallic leading edge shroud 71 which is disposed on a portion through the thickness leading edge 72. The SMA insert 100 is disposed in the frangible airfoil portion 94 between the suction side 70 and the pressure side 68 of the airfoil 62. The SMA insert 100 includes a top edge 110 and a bottom edge 120, the top edge 110 is disposed to face the airfoil tip 66 in the spanwise direction S, and the bottom edge 120 is disposed to face the airfoil root 64 in the spanwise direction S relative to the top edge 110. The SMA insert 100 includes side faces 130a, 130b which are disposed between the top edge 110 and the bottom edge 120 and connect the top edge 110 to the bottom edge 120. The midpoint M generally defines a point or axis located at approximately 50% of the spanwise height of the SMA insert 100. Although the midpoint M is generally shown in the middle of the SMA insert 100, it should be understood that the midpoint M does not necessarily correspond to the exact midpoint M of the SMA insert 100 along the spanwise direction S. For example, although the midpoint M is shown as approximately 50% of the spanwise height, it should be understood that the midpoint can be defined at any position from approximately 25% to approximately 75% of the spanwise height. The SMA insert 100 further includes a first leading edge face 140a and a second leading edge face 140b, which will be discussed in more detail with reference to Figure 5 The first leading edge face 140a and the second leading edge face 140b will be discussed in more detail.

[0039] The SMA insert 100 is generally disposed between or within one or more composite material layers 82 in the frangible airfoil portion 94. Although only two exemplary composite material layers 82 are shown, the present disclosure is not limited thereto. In fact, any number of composite material layers 82 located in various directions can be used to form the frangible airfoil portion 94 of the airfoil 62 as provided herein. Additionally, although one or more composite material layers 82 are shown in the spanwise direction S, the present disclosure is not limited thereto. In fact, any number of composite material layers 82 oriented in various directions (such as the spanwise direction S, the chordwise direction C, the thickness direction T) can be used in the frangible airfoil portion 94 according to the exemplary embodiments disclosed herein.

[0040] The SMA insert 100 includes an insert geometry that facilitates mechanical interlocking between one or more SMA inserts 100 and one or more composite material layers 82 used to form the frangible airfoil portion 94. Although in Figure 4Exemplary geometries are shown, but the present disclosure is not limited thereto. In fact, any geometry that can facilitate the mechanical interlock of the SMA insert 100 within the composite layer 82 for forming the frangible airfoil section 94 can be used. Suitable geometries can include an I-shaped geometry, a T-shaped geometry, a Y-shaped geometry, an X-shaped geometry, and / or combinations and variations thereof.

[0041] Figure 5 A perspective view of an exemplary SMA insert 100 is shown. The SMA insert 100 has a top edge 110 and a bottom edge 120, the top edge 110 being spaced from the airfoil tip 66 in the spanwise direction S, and the bottom edge 120 being spaced from the top edge 110 in the spanwise direction S. The sides 130a, 130b of the SMA insert 100 are arranged in a tapered manner to facilitate the mechanical interlock of the SMA insert 100 within the composite layer 82 ( Figure 5 not shown) of the frangible airfoil section 94. For example, the sides 130a, 130b taper in the spanwise direction S from the top edge 110 to the midpoint M. Similarly, the sides 130a, 130b taper in the spanwise direction S from the bottom edge 120 to the midpoint M. This tapering of the sides 130a, 130b facilitates the mechanical interlock of the SMA insert 100 between and within one or more composite layers 82 ( Figure 5 not shown) for forming the frangible airfoil section 94. Additionally, it should be understood that the SMA insert 100, specifically the sides 130a and 130b, can taper from the top edge 110 or the bottom edge 120 to the midpoint M in any direction including the thickness direction T, the chordwise direction C, the spanwise direction S, and combinations thereof. Further, as shown, the width of the top edge 110 and the bottom edge 120 in the thickness direction T is greater than the width of the midpoint M in the thickness direction T. In an embodiment, the SMA insert 100 has a span height ranging from about 0.1 inches to about 6.0 inches, such as from about 1 inch to about 5 inches, such as from about 2 inches to about 4 inches.

[0042] Figure 5The SMA insert 100 can be placed between the leading edge 72 and the trailing edge 74 in the chordwise direction C. As shown, the SMA insert 100 includes a first leading edge face 140a that extends from the top edge 110 of the SMA insert 100 to the midpoint M and extends between the sides 130a, 130b. A second leading edge face 140b extends from the bottom edge 120 of the SMA insert 100 to the midpoint M and extends between the sides 130a, 130b. A first leading edge angle α is formed between the surface of the first leading edge face 140a and the axis at the midpoint M. As shown, in an embodiment, the first leading edge angle α can be in the range of about 110° to about 15°, such as in the range from about 90° to 30°, such as in the range from about 75° to 50°. In an embodiment, the first leading edge angle α includes an angle less than about 90° to facilitate interlocking the SMA insert 100 in the fragile airfoil section 94. Similarly, a second leading edge angle θ is formed between the surface of the second leading edge face 140b and the axis at the midpoint M. As shown, in an embodiment, the second leading edge angle θ can be in the range of about 110 degrees to about 15°, such as about 90° to about 30°, such as about 75° to about 50°. In an embodiment, the second leading edge angle θ includes an angle less than about 90° to facilitate interlocking the SMA insert 100 in the fragile airfoil section 94. The first leading edge angle α and the second leading edge angle θ can be the same or different. For example, in some embodiments, the first leading edge angle α can be greater than or less than the second leading edge angle θ. For example, in order to adjust the fragility or deformability of the airfoil 62, more specifically the fragile airfoil section 94, to a desired parameter, the first leading edge angle α can be less than or greater than the second leading edge angle θ.

[0043] Similar features also exist along the trailing edge 74 on the SMA insert 100. Still referring to Figure 5, the SMA insert 100 includes a first trailing edge surface 150a that extends from the top edge 110 of the SMA insert 100 to the midpoint M and extends between the side surfaces 130a, 130b. A second trailing edge surface 150b extends from the bottom edge 120 of the SMA insert 100 to the midpoint M and extends between the side surfaces 130a, 130b. A first trailing edge angle β is formed between the surface of the first trailing edge surface 150a and the axis with respect to the midpoint M. As shown, in an embodiment, the first trailing edge angle β can be in the range of about 110° to about 15°, such as in the range of about 90° to about 30°, such as in the range of about 75° to about 50°. In an embodiment, the first trailing edge angle β includes an angle less than about 90° to facilitate interlocking the SMA insert 100 in the frangible airfoil portion 94. Similarly, a second trailing edge angle γ is formed between the surface of the second trailing edge surface 150b and the axis with respect to the midpoint M. As shown, in an embodiment, the second trailing edge angle γ can be in the range of about 110° to about 15°, such as in the range of about 90° to about 30°, such as in the range of about 75° to about 50°. In an embodiment, the second trailing edge angle γ includes an angle less than about 90° to facilitate interlocking the SMA insert 100 in the frangible airfoil portion 94. The first trailing edge angle β and the second trailing edge angle γ can be the same or different. For example, in some embodiments, the first trailing edge angle β can be greater than or less than the second trailing edge angle γ. For example, in order to adjust the fragility or deformability of the airfoil 62, more specifically the frangible airfoil portion 94, to a desired parameter, the first trailing edge angle β can be less than or greater than the second trailing edge angle γ.

[0044] Figures 6 - 7 The SMA insert 100 disposed in the frangible airfoil portion 94 is shown relative to the chordwise direction C. As shown, the SMA insert 100 is disposed relative to the airfoil tip 66 in the chordwise direction C and the spanwise direction S. Specifically, Figure 6 The SMA insert 100 including a single SMA insert 100 is shown, and the single SMA insert 100 unfolds from the leading edge 72 of the airfoil to the trailing edge 74. As shown, the insert 100 includes a top edge 110, a bottom edge 120, and a side surface 130a. In other embodiments, as Figure 7 shown, the SMA insert 100 can include one or more, such as a plurality of, SMA inserts 100 sequentially placed in the chordwise direction from the leading edge 72 of the airfoil to the trailing edge 74. Similar to Figure 6 the insert, as shown, the SMA insert 100 includes a top edge 110, a bottom edge 120, a side surface 130a, and a midpoint M.

[0045] In some embodiments, the SMA insert may include an SMA material as a main component in an amount greater than 50 wt.% (weight percentage) of the SMA insert. In certain embodiments, the SMA insert may consist essentially of the SMA material. The SMA material is generally an alloy that is capable of returning to its original shape after deformation. For example, the SMA material may define a hysteresis effect, where the load path on a stress-strain diagram is different from the unloading path on the stress-strain diagram. The SMA material may also provide a varying stiffness in a predetermined manner in response to a specific range of stress and temperature. The change in hardness of the shape memory alloy is due to a temperature-related solid-state microstructure phase change, which enables the alloy to change from one physical shape to another. The change in stiffness of the SMA material can be produced by processing and annealing a preform of the alloy at or above the temperature at which the solid-state microstructure phase change of the shape memory alloy occurs. The temperature at which this phase change occurs is generally referred to as the critical temperature or transformation temperature of the alloy. In the manufacture of an SMA insert for changing stiffness during the operation of an airfoil 62, the SMA insert may be formed to have one operating stiffness (e.g., a first stiffness) below the transformation temperature and another stiffness (e.g., a second stiffness) at or above the transformation temperature.

[0046] Some shape memory alloys used herein are characterized by temperature-related phase changes. These phases include the martensite phase and the austenite phase. The martensite phase generally refers to the phase at a lower temperature. The austenite phase generally refers to the phase at a higher temperature. The martensite phase is generally more deformable, while the austenite phase is generally less deformable. When a shape memory alloy is in the martensite phase and is heated above a certain temperature, the shape memory alloy begins to transform into the austenite phase. The temperature at which this phenomenon begins is called the austenite start temperature (As). The temperature at which this phenomenon is completed is called the austenite finish temperature (Af). When a shape memory alloy in the austenite phase is cooled, it begins to transform into the martensite phase. The temperature at which this transformation begins is called the martensite start temperature (Ms). The temperature at which the martensite phase transformation is completed is called the martensite finish temperature (Mf). As used herein, the term "transformation temperature" without any further qualifier may refer to either the martensite transformation temperature or the austenite transformation temperature. Additionally, "below the transformation temperature" without the "start temperature" or "finish temperature" qualifier generally refers to a temperature below the martensite finish temperature, and "above the transformation temperature" without the "start temperature" or "finish temperature" qualifier generally refers to a temperature above the austenite finish temperature.

[0047] In some embodiments, the SMA insert may define a first stiffness at a first temperature and a second stiffness at a second temperature, where the second temperature is different from the first temperature. Additionally, in some embodiments, one of the first temperature or the second temperature is below the transformation temperature and the other may be at or above the transformation temperature. Thus, in some embodiments, the first temperature may be below the transformation temperature and the second temperature may be at or above the transformation temperature. Although in some other embodiments, the first temperature may be at or above the transformation temperature and the second temperature may be below the transformation temperature. Further, the various SMA inserts 100 described herein may be configured to have different first rigidities and different second rigidities at the same first temperature and second temperature.

[0048] Non-limiting examples of SMAs that may be suitable for forming the various SMA inserts described herein may include nickel-titanium (NiTi) and other nickel-titanium-based alloys such as nickel-titanium hafnium (NiTiHf) and nickel-titanium palladium (NiTiPd). However, it should be understood that other SMA materials may be equally applicable to the present disclosure. For example, in certain embodiments, the SMA material may include nickel-aluminum-based alloys, copper-aluminum-nickel alloys, or alloys containing zinc, zirconium, copper, gold, platinum, and / or iron. The alloy composition may be selected to provide the stiffness effects required for the application, such as but not limited to damping capacity, transformation temperature, and strain, strain hysteresis, yield strength (martensitic and austenitic phases), oxidation resistance, and thermal corrosion, the ability to change shape through repeated cycling, the ability to exhibit one-way or two-way shape memory effects, and / or many other engineering design criteria. Suitable shape memory alloy compositions that may be used in the embodiments of the present disclosure may include but are not limited to NiTi, NiTiHf, NiTiPt, NiTiPd, NiTiCu, NiTiNb, NiTiVd, TiNb, CuAlBe, CuZnAl, and some iron-based alloys. In some embodiments, a NiTi alloy having a transformation temperature between 5°C and 150°C is used. The NiTi alloy transforms from austenite to martensite upon cooling.

[0049] Additionally, the SMA material may also exhibit superelasticity. Superelasticity is generally characterized by the recovery of large strains, with some dissipation possible. For example, the martensitic and austenitic phases of the SMA material may respond to mechanical stress as well as temperature-induced phase transformations. For example, the SMA may be loaded in the austenitic phase (i.e., above a certain temperature). Thus, when the critical stress is reached, the material may begin to transform into the (twinned) martensitic phase. Under continued loading and assuming isothermal conditions, the (twinned) martensite may begin to detwin, allowing the material to undergo plastic deformation. If unloading occurs before plasticity, the martensite generally transforms back to austenite and the material may recover its original shape by generating hysteresis.

[0050] Prior to the SMA insert being assembled within or associated with the airfoil 62, the various SMA inserts disclosed herein can be in a non-stressed state. Additionally, after the assembly or arrangement of the SMA insert relative to the airfoil, the SMA insert can be in a prestressed state. For example, the SMA insert can be in a prestressed state in a compressed manner after being assembled or arranged relative to the airfoil 62.

[0051] In some embodiments, a single SMA insert or some or all of the SMA inserts can be in a pre-strained or prestressed state (e.g., prestressed state). The prestressed SMA insert can shift the hysteresis loop of the SMA insert to a different stress range compared to a non-prestressed SMA insert (e.g., when the SMA insert is in a non-stressed state). The prestress is further used to maximize the damping function of the SMA insert, thereby making the material effective at the maximum stress generated. More specifically, placing the SMA insert in a prestressed position or state can allow the insert to enter a hysteretic bending state without requiring a relatively large amount of displacement. For example, in certain embodiments, the various SMA inserts disclosed herein can be prestressed between 70 GPa and 150 GPa. Further, it should be understood that in embodiments including more than one SMA insert, the SMA inserts can be prestressed to the same or approximately the same stress or strain. However, additionally or alternatively, one or more SMA inserts can be prestressed or pre-strained to different degrees in order to provide an appropriate amount of damping for the position or condition of a particular SMA insert.

[0052] In an embodiment, the airfoil 62 is at least partially formed of a ceramic matrix composite. The composite material can include, but is not limited to, metal matrix composites (MMC), polymer matrix composites (PMC), or ceramic matrix composites (CMC). Composite materials, such as those that can be used in the airfoil 62, generally include a fiber reinforcement material embedded in a matrix material, such as a polymer, ceramic, or metal material. The reinforcement material serves as the load-bearing component of the composite material, while the matrix of the composite material is used to bind the fibers together and serves as a medium for transferring and distributing externally applied stresses to the fibers.

[0053] Exemplary CMC materials can include silicon carbide (SiC), silicon, silica, or alumina matrix materials and combinations thereof. Ceramic fibers can be embedded in the matrix, such as oxidation-stable reinforcing fibers, including monofilaments, such as sapphire and silicon carbide (e.g., SCS-6 from Textron), and rovings and yarns including silicon carbide (e.g., from Ube Industries and ), aluminosilicates (such as Nextel 440 and 480), chopped whiskers and fibers (such as Nextel 440 and ), and optionally ceramic particles (such as oxides of Si, Al, Zr, Y and combinations thereof) and inorganic fillers (such as pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For example, in certain embodiments, a fiber bundle including a ceramic refractory coating can be formed into a reinforcing tape, such as a unidirectional reinforcing tape. A plurality of tapes can be stacked together (e.g., as composite layer 82) to form a preform. The fiber bundle can be impregnated with a slurry composition before or after forming the preform. Then, the preform can be subjected to heat treatment, such as curing or burnout to produce a high char residue in the preform, and subsequent chemical treatment, such as molten silicon infiltration, to obtain a component formed of a CMC material having a desired chemical composition.

[0054] Similarly, in various embodiments, a PMC material can be manufactured by impregnating a fabric or unidirectional tape with a resin (prepreg) and then curing. For example, multiple layers of prepreg layers (e.g., composite layer 82) can be stacked to an appropriate thickness and orientation of the component, and then the resin can be hardened and cured to present a fiber-reinforced composite component. As another example, a mold can be utilized, and an uncured layer of prepreg can be stacked onto the mold to form at least a portion of a composite component. The mold can be in a closed configuration (e.g., compression molding) or an open configuration utilizing a vacuum bag molding. For example, in the open configuration, the mold forms one side of a blade (e.g., pressure side 68 or suction side 70). The PMC material is placed inside the bag, and the PMC material is held onto the mold using a vacuum during hardening. In other embodiments, the airfoil 62 can be formed at least in part via resin transfer molding (RTM), light resin transfer molding (LRTM), vacuum-assisted resin transfer molding (VARTM), a forming process (such as thermoforming), or a similar process.

[0055] Before impregnation, the fabric can be referred to as a "dry" fabric and generally includes a laminate of two or more fiber layers. The fiber layers can be formed of a variety of materials, non-limiting examples of which include carbon (such as graphite), glass (such as fiberglass), polymer (such as ), fibers, and metal fibers. The fiber reinforcement can be used in the form of relatively short chopped fibers, generally less than 2 inches in length, more preferably less than 1 inch, or long continuous fibers, the latter often being used to produce woven fabrics or unidirectional tapes. Other embodiments can include other textile forms, such as plain weave, twill, or satin.

[0056] In one embodiment, a PMC material can be manufactured by dispersing dry fibers into a mold and then allowing a matrix material to flow around the reinforcing fibers. Resins for PMC matrix materials can generally be classified as thermosetting or thermoplastic. Thermoplastic resins are generally classified as polymers that can be repeatedly softened and flowed when heated and cured due to physical changes rather than chemical changes when sufficiently cooled. Distinctive example classes of thermoplastic resins include nylon, thermoplastic polyester, polyaryletherketone, and polycarbonate resins. Specific examples of high-performance thermoplastic resins for aerospace applications include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), and polyphenylene sulfide (PPS). In contrast, once fully hardened into a hard rigid solid, thermosetting resins do not undergo significant softening when heated but instead thermally decompose when sufficiently heated. Distinctive examples of thermosetting resins include epoxy resins, bismaleimide (BMI), and polyimide resins.

[0057] During the manufacture of the airfoil 62 described herein, the SMA insert 100 can be placed between one or more material layers and processed accordingly to provide a frangible airfoil 62, as will be further discussed below.

[0058] Now referring Figure 8 , a method 200 for forming a frangible airfoil is described in accordance with aspects of the present subject matter. In particular, the method 200 can be used to form various embodiments of an airfoil as Figures 2 - 7 shown. For example, a frangible airfoil can define a span extending in a spanwise direction S between a root and a tip, a chordwise direction C extending between a leading edge and a trailing edge, and a thickness direction T extending between a pressure side and a suction side.

[0059] The method 200 can include 202 laying a plurality of composite material layers to form a remaining airfoil portion. The plurality of composite material layers can include a composite material, such as a CMC material. The composite material layers can be laid on a tool, mandrel, mold, or other suitable support device or surface. At 204, the method includes laying a plurality of composite material layers and one or more shape memory alloy inserts to form a frangible portion of the airfoil. Depending on the desired characteristics of the airfoil, the composite material layers can be oriented in the chordwise direction C, thickness direction T, spanwise direction S, or a combination thereof. When laying one or more material layers, one or more SMA inserts can be placed between one or more of the layers. For example, a material layer can be laid, the SMA insert can be positioned on the material layer, and then one or more material layers can be laid on the material layer and the SMA insert. One or more SMA inserts can be included between one or more of the composite material layers used to form the frangible portion of the airfoil. The plurality of composite material layers can include a composite material, such as a CMC material. The composite material layers can be laid on a tool, mandrel, mold, or other suitable support device or surface.

[0060] Another step of method 200 may include 206 processing multiple material layers to form a frangible airfoil. In one embodiment, processing the composite material layer may include compacting the composite material layer. In another embodiment of method 200, processing the composite material layer may include autoclaving the composite material layer. In yet another embodiment of method 200, processing the composite material layer may include compacting the composite material layer and autoclaving the composite material layer. For example, the composite material layer may be compacted and then processed in an autoclave. Compaction may be performed in air, i.e., at room temperature and pressure. The autoclave cycle may impart stiffness to the final layer and / or laminate assembly through complete drying and / or curing of the composite material composition and produce the final dimensions of the composite part through complete consolidation of the layers and / or sub-assemblies.

[0061] Furthermore, in embodiments where the composite material layer is processed in an autoclave, the composite material layer may be autoclaved using soft and / or hard tools. For example, the composite material layer may be autoclaved using a metal tool (i.e., a hard tool) shaped to impart the desired shape of the frangible airfoil. As another example, a soft tool such as a vacuum bag may be used to autoclave the composite material layer. For example, the composite material layer may be supported on a metal tool, and then the composite material layer and the tool may be bagged, and air may be removed from the bag to apply pressure to the composite material layer and compact the composite material layer before the composite material layer is processed in an autoclave cycle. For example, processing the composite material layer may include autoclaving the composite material layer to form an autoclaved body. Additionally, another step may include firing the autoclaved body to form a fired body. Processing the composite material layer may further include densifying the fired body to form a composite part. In certain embodiments, processing the composite material layer may include at least one of melt infiltration or polymer infiltration and pyrolysis.

[0062] In an embodiment where the composite material is a CMC material, the autoclaved body can be subjected to firing (or burnout) to form a fired body, which is then densified to produce a densified CMC component. The densified CMC component is a monolithic component, i.e., the component is a continuous piece of CMC material. For example, after autoclaving, the component can be placed in a furnace to burn off any mandrel forming material and / or solvent used to form the CMC layer and to decompose the binder in the solvent, and then placed in a furnace with silicon to convert the ceramic matrix precursor of the layer into the ceramic material of the matrix of the CMC component. The silicon melts and infiltrates any pores generated within the matrix due to binder decomposition during combustion / firing; the CMC component is densified by silicon melt infiltration into the CMC component. However, any known densification technique can be used for densification, including but not limited to Silcomp, melt infiltration (MI), chemical vapor infiltration (CVI), polymer infiltration and pyrolysis (PIP), and oxide / oxide processes. In one embodiment, densification and firing can be carried out in a vacuum furnace or an inert atmosphere with an atmosphere established at a temperature above 1200 °C to allow silicon or another or multiple suitable materials to melt and infiltrate into the component.

[0063] Optionally, after processing, if desired, the composite component can be finished and coated with one or more coatings, such as an environmental barrier coating (EBC). For example, the composite material layer wound around the core may be oversized such that a portion of the composite material layer extends beyond the desired trailing edge of the airfoil. Thus, after processing, the composite material layer can be machined to define the trailing edge.

[0064] Of course, the method 200 is provided by way of example only with respect to Figure 8 the description. Thus, other known methods for compressing and / or hardening composite material layers and for densifying CMC components can be used. Optionally, any combination of these or other known methods can be used in any suitable order. Additionally, although the method 200 has been described with respect to a fan blade, the method 200 can also be used to form other composite components, such as turbine nozzle vanes and turbine stator vanes including airfoils and / or compressor blades and vanes, as exemplary composite components. Figure 8 the method 200,

[0065] Other aspects of the present disclosure are provided by the subject matter of the following clauses:

[0066] An airfoil defines a span extending in the spanwise direction between a root and a tip; a chordwise direction extending between a leading edge and a trailing edge; a thickness direction extending between a pressure side and a suction side; and a frangible line at a location along the span. The airfoil includes: a frangible airfoil portion extending between the leading edge and the trailing edge and along the span between a tip and the frangible line, the frangible airfoil portion including a plurality of composite material layers and one or more shape memory alloy (SMA) inserts; and a remaining airfoil portion extending along the span from the frangible line to the root, the remaining airfoil portion including a plurality of composite material layers, wherein after an event that creates an imbalance, the frangible airfoil portion deforms or partially detaches relative to the remaining airfoil portion at or radially outside of the frangible line.

[0067] The airfoil according to any of the preceding clauses, wherein the frangible airfoil portion extends from the tip along at least 10% but less than 50% of the span.

[0068] The airfoil according to any of the preceding clauses, wherein the one or more SMA inserts include nickel-titanium (NiTi), NiTi-based alloys, and combinations thereof.

[0069] The airfoil according to any of the preceding clauses, wherein the one or more SMA inserts include an insert geometry to facilitate mechanical interlocking between the one or more SMA inserts and the plurality of composite material layers.

[0070] The airfoil according to any of the preceding clauses, wherein the one or more SMA inserts include one or more sides that taper from a top edge towards a midpoint in the thickness direction, spanwise direction, chordwise direction, or a combination thereof, one or more sides that taper from a bottom edge towards a midpoint in the thickness direction, spanwise direction, chordwise direction, and / or both.

[0071] The airfoil according to any of the preceding clauses, wherein the one or more SMA inserts include a span height of from about 0.1 inches to about 6.0 inches.

[0072] The airfoil according to any of the preceding clauses, wherein the one or more SMA inserts include a plurality of SMA inserts arranged sequentially in the chordwise direction.

[0073] The airfoil according to any of the preceding clauses, wherein the plurality of composite material layers include carbon-based composite material layers.

[0074] The airfoil according to any of the preceding clauses, including a metallic leading edge shroud covering at least a portion of the leading edge.

[0075] The airfoil according to any of the preceding clauses, wherein the airfoil is a fan blade of a gas turbine engine.

[0076] A gas turbine engine includes: a turbine including a compressor, a combustor, and a turbine arranged in a series flow order; and a plurality of airfoils rotatable with the turbine, each of the plurality of airfoils defining a span extending in a spanwise direction between a root and a tip; a chordwise direction extending between a leading edge and a trailing edge; a thickness direction extending between a pressure side and a suction side; and a frangible line at a location along the span, each airfoil including: a frangible airfoil portion extending between the leading edge and the trailing edge and along the span between the tip and the frangible line, the frangible airfoil portion including a plurality of composite material layers and one or more shape memory alloy (SMA) inserts; and a remaining airfoil portion extending along the span from the frangible line to the root, the remaining airfoil portion including a plurality of composite material layers, wherein after an event that creates an imbalance, the frangible airfoil portion deforms or partially detaches relative to the remaining airfoil portion at or radially outside of the frangible line.

[0077] The gas turbine engine according to any of the preceding clauses further includes a fan section including a plurality of airfoils configured as fan blades.

[0078] The gas turbine engine according to any of the preceding clauses, wherein the frangible airfoil portion extends from the tip for at least 10% but less than 50% of the span.

[0079] The gas turbine engine according to any of the preceding clauses, wherein the one or more SMA inserts include nickel-titanium (NiTi), NiTi-based alloys, and combinations thereof.

[0080] The gas turbine engine according to any of the preceding clauses, wherein the one or more SMA inserts include an insert geometry to facilitate mechanical interlock between the one or more SMA inserts and the plurality of composite material layers.

[0081] The gas turbine engine according to any of the preceding clauses, wherein the one or more SMA inserts include one or more sides that taper from a top edge towards a midpoint in the thickness direction, spanwise direction, chordwise direction, or a combination thereof, one or more sides that taper from a bottom edge towards a midpoint in the thickness direction, spanwise direction, chordwise direction, and / or both.

[0082] The gas turbine engine according to any of the preceding clauses, wherein the one or more SMA inserts include a span height of from about 0.1 inches to about 6.0 inches.

[0083] The gas turbine engine according to any of the preceding clauses, wherein the plurality of composite material layers include carbon-based composite material layers.

[0084] A method for forming a frangible airfoil, the frangible airfoil defining a span extending in a spanwise direction between a root and a tip; a chordwise direction extending between a leading edge and a trailing edge; and a thickness direction extending between a pressure side and a suction side, the method comprising: laying a plurality of composite material layers to form a remaining airfoil portion; laying a plurality of composite material layers and one or more SMA inserts to form a frangible airfoil portion; and processing the plurality of composite material layers and the one or more SMA inserts to form the frangible airfoil.

[0085] The method according to any of the preceding clauses, wherein the one or more SMA inserts comprise nickel-titanium (NiTi), NiTi-based alloys, and combinations thereof.

[0086] This written description uses examples to disclose aspects of the present disclosure, including the best mode, and also enables those skilled in the art to practice aspects of the present disclosure, including making and using any device or system and performing any combined method. The patentable scope of the present 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 have structural elements that do not differ from the literal language of the claims or if they include equivalent structural elements that have insubstantial differences from the literal language of the claims.

Claims

1. An airfoil, the airfoil defining a span extending in a spanwise direction between a root and a tip; a chordwise direction extending between a leading edge and a trailing edge; a thickness direction extending between a pressure side and a suction side ； and a deformation line at a location along the spanwise direction, wherein, the airfoil comprises: a deformable airfoil portion extending between the leading edge and the trailing edge and along the span between a tip and the deformation line, the deformable airfoil portion comprising a plurality of composite material layers and one or more shape memory alloy - SMA inserts; and a remaining airfoil portion extending along the span from the deformation line to the root, the remaining airfoil portion comprising a plurality of composite material layers, wherein after an event that creates an imbalance, the deformable airfoil portion deforms or partially detaches relative to the remaining airfoil portion at or radially outside of the deformation line, wherein the one or more SMA inserts comprise one or more sides that taper from a bottom edge towards a midpoint along the thickness direction, the spanwise direction, the chordwise direction, or a combination thereof.

2. The airfoil according to claim 1, wherein, the deformable airfoil portion extends from the tip along at least 10% but less than 50% of the span.

3. The airfoil according to claim 1, wherein, the one or more SMA inserts comprise nickel - titanium (NiTi), NiTi - based alloys, and combinations thereof.

4. The airfoil according to claim 1, wherein, the one or more SMA inserts comprise an insert geometry to facilitate mechanical interlocking between the one or more SMA inserts and the plurality of composite material layers.

5. The airfoil according to claim 1, wherein, the one or more SMA inserts comprise one or more sides that taper from a top edge towards a midpoint along the thickness direction, the spanwise direction, the chordwise direction, or a combination thereof.

6. The airfoil according to claim 1, wherein, the one or more SMA inserts comprise a span height of from 0.1 inches to 6.0 inches.

7. The airfoil according to claim 1, wherein, the one or more SMA inserts comprise a plurality of SMA inserts arranged in sequence along the chordwise direction.

8. The airfoil according to claim 1, wherein, the plurality of composite material layers comprise carbon - based composite material layers.

9. The airfoil according to claim 1, wherein, comprises a metallic leading - edge shroud covering at least a portion of the leading edge.

10. The airfoil according to claim 1, wherein, the airfoil is a fan blade of a gas turbine engine.

11. A gas turbine engine, wherein, comprises: a turbine, the turbine comprising a compressor, a combustor, and a turbine arranged in a series - flow sequence; and Multiple airfoils that are capable of rotating with the turbine, each of the multiple airfoils defining a span that extends in the spanwise direction between a root and a tip; a chordwise direction that extends between a leading edge and a trailing edge; a thickness direction that extends between a pressure side and a suction side; and a deformation line at a location along the span, each airfoil comprising: a deformable airfoil portion that extends between the leading edge and the trailing edge and along the span between a tip and the deformation line, the deformable airfoil portion including a plurality of composite material layers and one or more shape memory alloy - SMA inserts; and a remaining airfoil portion that extends along the span from the deformation line to the root, the remaining airfoil portion including a plurality of composite material layers, wherein after an event that creates an imbalance, the deformable airfoil portion deforms or partially detaches relative to the remaining airfoil portion at or radially outside of the deformation line; wherein the one or more SMA inserts include one or more sides that taper from a bottom edge to a midpoint in the thickness direction, the spanwise direction, and the chordwise direction.

12. The gas turbine engine according to claim 11, wherein, it further includes a fan section that includes the multiple airfoils configured as fan blades.

13. The gas turbine engine according to claim 11, wherein, the deformable airfoil portion extends from the tip for at least 10% but less than 50% of the span.

14. The gas turbine engine according to claim 11, wherein, the one or more SMA inserts include nickel - titanium (NiTi), NiTi - based alloys, and combinations thereof.

15. The gas turbine engine according to claim 11, wherein, the one or more SMA inserts include an insert geometry to facilitate mechanical interlocking between the one or more SMA inserts and the plurality of composite material layers.

16. The gas turbine engine according to claim 11, wherein, the one or more SMA inserts include one or more sides that taper from a top edge to a midpoint in the thickness direction, the spanwise direction, the chordwise direction, or a combination thereof.

17. The gas turbine engine according to claim 11, wherein, the one or more SMA inserts include a span height of 0.1 inches to 6.0 inches.

18. The gas turbine engine according to claim 11, wherein, the plurality of composite material layers include carbon - based composite material layers.

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