Airfoil arrangement for a gas turbine engine utilizing shape memory alloys

Abstract

A wing profile arrangement for a gas turbine engine can include a support device that uses a shape memory alloy to support and control a wing profile. The support device can be formed as part of a fan blade. The arrangement can be configured to reduce the overall weight and size of the gas turbine engine.

Description

Technical Field

[0001] This disclosure relates to an airfoil arrangement for a gas turbine engine. Background Technology

[0002] Gas turbine engines typically include a turbine and a rotor assembly. Gas turbine engines, such as turbofan engines, can be used for aircraft propulsion. In the case of a turbofan engine, the rotor assembly can be configured as a fan assembly. The turbine can include a spool arrangement. For example, the spool arrangement can include a high-pressure, high-speed spool and a low-pressure, low-speed spool. The combustion section of the turbine receives pressurized air, which mixes with fuel in the combustion chamber and burns to produce combustion gases. The combustion gases are supplied to the spool arrangement. For example, the combustion gases can be first supplied to a high-pressure turbine on a high-pressure spool to drive the high-pressure spool, and then subsequently supplied to a low-speed turbine on a low-speed spool to drive the low-speed spool.

[0003] In a turbofan engine, the fan assembly typically includes a fan and a fan housing. The fan usually comprises multiple airfoils or fan blades extending radially outward from the center hub and / or disk. During certain operations, the fan blades may rub against or otherwise contact the fan housing, transferring load to the fan housing and potentially causing damage to the fan blades, the fan housing, or both.

[0004] In some configurations, the fan housing may include honeycomb or groove-filled materials, configured to reduce load transfer to and through the fan housing. However, this approach is typically costly. Furthermore, it can result in larger, heavier, and / or less efficient fan housings.

[0005] Therefore, an improved airfoil is needed that reduces the load transfer to the fan housing when the rotor blades rub against the fan housing. Attached Figure Description

[0006] Figure 1 It is a cross-sectional view of a gas turbine engine;

[0007] Figure 2 This is a perspective view of the fan assembly of a gas turbine engine;

[0008] Figure 3 This is a perspective view of the fan blades of a gas turbine engine;

[0009] Figure 4 This is a schematic cross-sectional view of an embodiment of a fan blade for a gas turbine engine;

[0010] Figure 5 This is a schematic cross-sectional view of an embodiment of a fan blade for a gas turbine engine;

[0011] Figure 6This is a schematic cross-sectional view of an embodiment of a fan blade for a gas turbine engine;

[0012] Figure 7 This is a schematic diagram of the detachable components of the airfoil used in a gas turbine engine;

[0013] Figure 8 This is a schematic diagram of the matrix arrangement of components in a gas turbine engine;

[0014] Figure 9 yes Figure 7 A schematic cross-sectional view of the airfoil component; and

[0015] Figure 10 It is a perspective view of the fan blades of a gas turbine engine in a deformed state.

[0016] Other aspects and advantages of the embodiments disclosed herein will become apparent upon consideration of the following detailed description, in which similar or identical structures may have similar or identical reference numerals. Detailed Implementation

[0017] Reference will now be made in detail to the present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. Detailed description uses numerals and letter names to refer to features in the drawings. Similar or analogous reference numerals in the drawings and description are used to denote similar or analogous portions of the invention.

[0018] The term “exemplary” is used herein to mean “as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as superior to or better than other implementations.

[0019] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0020] The terms "front" and "rear" refer to relative positions within a gas turbine engine or carrier, and are used with reference to the normal operating posture of the gas turbine engine or carrier. For example, "front" refers to the position closer to the engine inlet, and "rear" refers to the position closer to the engine nozzle or exhaust port.

[0021] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction in which the fluid flows.

[0022] The terms “connection,” “fixation,” “attachment,” etc., refer to direct connection, fixation, or attachment, as well as indirect connection, fixation, or attachment via one or more intermediate components or features, unless otherwise specified herein.

[0023] The singular forms “a,” “one,” and “the” include plural references unless the context clearly specifies otherwise.

[0024] The approximate language used throughout the specification and claims is applied to modify any quantitative expression that allows for variation without altering its associated essential function. Therefore, values ​​modified by one or more terms such as "approximately," "approximately," and "essentially" are not limited to specified exact values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to configure or manufacture the component and / or system. For example, approximate language may refer to a range of 1%, 2%, 4%, 10%, 15%, or 20%.

[0025] Scope limitations are combined and interchanged herein, and unless the context or language otherwise indicates otherwise, such scopes are identified and include all subscopes contained herein. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0026] In some embodiments, one or more components of the gas turbine engine described below may be manufactured or formed using any suitable process, such as additive manufacturing processes, for example, 3D printing. The use of such a process can allow such components to be formed integrally, as a single monolithic part, or as any suitable number of sub-parts. In particular, additive manufacturing processes can allow such components to be formed integrally and include a variety of features that are not possible using existing manufacturing methods. For example, the additive manufacturing methods described herein are capable of producing airfoils, turbine blades and rotors, compressor blades and rotors, and / or fan blades. Such components may have unique features, configurations, thicknesses, materials, densities, fluid channels, manifolds, and mounting structures that may be impossible or impractical using existing manufacturing methods. Some of these features are described herein.

[0027] As used herein, the term "additive manufacturing" or "additive manufacturing technology or process" generally refers to a manufacturing process in which consecutive layers of material are provided on one another to "build" a three-dimensional part layer by layer. Consecutive layers are often fused together to form a monolithic part that may have multiple integral sub-parts. Although additive manufacturing technology is described herein as capable of manufacturing complex objects by building them point-by-point, layer-by-layer, typically in a vertical direction, other manufacturing methods are possible and within the scope of this subject matter. For example, while the discussion herein relates to the addition of material to form consecutive layers, those skilled in the art will understand that the methods and structures disclosed herein can be practiced with any additive manufacturing technology or manufacturing technique. For example, embodiments of the invention may use layer-additive processes, layer-subtractive processes, or hybrid processes.

[0028] In a gas turbine engine, the fan assembly is typically connected to a spool arrangement. For example, the rotor of the fan assembly may be coupled to a low-speed spool. The function of the fan assembly is to direct air toward the turbine. A portion of the air from the fan assembly enters the turbine to move through at least one compression stage, while another portion of the air bypasses at least a portion of the turbine. The air initially bypassing the turbine may enter the turbine at a stage furthest from the fan assembly, and the air initially entering the turbine may bypass the other stages of the turbine.

[0029] The rotor's rotation, together with the fan's airfoil, accelerates air toward the turbine. The fan casing extends radially around the far end of the airfoil and can be used to improve performance by reducing blade tip losses. The performance of the fan assembly typically requires maintaining a very small clearance between the fan and the fan casing.

[0030] To achieve high performance, the airfoil's tip rotates very close to the fan casing. During certain events, such as foreign object ingestion in a gas turbine engine, the airfoil may come into contact with the fan casing. In these events, the blades may be released from the engine, a condition known as fan blade detachment (“FBO”). This situation poses a threat to the gas turbine engine and its surrounding environment.

[0031] One object of this disclosure is to describe various airfoil arrangements for managing adverse events such as foreign object ingestion (FBO). The airfoil of the fan assembly can be configured with support devices that utilize shape memory alloys to support and control the airfoil during these events. The described airfoil arrangements can advantageously provide a reliable, lightweight, and compact configuration for managing these adverse events and reducing and controlling FBO events.

[0032] To mitigate damage and potentially reduce the need for immediate repairs, some fan components are equipped with abrasive materials. For example, abrasive materials may be provided inside the airfoil and / or the fan housing. The fan housing may be provided with grooves arranged circumferentially within the fan housing and typically aligned with the fan's rotation. The abrasive groove filler may be positioned within the grooves and configured to dissipate some of the impact energy from the contact between the fan and the fan housing. Such a configuration may require additional weight and space to effectively manage these events.

[0033] A clearance device for an airfoil can be provided, for example, in a manner capable of maintaining a controlled and consistent clearance between the airfoil and the fan housing, which can reduce cost and weight as well as load transfer to the surrounding housing. The airfoil may define a tip portion, a root portion, and an intermediate portion between the tip and root portions. The intermediate portion may include a clearance device or a portion thereof. In some embodiments, the clearance device may extend into some or all of the tip and / or root portions. The root portion may extend along the wingspan S between the intermediate portion and the airfoil root. The tip portion may extend along the wingspan S between the intermediate portion and the airfoil tip. One or more transition portions may be provided. For example, a root transition portion may be disposed between the intermediate and root portions. A tip transition portion may be disposed between the intermediate and tip portions. One or more transition portions may be configured to include the intermediate portion and at least one of the root and tip portions. For example, the transition portion may define an overlap between the intermediate portion and another portion. In one embodiment, the tip transition portion defines an overlap of the airfoil thickness along the tip and intermediate portions. In another embodiment, the root transition portion defines the overlap of the airfoil thickness along the root portion and the middle portion.

[0034] The embodiments generally shown and described herein enable controlled and consistent reduction of the airfoil (e.g., fan blade) span after a damage event, such as hard friction against the surrounding fan housing. These embodiments enable the airfoil to fold and / or retract at a desired span, for example, to reduce load transfer to the surrounding housing. These embodiments can further enable the airfoil to fold and / or retract, thereby reducing excessive or extreme fan imbalances after a damage event, such as airfoil release, foreign object damage (e.g., bird strikes, icing, etc.), or loss of lubrication or damping of bearing assemblies.

[0035] Now refer to the attached diagram, Figure 1 This is a cross-sectional schematic diagram of an embodiment of the gas turbine engine 10. The illustrated embodiment can be used within an aircraft according to aspects of this subject matter. More specifically, for Figure 1 In one embodiment, the gas turbine engine is a high-bypass turbofan jet engine. The illustrated gas turbine engine 10 has a longitudinal or axial centerline axis 12 extending through it in an axial direction A, for reference. The gas turbine engine 10 further defines a radial direction R extending from the centerline 12. Although an exemplary turbofan embodiment is shown, it is contemplated that this disclosure is equally applicable to general turbomachinery, such as open rotor, turbine shaft, turbojet, or turboprop engine configurations, including marine and industrial turbine engines and auxiliary power units.

[0036] Generally, the gas turbine engine 10 includes a turbine (generally indicated by reference numeral 14) and a fan assembly 16 positioned upstream thereon. The turbine 14 may include a housing 18. The housing 18 may have a generally tubular configuration and / or define an annular inlet 20. Furthermore, the housing 18 may further enclose and support a low-pressure (“LP”) compressor 22 for increasing the pressure of air entering the turbine 14 to a first pressure level. A high-pressure (“HP”) compressor may be included. For example, a multi-stage axial flow configuration of the HP compressor 24 may receive pressurized air from the LP compressor 22 and further increase the pressure of such air.

[0037] The pressurized air exiting the HP compressor 24 can flow to the combustor 26, where fuel is injected into the pressurized air stream, and the resulting mixture is combusted within the combustor 26. Combustion products 60 can be directed from the combustor 26 along the hot gas path 67 of the gas turbine engine 10 to the high-pressure (“HP”) turbine 28, which drives the HP compressor 24 via the high-pressure (“HP”) shaft or spool 30. The combustion products 60 can also flow to the low-pressure (“LP”) turbine 32, which drives the LP compressor 22. The LP turbine 32 can also drive the fan assembly 16. For example, the LP turbine 32 can drive the LP compressor 22 and / or the fan assembly 16 via the low-pressure (“LP”) shaft or spool 34. The LP shaft 34 can be substantially coaxial with the HP shaft 30. After driving turbines 28 and 32, the combustion products 60 can be discharged from the turbine 14 via exhaust nozzles 36, which provides propulsive thrust.

[0038] The turbine section of the gas turbine engine 10 may include one or more non-rotating components. For example, the HP turbine 28 may include multiple HP turbine stator blades 29. The LP turbine may include multiple LP turbine stator blades 33. The HP turbine stator blades 29 and LP turbine stator blades may be configured to span the hot gas flow path 67 in the radial direction R. The HP turbine stator blades 29 and LP turbine stator blades may be non-rotating relative to the housing 18 about the longitudinal centerline axis 12 and may be used to support the housing 18. The HP turbine stator blades 29 and LP turbine stator blades 33 may be adjustable or adjustable to increase the performance of the gas turbine engine 10 by controlling the flow rate and pressure.

[0039] The turbine section of the gas turbine engine 10 may include multiple rotor blades. For example, the HP turbine 28 may include multiple HP turbine rotor blades 31. The LP turbine 32 may include multiple LP turbine rotor blades 35. The HP turbine rotor blades 31 and LP turbine rotor blades 35 may rotate relative to the housing 18. The relative rotation between the HP turbine stator blades 29 and HP turbine rotor blades 31, and between the LP turbine stator blades 33 and LP turbine rotor blades 35, may be adjusted or adjustable. For example, the pitch of the HP turbine stator blades 29, HP turbine rotor blades 31, LP turbine stator blades 33, and / or LP turbine rotor blades may be adjusted or adjustable to optimize flow through the hot gas path 67.

[0040] The performance of the gas turbine engine 10 can be improved by minimizing the clearances between components such as impellers and blades and duct components that can rotate relative to them. For example, minimizing and maintaining the clearance between turbine rotor blades 31, 35 and housing 18 can be used to improve engine performance. Minimizing and maintaining the clearance between turbine stator blades 29, 33 and rotatable components (such as LP shaft or spool 34) or components coupled thereto can also be used to improve engine performance.

[0041] like Figure 1 and 2 As shown, the fan assembly 16 of the gas turbine engine 10 may include a fan rotor 38. For example, the rotational, axial-flow configuration of the fan rotor 38 may be configured to be surrounded or guided by the fan housing 40. In some embodiments, the LP shaft 34 may be directly connected to the fan rotor 38 and / or the rotor disk 64, such as in a direct drive configuration. In some configurations, 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. The reduction gear 37 may be included between any suitable shaft / spindle within the gas turbine engine 10, as needed or required.

[0042] Those skilled in the art will understand that the fan housing 40 can be configured to be supported relative to the turbine 14. For example, the fan housing 40 can be supported by a plurality of substantially radially extending, circumferentially spaced supports. In one embodiment, a plurality of outlet guide vanes 42 are provided as such supports. Thus, the fan housing 40 may surround the fan rotor 38. The fan housing may include an outer nacelle 21 configured for aerodynamic flight characteristics, such as drag reduction. The fan rotor 38 may be connected to a plurality of fan blades 44. For example, a disk 64 may be configured to be coupled to a plurality of fan blades 44. Furthermore, a downstream portion 46 of the fan housing 40 may extend over an external portion of the turbine 14 to define an airflow duct 48. The airflow duct 48 may be configured as an auxiliary or bypass airflow duct. The airflow duct 48 may be configured to provide additional propulsive jet thrust and / or cooling effects for the gas turbine engine 10, for example, using fan nozzles 47.

[0043] The hub 65 can be configured to cover the disc 64. For example, the hub 65 can rotate together with the disc 64 and have an aerodynamic profile to control airflow through the multiple fan blades 44.

[0044] A pitch control mechanism (“PCM”) 61 is provided. For example, the PCM 61 may be operatively coupled to a plurality of fan blades 44 and configured to change the pitch of at least some of the fan blades. Individual fan blades 44 may rotate relative to disk 39 about pitch axis P and are controlled by the PCM 61. The PCM 61 may be configured to change the pitch of the fan blades 44 individually or uniformly.

[0045] However, it should be understood that the gas turbine engine 10 is provided as an example only. In other exemplary embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, in other exemplary embodiments, the gas turbine engine 10 may include a fixed-pitch fan, may be configured to directly drive the engine (i.e., may not include the gearbox 37), and may include any suitable or desired number or configuration of shafts or spools, compressors, and / or turbines, etc.

[0046] Still referencing Figure 1During operation of the gas turbine engine 10, it should be understood that the initial airflow 50 can enter the gas turbine engine 10 through the initial inlet 52 of the fan housing 40. The initial airflow 50 then passes through the fan blades 44. Downstream of the fan blades 44, the initial airflow 50 can be split. For example, the initial airflow 50 can be split into a first compressed airflow 54 moving through the airflow duct 48 and a second compressed airflow 56 entering the LP compressor 22. The second compressed airflow 56 can travel along the core airflow path 15. When the second compressed airflow 56 enters the HP compressor 24, its pressure can subsequently increase, becoming a third compressed airflow 58. After mixing with fuel and burning within the combustor 26, the combustion products 60 can exit the combustor 26 and flow through the HP turbine 28. Thereafter, the combustion products 60 can flow through the LP turbine 32 and exit the exhaust nozzle 36 to provide thrust to the gas turbine engine 10.

[0047] For reference Figure 2 and Figure 3 An exemplary embodiment of the airfoil 62 is provided against the background of the fan blade 44. As described above. Figure 2 The components of fan assembly 16 are depicted. Figure 3 A single fan blade 44 for use in fan assembly 16 is depicted. Although airfoil 62 is shown as part of fan blade 44, it should be understood that the following discussion of airfoil 62 can also be applied to another airfoil embodiment, such as stator blades or rotor blades of compressors 22, 24 and / or turbines 28, 32 (see Figure 1 As shown in the figure, each fan blade 44 extends radially outward along the wingspan S, which defines the spanwise direction from the airfoil root 64 to the airfoil tip 66. The wingspan S is defined by the distance along the spanwise centerline of the fan blade 44 from the airfoil root 64 to the airfoil tip 66 in the spanwise direction. The pressure side 68 and suction side 70 of the airfoil 62 extend from the leading edge 72 of the fan blade 44 to the trailing edge 74 and extend along the wingspan S between the airfoil root 64 and the airfoil tip 66. Furthermore, it should be appreciated that the airfoil 62 can define a chordwise direction along a chord C at each point along the wingspan S and extend between the leading edge 72 and the trailing edge 74. The chord C is defined by the distance along the chord centerline in the chordwise direction, which is generally orthogonal to the spanwise centerline of the fan blade 44 from the leading edge 72 to the trailing edge 74. Furthermore, the chord C can vary along the wingspan S of the airfoil 62. For example, in the depicted embodiment, the chord C increases along the wingspan S toward the airfoil tip 66. However, in other embodiments, the chord C may be substantially constant over the entire wingspan S or may decrease from the airfoil root 64 toward the airfoil tip 66.

[0048] like Figure 3Specifically, it is shown that the airfoil 62 may be defined with a thickness extending in the thickness direction T between the pressure side 68 and the suction side 70 at each point along the wingspan S. In some embodiments, the thickness may be substantially constant over the entire wingspan S of the airfoil 62. In other embodiments, the airfoil 62 may be defined with a variable thickness between the airfoil root 64 and the airfoil tip 66. For example, the thickness may generally decrease along the wingspan S toward the airfoil tip 66. Additionally, the airfoil 62 may be defined with a substantially constant thickness along chord C at each point along the wingspan S. Alternatively, in other embodiments, at least one point along the wingspan S of the airfoil 62 may be defined with a variable thickness along chord C. For example, the airfoil 62 may be defined with a maximum thickness at the location of chord C at each point along the wingspan S.

[0049] The airfoil 62 may also include multiple surfaces. For example, the airfoil 62 may include a first outer surface, a second outer surface opposite to the first outer surface, and a support device 84 formed of a shape memory alloy and disposed between the first and second outer surfaces. The first outer surface may, for example, be the pressure side 68 of the airfoil 62, and the second outer surface may be the suction side 70 of the airfoil 62, such as... Figure 2 and 3 As shown. The support device 84 can be enclosed within a housing, which is at least partially composed of a first outer surface, a second outer surface, a leading edge 72 surface connecting the first and second outer surfaces, and a trailing edge 74 surface spaced apart from the leading edge surface and connecting the first and second outer surfaces.

[0050] Figure 3 The fan blade 44 may have a hollow configuration. In another example, the fan blade 44 may include one or more cavities for purposes such as cooling. Alternatively, the fan blade 44 may be a solid structure.

[0051] In the illustrated embodiment, each fan blade 44 includes an integral component with an axial dovetail 76 having a pair of opposing pressure surfaces 78 leading to the transition section 80. When mounted within the fan rotor disk 39, as... Figure 2 As shown, the dovetail tenon 76 is provided in the dovetail groove of the fan rotor disk 39, thereby attaching the fan blade 44 as part of the fan rotor 38.

[0052] In one embodiment, the airfoil 62 may include at least one composite layer (as described below, such as...) Figure 7 (One of the layers 82 shown). For example, the airfoil 62 may be formed at least partially of a ceramic matrix composite. More specifically, in some embodiments, the airfoil 62 may be formed partially of one or more composite layers 82 configured as ceramic matrix composite prepreg layers.

[0053] Composite materials may include, but are not limited to, metal matrix composites (MMC), polymer matrix composites (PMC), or ceramic matrix composites (CMC). Composite materials, such as those used in airfoil 62, typically include fiber reinforcement materials, such as polymers, ceramics, or metals, embedded in a matrix material. The reinforcement material serves as the load-bearing component of the composite material, while the matrix of the composite material serves to bond the fibers together and as a medium for transferring and distributing externally applied stress to the fibers.

[0054] Exemplary CMC materials may include silicon carbide (SiC), silicon, silica, or alumina matrix materials, and combinations thereof. Ceramic fibers may be embedded in the matrix, such as oxidation-stabilized reinforcing fibers, including monofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6), and rovings and yarns comprising silicon carbide (e.g., Nippon Carbon's). UbeIndustries And Dao Corning ), aluminum silicate (e.g., Nextel's 440 and 480) and chopped whiskers and fibers (e.g., Nextel's 440 and 480) The preform may include, for example, ceramic particles (e.g., Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For instance, in some embodiments, fiber bundles comprising a ceramic refractory coating may be formed as reinforcing bands, such as unidirectional reinforcing bands. Multiple bands may be stacked together (e.g., as a composite layer 82) to form a preformed part. The fiber bundles may be impregnated with a slurry composition before or after the formation of the preform. The preform may then be subjected to heat treatment, such as curing or burn-out, to produce a significant amount of carbon residue in the preform, and subsequent chemical treatment, such as infiltration with a silicon melt, to obtain a part formed from a CMC material having the desired chemical composition.

[0055] The airfoil 62 may include various metallic components. For example, the leading edge 72, trailing edge 74, tip 66, and / or root 64 may be formed of at least one metallic layer. For example, reinforcing and / or impact-resistant features may be provided in these components of the airfoil 62. These reinforcing and / or impact-resistant features may be made of metals such as titanium. Titanium may be a titanium alloy, such as Ti-6Al-4V. Alternatively, ferrous alloys, steel, and / or aluminum alloys may be provided. For example, the reinforcing and / or impact-resistant features may be made of stainless steel such as 15-5PH.

[0056] The gas turbine engine 10 may include a support device 84 configured to support another component. For example, the support device 84 may be provided to support and control the deflection of components of the gas turbine engine 10, such as airfoil 62. Such airfoil 62 may be provided on various components of the gas turbine engine 10. For example, one or more turbine rotor blades 31, 35; turbine stator blades 29, 33; and / or fan blades 44 may be provided with such a support device 84.

[0057] More specifically, now refer to Figure 4 A schematic cross-sectional view of the airfoil 62 is provided, including a support device 84. The support device 84 may be manufactured using a shape memory alloy (“SMA”). The support device 84 may include one or more SMA components or features, or may be formed entirely of SMA.

[0058] The shape memory alloy (SMA) described with reference to support device 84 can be a metallic alloy undergoing a temperature-dependent, solid-state, microstructure phase transformation. This configuration of support device 84 can facilitate a change from one physical shape to another. The temperature at which the phase transformation occurs is commonly referred to as the critical or transformation temperature of the alloy. Support device 84 can be made of various SMA materials. For example, support device 84 can include a titanium-nickel alloy (commonly referred to as a nickel-titanium alloy). Support device 84 may optionally or additionally include a high-temperature SMA. For example, support device 84 can be made of a Ru alloy that alloys with Nb or Ta. In one embodiment, a high-temperature SMA can facilitate support device 84 to produce a shape memory transformation temperature ranging from about 20 degrees Celsius to about 1400 degrees Celsius. The transformation temperature of support device 84 can be tuned for a specific application. For example, support device 84 used in fan blades can benefit from a relatively low transformation temperature, while support device 84 used in HP turbine rotor blades 31 can benefit from a relatively high transformation temperature.

[0059] In some embodiments, the support device 84 may include an SMA material as a major component, in an amount greater than 50 weight percent (“wt.%”) of the support device 84. In some embodiments, the support device 84 may be substantially composed of an SMA material (e.g., at least 90 wt.%, such as at least 95 wt.%, such as 100 wt.%). An SMA material is typically an alloy capable of recovering its original shape after deformation. For example, an SMA material can define a hysteresis effect, where the loading path on a stress-strain diagram differs from the unloading path on a stress-strain diagram. Therefore, an SMA material can provide improved hysteresis damping compared to conventional elastic materials. Furthermore, an SMA material can serve as a lightweight, solid-state alternative to conventional actuators. For example, certain SMA materials can be heated to restore a deformed SMA to its pre-deformation shape. An SMA material can also provide varying stiffness in a predetermined manner in response to certain temperature ranges. The stiffness variation of shape memory alloys is due to temperature-dependent, solid-state, microstructural phase transitions, enabling the alloy to change from one physical shape to another. The stiffness variation of SMA materials can be produced by processing and annealing preforms of the alloy at or above the temperature at which the solid-state, microstructure phase transformation of the shape memory alloy occurs. The temperature at which this phase transformation occurs is generally referred to as the critical temperature or transformation temperature of the alloy. In the manufacture of a support device 84 (also called an SMA support device 84) that includes an SMA designed to change stiffness during operation of the gas turbine engine 10, the support device 84 may be formed to have an 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.

[0060] Some shape memory alloys used in this article are characterized by temperature-dependent phase transformations. These phases include martensitic and austenitic phases. Martensitic phase generally refers to the low-temperature phase, while austenitic phase usually refers to the higher-temperature phase. Martensitic phase is generally more deformable, while austenitic phase is generally less deformable. When a shape memory alloy is in the martensitic phase and heated above a certain temperature, it begins to transform into the austenitic phase. The temperature at which this phenomenon begins is called the austenitic initiation temperature (As). The temperature at which this phenomenon completes is called the austenitic termination temperature (Af). When a shape memory alloy in the austenitic phase cools, it begins to transform into the martensitic phase. The temperature at which this transformation begins is called the martensitic initiation temperature (Ms). The temperature at which the transformation to the martensitic phase is completed is called the martensitic termination temperature (Mf). As used herein, the term "transformation temperature" without further qualifiers can refer to either the martensitic transformation temperature or the austenitic transformation temperature. Furthermore, "below the transformation temperature" without the modifiers "starting temperature" or "ending temperature" generally refers to a temperature below the martensite termination temperature, while "above the transformation temperature" without the modifiers "starting temperature" or "ending temperature" generally refers to a temperature above the austenite termination temperature.

[0061] In some embodiments, the support device 84 including the SMA can define a first stiffness at a first temperature and a second stiffness at a second temperature, wherein the second temperature is different from the first temperature. Furthermore, in some embodiments, one of the first or second temperature is below the transition temperature, while the other may be at or above the transition temperature. Thus, in some embodiments, the first temperature may be below the transition temperature and the second temperature may be at or above the transition temperature. In other embodiments, the first temperature may be at or above the transition temperature, while the second temperature may be below the transition temperature. Moreover, various embodiments of the SMA support device 84 described herein can be configured to have different first stiffnesses and different second stiffnesses at the same first and second temperatures.

[0062] Non-limiting examples of SMAs applicable to various embodiments of the SMA support device 84 described herein may include nickel-titanium (NiTi) and other nickel-titanium-based alloys, such as nickel-titanium hydrogen fluoride (NiTiHf) and nickel-titanium palladium (NiTiPd). However, it should be understood that other SMA materials may be equally applicable to this disclosure. For example, in some 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 can 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 (of martensitic and austenitic phases), resistance to oxidation and hot corrosion, the ability to change shape through repeated cycles, the ability to exhibit unidirectional or bidirectional shape memory effects, and / or many other engineering design criteria. Suitable shape memory alloy compositions that can be used in embodiments of this 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 may transform from austenite to martensite upon cooling.

[0063] Furthermore, SMA materials may also exhibit superelasticity. Superelasticity is typically characterized by recovery from large strains and may involve some dissipation. For example, the martensitic and austenitic phases of SMA materials can respond to mechanical stress as well as temperature-induced phase transformations. For instance, SMA can be loaded into the austenitic phase (i.e., above a certain temperature). Therefore, when a critical stress is reached, the material may begin to transform into the (twinned) martensite phase. Under continued loading and assumed isothermal conditions, the (twinned) martensite may begin to detwin, causing the material to undergo plastic deformation. If unloading occurs before plastic deformation, the martensite typically transforms back into austenite, and the material may recover its original shape by generating hysteresis.

[0064] Still referencing Figure 4 It is understandable that airfoil 62 can be combined with... Figure 3 In fan blade 44. For example, Figure 4 The airfoil 62 is depicted as a component including an aerodynamic working surface of fan blades 44.

[0065] Figure 4 Depicting Figure 3 A schematic cross-sectional view of a fan blade 44, including a support device 84. As depicted, the support device 84 is at least partially enclosed within the fan blade 44. For example, although not depicted, the support device 84 may include a coating or other separate layer bonded to another portion of the fan blade 44. In any case, the support device 84 may be inserted into the fan blade 44 during manufacturing or assembly of the fan blade 44, or embedded in the fan blade 44 during its formation.

[0066] Still referencing Figure 4 The airfoil 62 includes a fragile portion 86. The fragile portion 86 is configured to selectively break, disintegrate, or otherwise lose mass. For example, the fragile portion 86 of the fan blade 44 can be configured to preferentially break upon contact with the fan housing 40. In this way, the use of protective barriers or abrasive materials (e.g., groove fillers) in the fan housing 40 can be avoided or minimized. Specific configurations of the groove support 84, alone or in combination with the fragile portion 86, can facilitate this weight-reducing configuration.

[0067] The fragile portion 86 is disposed on the outer side of the root portion in the radial direction R. For example, in Figure 4 In the exemplary embodiment shown, the fragile portion 86 constitutes at least a portion of the tip 66 of the airfoil 62. In one embodiment, the entire outer surface of the tip 66 is composed of the fragile portion 86. The fragile portion 86 may generally surround at least a portion of the support device 84. In one embodiment, the fragile portion 86 is configured to permanently deform in response to an external load.

[0068] Still referencing Figure 4 The support device 84 is configured to connect two or more components of the fan blade 44. For example, in the depicted embodiment, the support device 84 extends through the root portion and connects the root portion to the fragile portion 86. The root portion of the airfoil 62 may be configured to selectively maintain structural integrity. For example, the fragile portion 86 may not extend to the root portion. The support device 84 may be permanently attached to the body portion 88 and the fragile portion 86. For example, the support device 84 may be embedded in the body portion and the fragile portion 86.

[0069] The support device 84 can move between a baseline position and a deflection position. For example, the airfoil 62 including the support device 84 may be subjected to disturbance or impact, causing the support device 84 to deflect. In this way, the support device 84 can take advantage of the advantageous properties of its shape memory alloy structure, as described above. Embodiments of the airfoil 62 configured with the support device 84 can be... Figure 3 As shown in the diagram. The deflection of the airfoil 62 can be directly translated into the deflection of the support device 84 to its deflection position. The deflection zone 90 can be defined in one or more components. This deflection zone can be defined across the chord C of the fan blade 44. In some embodiments, the support device 84 can be configured to control the deflection to one or more graded deflection portions or folding portions, as will be shown in the following reference. Figure 4-6 Detailed description is provided.

[0070] The support device 84 may be disposed within the fragile portion 86. For example, the support device 84 may be completely enclosed within the fragile portion 86. The support device 84 may also be integral with the fragile portion 86. For example, the support device 84 may be co-molded with the fragile portion 86. It should be understood that the support device 84 and the fragile portion 86 may be manufactured in the same process or in different processes. In one embodiment, one or both of the support device 84 and the fragile portion 86 are additively manufactured.

[0071] Still referencing Figure 4-6 The support device 84 further includes at least one folding portion F1-F4 configured to deform along the axial direction A. More specifically, the folding portion is configured to allow the outer portion of the fan blade 44 in the radial direction R to pivot or bend relative to the inner portion of the fan blade 44 in the radial direction R at the folding portion. For example, in the event of foreign object ingestion, the support device 84 may deflect toward the turbine 14 along with the airfoil 62 of the fan blade 44. In one embodiment, the deformation is temporary. For example, the SMA support device 84 may enable a return to an undeformed state by utilizing a phase change.

[0072] The support device 84 may be configured with a deformation threshold. For example, the support device 84 may be adjusted to resist folding under certain conditions. In one embodiment, the support device 84 is adjusted to remain undeformed during crosswinds, minor friction, and / or small to moderate flocked bird strikes. The support device 84 may be adjusted using a shape memory alloy to have a relatively small deflection during minor impact events below the deformation threshold and a relatively large deflection during heavy impact events above the deformation threshold. The deflection of the support device 84 relative to the impact force may be non-linear, for example, having a disproportionately large deflection per unit force above the deformation threshold. In this way, the support device 84 can deflect within a small range below the deformation threshold and within a large range above the deformation threshold.

[0073] At least one of the folded portions F1-F4 can be configured to cause deflection, deformation, and / or folding in various ways. For example, at least one of the folded portions F1-F4 can be configured as a structural weak point in the support device 84. In one embodiment, one or more grooves 92 are provided along the perimeter of the support device 84 to at least partially form at least one of the folded portions F1-F4. Optionally or additionally, the support device 84 may be formed of a thinner material at at least one of the folded portions F1-F4 and / or the material properties may be configured for localized structural weaknesses relative to adjacent portions of the support device 84.

[0074] At least one of the folding portions F1-F4 can be multiple folding portions. Each of the multiple folding portions F1-F4 can be spaced apart along the span S of the airfoil 62 and / or the fan blade 44. In one embodiment, each of the folding portions F1-F4 is configured identically.

[0075] Multiple folding portions F1-F4 may include a first folding portion F1 and a second folding portion F2 disposed between the root portion 64 of the airfoil 62 and the first folding portion F1. A third folding portion F3 and a fourth folding portion F4 may be further provided disposed between the root portion 64 and the second folding portion. Subsequent folding portions F1-F4 may be configured to resist folding to a greater extent than the preceding folding portion F1, which is disposed further outward in the radial direction R. For example, the second folding portion F2 may be configured to resist folding to a greater extent than the first folding portion F1. In one embodiment, the support device 84 tapers such that it thins along the chord C of the airfoil as it approaches its tip. Such a configuration naturally facilitates the aforementioned difference in relative folding resistance. For example, the first folding portion F1 may define a first foldable width W1 along the chord C of the airfoil 62, and the second folding portion F2 may define a second foldable width W2 along the chord C, the second foldable width being greater than the first foldable width.

[0076] The folded sections F1-F4 can be configured in various ways to achieve relative stiffness. For example, as... Figure 4 and 5As shown, at least one of the folding portions F1-F4 may be configured with a recess 92. In one embodiment, at least one of the folding portions F1-F4 is configured as at least one recess 92 in the support device 84. In another embodiment, opposing pairs of recesses 92 may be provided across the chord C of the support device. For example, each of the plurality of folding portions F1-F4 may include a pair of recesses 92 in the support device 84. Although the recesses 92 are depicted as curved or rounded recesses, it should be understood that other configurations of recesses may be employed. For example, one or more recesses 92 may be angular or polygonal in configuration. Optionally or additionally, other features may be employed for the recesses 92 to facilitate folding. For example, localized thickness reduction, perforation, or adjustment of material properties may be used to facilitate folding.

[0077] The support device 84 can be configured to undergo a phase transformation in response to an adverse event. For example, the support device 84 can undergo a phase transformation to the martensitic phase in response to an external load. The support device 84 can also be configured to return to the austenitic phase in response to the release of the external load. The transformation temperature can be an adjustable characteristic corresponding to the temperature conditions of the gas turbine engine. For example, during takeoff and high-power events, components of the gas turbine engine will experience relatively higher temperatures compared to cruise or low-power events.

[0078] In various embodiments, the transition temperature of the support device 84 may be adjusted or adjustable for intended use. For example, the transition temperature may be adjusted to fall within a temperature range experienced by the support device 84 during low-power and high-power events. In one embodiment, the support device 84 is configured for use in a fan blade 44 having a transition temperature of at least 100 degrees Celsius (100°C). In another embodiment, the support device 84 is configured for use in a fan blade 44 having a transition temperature between 20 and 100 degrees Celsius (20-100°C). In other embodiments, the transition temperature can be much higher. For example, in an embodiment where the support device 84 is configured for an HP turbine rotor blade 31, the transition temperature may be between 500 and 1000 degrees Celsius (500-1000°C) or may be greater than 1000 degrees Celsius (1000°C).

[0079] The support device 84 can be configured to extend over various dimensions of the airfoil 62. For example, the support device 84 can extend over 5%, 10%, 25%, 50%, or 75% of the wingspan S of the airfoil 62. Figure 4 As shown, the support device 84 may be defined with a support device length D. The support device 84 may also extend at 5%, 10%, 25%, 50%, or 75% of the chord of the airfoil 62. Similarly, as... Figure 4As shown, the support device 84 may define one or more support device widths C1, C2. To determine the relative widths, the support device widths C1, C2 may be compared with a first foldable width W1 and a second foldable width W2. In one embodiment, a portion of the support device 84 extends over at least fifty percent (50%) of the wingspan S. In another embodiment, a portion of the support device 84 extends over at least fifty percent (50%) of the chord.

[0080] Shape memory alloy components can be configured to deform to reduce the wingspan S of an airfoil. For example, a support device 84 formed of SMA can be used to reduce the wingspan S by deforming toward the turbine 14 as described above. The support device 84 can be formed entirely of SMA or can be formed partially of SMA.

[0081] The support device 84 may be additionally or optionally configured to regulate the mass release event of the fragile portion 86. In one embodiment, the support device 84 is at least partially configured to maintain integrity while the fragile portion 86 loses mass. Optionally or in conjunction with this embodiment, the support device 84 may be configured to break along with the fragile portion 86. For example, the support device 84 may include one or more fragile portions configured to selectively break along with a certain amount of the fragile portion 86.

[0082] As depicted in the figure, the fragile tip 86 can be positioned between the root 64 and the tip 66 of the fan blade 44. In this arrangement, as... Figure 4 and Figure 5 As shown in the embodiment, the tip 66 is defined as the outermost radial point of the fragile tip portion. As shown, the support device 84 connects the root 64 and the fragile portion 86.

[0083] It should be understood that in other embodiments, the fan blades 44 can be configured in any other suitable manner. For example, now refer to Figure 5 A cross-sectional schematic diagram according to another embodiment of the present disclosure is provided. As shown, an intermediate portion 94 is disposed in a fan blade 44. In the illustrated embodiment, the intermediate portion 94 is disposed between a root portion and a brittle tip portion. The intermediate portion 94 may comprise various materials. For example, the intermediate portion 94 may be formed of a shape memory alloy. The intermediate portion 94 may be configured to facilitate local deformation or elasticity, such as allowing local bending, compression, and / or expansion.

[0084] refer to Figure 6 The winglet 62 can also be equipped with an axial spring 97. For example... Figure 6As shown, an axial spring 97 is provided in the radial direction R inside the support 84. In one embodiment, the axial spring 97 is compressible to facilitate movement of the support 84 and thus reduce the wingspan S of the airfoil 92. The axial spring 97 may be formed of an SMA with reference to the support 84 in the various ways described above. The configuration employing the axial spring 97 may also include the fragile portion 86 as described above. The axial spring 97 may be configured to facilitate radially inward movement of at least a portion of the airfoil 62 toward the axial centerline 12. For example, the axial spring 97 may be deformable to facilitate radially inward movement of the support 84.

[0085] Figure 7 A schematic diagram of a partially disassembled airfoil 62 is depicted. As shown, the airfoil 62 may be provided with components that separate at least a portion of the support device 84. For example, an intermediate component 96 is shown disposed between multiple layers of the support device 84. For example, the support device 84 may include a first support device layer 84a and a second support device layer 84b. The intermediate component 96 may be disposed between the first support device layer 84a and the second support device layer 84b. The intermediate component 96 may include various materials. For example, the intermediate component 96 may be formed of epoxy resin or a polymer material such as polyurethane filler.

[0086] The support layers 84a and 84b can be configured in various ways as described above with reference to support 84. In one embodiment, the support layers 84a and 84b form an SMA matrix. For example, refer to... Figure 8 The schematic diagram of the matrix arrangement depicts a first support layer 84a that may include matrix components aligned with the wingspan S of the airfoil 62, and a second support layer 84b that may include matrix components aligned with the chord C of the airfoil 62. In one embodiment, each of the first and second support layers 84a, 84b includes a matrix of a first matrix component 98 and a second matrix component 99 intersecting the first matrix component. The first and second matrix components may be orthogonally arranged.

[0087] Figure 9 Depicting the assembly configuration Figure 7 A schematic cross-sectional view of the airfoil 62 component. Figure 9 The support device 84 shown may include SMA fibers embedded in a matrix composite material. When components such as intermediate component 96 or first support device layer 84a and second support device layer 84b are separate layers, they can be bonded to another portion of fan blade 44 by suitable processes such as welding or brazing. Intermediate component 96 may be inserted into fan blade 44 or manufactured using additive manufacturing. It should be understood that first support device layer 84a and second support device layer 84b can be embedded in any suitable matrix composite material, such as a metal matrix or a ceramic matrix.

[0088] Figure 10 Describing as Figure 3 The deflection position of fan blade 44 is shown. Due to the deflection, and... Figure 3 Compared to the fan blades shown, Figure 10 The wingspan S is reduced. When a support device 84 as described herein is provided, the reduction in wingspan S may be limited to the radially outer portion or the tip portion of the airfoil 62.

[0089] The gas turbine engine 10 may be provided with any of the features and elements shown and described herein. The illustrations of the embodiments described herein are intended to provide a general understanding of the structure of various embodiments. The illustrations are not intended to be a complete description of all elements and features of devices and systems utilizing the structures or methods described herein. Many other embodiments may be apparent to those skilled in the art upon reading this disclosure. Other embodiments may be utilized and derived from this disclosure, allowing structural and logical substitutions and changes to be made without departing from the scope of this disclosure. Furthermore, the illustrations are representative only and may not be drawn to scale. Some scales in the illustrations may be exaggerated, while others may be minimized. Therefore, this disclosure and the accompanying drawings should be considered illustrative rather than restrictive.

[0090] While this specification contains numerous details, these should not be construed as limiting the scope of the invention or its claimable content, but rather as descriptions of features specific to particular embodiments of the invention. Certain features described in the context of individual embodiments in this specification may also be implemented in combination in a single embodiment. Conversely, various features described in the context of a single embodiment may also be implemented individually or in any suitable sub-combination in multiple embodiments. Furthermore, although features may be described above as functioning in certain combinations and even initially claimed in this way, in some cases one or more features may be removed from the claimed combination, and the claimed combination may be for sub-combinations or variations thereof.

[0091] One or more embodiments of this disclosure may be referred to herein, individually and / or collectively, by the term "invention," merely for convenience and not intended to voluntarily limit the scope of this application to any particular invention or inventive concept. Furthermore, while specific embodiments have been illustrated and described herein, it should be understood that any subsequent arrangements designed to achieve the same or similar purpose may replace the specific embodiments shown. This disclosure is intended to cover any and all subsequent modifications or variations of the various embodiments. Combinations of the foregoing embodiments, as well as other embodiments not specifically described herein, will be apparent to those skilled in the art upon reading the specification.

[0092] The foregoing detailed description is intended to be illustrative rather than restrictive, and it should be understood that the following claims, including all equivalents, are intended to define the scope of the invention. The claims should not be construed as limited to the described order or elements unless stated otherwise. Therefore, all embodiments falling within the scope and spirit of the appended claims and their equivalents are claimed as part of the invention.

[0093] The following items provide further details:

[0094] On one hand, an airfoil capable of rotating about an axis to move air in an axial direction is provided, the airfoil comprising: a root portion having a root; a tip portion having a tip opposite to the root in a radial direction; a fragile portion disposed outside the root portion in the radial direction; and a support device comprising a shape memory alloy, wherein the support device connects the root portion to the fragile portion and is movable between a baseline position and a deflection position.

[0095] On the other hand, an airfoil is provided having a first outer surface; a second outer surface opposite to the first outer surface; and a support device formed of a shape memory alloy and disposed between the first and second outer surfaces, wherein the support device is movable between a baseline position and a deflection position.

[0096] One aspect provides an airfoil in which the support structure is entirely formed of shape memory alloy.

[0097] On the other hand, an airfoil is provided, wherein the support device is disposed within the fragile portion.

[0098] On the other hand, an airfoil is provided, wherein the support device is integrated with the fragile part.

[0099] On the other hand, an airfoil is provided, wherein the support device includes at least one folding portion configured to deform in the axial direction.

[0100] On the other hand, an airfoil is provided, wherein the airfoil defines a wingspan between the root and the tip, and wherein the at least one folded portion includes a plurality of folded portions spaced apart along the wingspan.

[0101] On the other hand, an airfoil is provided, wherein the plurality of folded portions include a first folded portion and a second folded portion, the first folded portion defining a first stiffness, the second folded portion defining a second stiffness, and the second folded portion being disposed between the root and the first folded portion, wherein the second stiffness is greater than the first stiffness.

[0102] On the other hand, an airfoil is provided, wherein the airfoil defines a leading edge, a trailing edge, and a chord between the leading edge and the trailing edge, wherein a first folding portion defines a first foldable width along the chord, wherein a second folding portion defines a second foldable width along the chord, and wherein the second foldable width is greater than the first foldable width.

[0103] On the other hand, an airfoil is provided, wherein at least one folded portion is configured as at least one recess in the support device.

[0104] On the other hand, an airfoil is provided, wherein the at least one folding portion comprises a plurality of foldable portions, and wherein each of the plurality of foldable portions comprises a pair of opposing recesses in the support device.

[0105] On the other hand, an airfoil is provided, wherein the support device is enclosed in a housing, the housing being formed at least in part by a first outer surface, a second outer surface, a leading edge surface connecting the first outer surface and the second outer surface, and a trailing edge surface spaced apart from the leading edge surface and connecting the first outer surface and the second outer surface.

[0106] On the other hand, an airfoil is provided, wherein the support device is configured to undergo a phase transformation to the martensitic phase in response to an external load; and to return to the austenitic phase in response to the release of the external load.

[0107] On the other hand, an airfoil is provided, further comprising a fragile portion substantially surrounding at least a portion of the support device, the fragile portion being configured to permanently deform in response to the external load.

[0108] On the other hand, an airfoil is provided, wherein the support device is further configured to have a transition temperature of at least one hundred degrees Celsius (100°C).

[0109] On the other hand, an airfoil is provided, further comprising: a tip; a root portion; and a fragile tip portion disposed between the root portion and the tip, wherein the support device connects the root portion and the fragile tip portion.

[0110] On the other hand, an airfoil is provided, further including an intermediate portion disposed between the root portion and the fragile tip portion, the intermediate portion being formed of the shape memory alloy.

[0111] On the other hand, an airfoil is provided, wherein the airfoil defines a wingspan, and wherein a portion of the support device extends over at least fifty percent (50%) of the wingspan.

[0112] On the other hand, an airfoil is provided, wherein the airfoil defines a chord, and wherein a portion of the support device extends over at least fifty percent (50%) of the chord.

[0113] On the other hand, an airfoil is provided, wherein the support device is configured to adjust the mass release event of the fragile tip portion.

Claims

1. An airfoil capable of rotating about an axis to move air in the axial direction, characterized in that, The airfoil includes: The root portion has a root; The tip portion has a tip that is radially opposite the root portion; A fragile portion, the fragile portion being disposed on the outer side of the root portion along the radial direction; and A support device comprising a shape memory alloy, wherein the support device connects the root portion to the fragile portion and is movable between a baseline position and a deflection position; and wherein the support device includes a folding portion configured to allow the outer portion of the airfoil in the radial direction to pivot or bend relative to the inner portion of the airfoil in the radial direction at the folding portion.

2. The airfoil according to claim 1, characterized in that, The support device is entirely formed of the shape memory alloy.

3. The airfoil according to claim 1, characterized in that, The support device is located within the fragile part.

4. The airfoil according to claim 3, characterized in that, The support device is integrated with the fragile part.

5. The airfoil according to claim 1, characterized in that, The support device includes at least one folding portion, which is configured to deform in the axial direction.

6. The airfoil according to claim 5, characterized in that, The airfoil defines a wingspan between the root and the tip, and wherein the at least one folded portion comprises a plurality of folded portions spaced apart along the wingspan.

7. The airfoil according to claim 6, characterized in that, The plurality of folded portions include a first folded portion and a second folded portion, the first folded portion defining a first stiffness, the second folded portion defining a second stiffness, and the second folded portion being disposed between the root and the first folded portion, wherein the second stiffness is greater than the first stiffness.

8. The airfoil according to claim 7, characterized in that, The airfoil defines a leading edge, a trailing edge, and a chord between the leading edge and the trailing edge, wherein the first folding portion defines a first foldable width along the chord, wherein the second folding portion defines a second foldable width along the chord, and wherein the second foldable width is greater than the first foldable width.

9. The airfoil according to claim 5, characterized in that, The at least one folded portion is configured as at least one recess in the support device.

10. The airfoil according to claim 9, characterized in that, The at least one folding portion includes a plurality of foldable portions, and each of the plurality of foldable portions includes a pair of opposing recesses in the support device.

11. An airfoil component, characterized in that, include: First outer surface; The second outer surface is opposite to the first outer surface; as well as A support device formed of shape memory alloy and disposed between a first outer surface and a second outer surface, wherein the support device is movable between a baseline position and a deflection position; The support device includes a folding portion configured to allow the outer portion of the airfoil in the radial direction to pivot or bend at the folding portion relative to the inner portion of the airfoil in the radial direction.

12. The airfoil according to claim 11, characterized in that, The support device is enclosed in a housing, which is at least partially formed by a first outer surface, a second outer surface, a leading edge surface connecting the first outer surface and the second outer surface, and a trailing edge surface spaced apart from the leading edge surface and connecting the first outer surface and the second outer surface.

13. The airfoil according to claim 11, characterized in that, The support device is configured as follows: In response to an external load, it undergoes a phase transformation to the martensitic phase; and It returns to the austenitic phase in response to the release of the external load.

14. The airfoil according to claim 13, characterized in that, It further includes a fragile portion substantially surrounding at least a portion of the support device, the fragile portion being configured to permanently deform in response to the external load.

15. The airfoil according to claim 13, characterized in that, The support device is further configured to have a transition temperature of at least 100 degrees Celsius (100°C).

16. The airfoil according to claim 11, characterized in that, Further includes: Tip; Root part; and A fragile tip portion is disposed between the root portion and the tip, wherein the support device connects the root portion and the fragile tip portion.

17. The airfoil according to claim 16, characterized in that, It further includes an intermediate portion disposed between the root portion and the fragile tip portion, the intermediate portion being formed of the shape memory alloy.

18. The airfoil according to claim 16, characterized in that, The airfoil defines the wingspan, and a portion of the support extends over at least fifty percent (50%) of the wingspan.

19. The airfoil according to claim 16, characterized in that, The airfoil defines a chord, and a portion of the support device extends over at least fifty percent (50%) of the chord.

20. The airfoil according to claim 16, characterized in that, The support device is configured to regulate the mass release event of the fragile tip portion.

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