Turbine engine with primary and secondary stage airfoils

Abstract

A turbine engine has an engine core defining an engine centerline and including a rotor assembly and a stator assembly. A turbine engine includes a primary airfoil disposed circumferentially about an engine centerline and defining at least a portion of a rotor assembly. The secondary stage airfoil is located upstream of the primary airfoil and circumferentially disposed about the engine centerline.

Description

Technical Field

[0001] The present disclosure relates generally to components for turbine engines and, more particularly, to a primary airfoil and a secondary-stage airfoil located upstream of the primary airfoil. Background Art

[0002] In a typical turbofan application, composite materials are used for flow path components and blades and vanes. Composite materials generally include a fiber reinforced matrix and exhibit a high strength-to-weight ratio. Due to the high strength-to-weight ratio and the plasticity of adopting relatively complex shapes, composite materials are used in various applications, such as turbine engines or aircraft. Composite materials can, for example, be mounted on or define a portion of a fuselage and / or a wing, rudder, manifold, airfoil, or other component of an aircraft or turbine engine. Extreme loads or sudden forces can be applied to composite components of an aircraft or turbine engine. For example, during the ingestion of various materials by a turbine engine, one or more airfoils may experience extreme loads. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] A full and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0004] Figure 1 is a schematic cross-sectional view of a gas turbine engine according to aspects disclosed herein.

[0005] Figure 2 yes Figure 1 An enlarged schematic cross-sectional illustration of a front end of an exemplary gas turbine engine.

[0006] Figure 3 is a schematic illustration of an airfoil in the form of a fan blade according to aspects disclosed herein.

[0007] Figure 4 is along Figure 3 A schematic cross-sectional view taken along line IV-IV of shows a leading edge protector according to aspects disclosed herein.

[0008] Figure 5 is with Figure 2 A similar schematic cross-sectional view is enlarged and shows a primary airfoil and a secondary airfoil located upstream of the primary airfoil.

[0009] Figure 6 yes Figure 1 A first exemplary axial view of a first inlet arrangement of a turbine engine of FIG. 1 is shown illustrating an arrangement of a secondary-stage airfoil according to one aspect disclosed herein.

[0010] Figure 7 yes Figure 1 FIG. 1 is a second exemplary axial view of a second inlet arrangement of a turbine engine of FIG. 1 , showing an arrangement of a secondary-stage airfoil according to another aspect disclosed herein. DETAILED DESCRIPTION

[0011] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are shown in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the present disclosure.

[0012] Aspects disclosed herein relate to a set of consecutive airfoil stages, including a secondary-stage airfoil and a primary airfoil located downstream of the secondary-stage airfoil. For ease of illustration, the present disclosure will be described with respect to a primary airfoil in the form of a fan blade and a secondary-stage airfoil immediately upstream of the primary airfoil as a partial span inlet guide vane (IGV). Although partial span IGVs and fan blades are shown within a high-speed engine, it will be understood that any consecutive set of stages are contemplated. Furthermore, it will be understood that the aspects disclosed herein are not limited thereto and are generally applicable to alternative engines as well as non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

[0013] The term "exemplary" is used herein 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. Furthermore, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0014] The terms “first,” “second,” and “third” may be used interchangeably herein to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0015] The terms "upstream" and "downstream" refer to the relative directions relative to the flow of fluid in a fluid channel. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows.

[0016] The terms "forward" or "front" mean in front of something, and "rearward" or "aft" means behind something. For example, for a turbine engine, forward means closer to the engine inlet, and rear means closer to the engine nozzle or exhaust.

[0017] The term "fluid" may be used herein to refer to a gas or a liquid or multiple phases. The term "fluid communication" means that a fluid is able to establish a connection between designated areas.

[0018] The term "radial" or "radially" refers to directions away from a common center. For example, in the general context of turbine engines, radial refers to directions along a ray extending between a central longitudinal axis of the engine and an outer circumference of the engine.

[0019] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, front, rear, etc.) are used only for identification purposes to help the reader understand the present disclosure and do not create limitations, particularly limitations on the position, orientation, or use of aspects of the disclosure described herein. Connection references (e.g., attachment, coupling, connection, and engagement) should be interpreted broadly and, unless otherwise indicated, may include intermediate structural elements between sets of elements and relative movement between elements. Therefore, a connection reference does not necessarily mean that two elements are directly connected and have a fixed relationship to each other. The exemplary drawings are for illustrative purposes only, and the dimensions, positions, orders, and relative sizes reflected in the drawings may vary.

[0020] Unless the context clearly dictates otherwise, the singular forms "a," "an," and "the" include plural referents. Also, as used herein, the term "set" or a "set of" elements may include any number of elements, including only one.

[0021] The term "composite material" as used herein refers to a component having two or more materials. A composite material can be a combination of at least two or more metals, non-metals, or a combination of metal and non-metal elements or materials. Examples of composite materials can be, but are not limited to, polymer matrix composites (PMC), ceramic matrix composites (CMC), metal matrix composites (MMC), carbon fibers, polymer resins, thermoplastics, bismaleimide (BMI), polyimide materials, epoxy resins, glass fibers, and silicon matrix materials.

[0022] As used herein, a "composite" component refers to a structure or component that includes any suitable composite material. A composite component (e.g., a composite airfoil) may include multiple layers or plies of composite material. The stiffness, material, and dimensions of these layers or plies may be varied to achieve a desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength.

[0023] One or more layers of adhesive may be used to form or join the composite parts. The adhesive may include resins and phenolics, where the adhesive may need to be cured at elevated temperatures or employ other hardening techniques.

[0024] As used herein, PMC refers to a class of materials. For example, PMC materials are defined in part as prepregs, which are reinforcement materials pre-impregnated with a polymer matrix material (e.g., a thermoplastic resin). Non-limiting examples of processes for producing thermoplastic prepregs include hot melt prepregs, in which a fiber reinforcement is pulled through a molten bath of resin; and powder prepregs, in which the resin is electrostatically deposited onto the fiber reinforcement by way of non-limiting example, and then adhered to the fibers in an oven or by means of heated rollers by way of non-limiting example. The prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked together to form the number of stacked plies required for the part.

[0025] The multilayer prepreg is stacked to a thickness and orientation suitable for the composite part, and then the resin is cured and solidified to form a fiber reinforced composite part. The resin of the matrix material of PMC can generally be divided into a thermosetting resin or a thermoplastic resin. Thermoplastic resins are generally classified as polymers that can be repeatedly softened and flowed when heated and hardened when fully cooled due to physical changes rather than chemical changes. The significant example categories of thermoplastic resins include nylon, thermoplastic polyesters, polyaryletherketones and polycarbonate resins. Specific examples of high-performance thermoplastic resins considered for aerospace applications include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK) and polyphenylene sulfide (PPS). On the contrary, once fully cured into a hard rigid solid, thermosetting resins do not significantly soften when heated, but thermally decompose when fully heated. The significant examples of thermosetting resins include epoxy resins, bismaleimide (BMI) and polyimide resins.

[0026] In another non-limiting example, in addition to using prepregs, in the case of using thermoplastic polymers, woven fabrics can also be used. Woven fabrics can include but are not limited to dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg woven architectures can be made in a similar manner. In this way, the fiber volume of a part can be adjusted by specifying the relative concentrations of thermoplastic fibers and reinforcing fibers that have been woven or woven together. In addition, different types of reinforcing fibers can be woven or woven together to adjust the properties of the part at various concentrations. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at various concentrations to adjust the properties of the component. Carbon fibers provide the strength of the system, glass fibers can be added to enhance impact properties, which is a design feature of parts located near the engine inlet, and thermoplastic fibers provide bonding for the reinforcing fibers.

[0027] In yet another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of a composite component. Typically, RTM includes applying dry fiber or matrix material to a mold or cavity. The dry fiber or matrix material can include prepreg, braided material, woven material, or any combination thereof.

[0028] The resin may be pumped or otherwise provided to the mold or cavity to impregnate the dry fibers or matrix material. The impregnated fiber or matrix material and resin combination is then cured and removed from the mold. When removed from the mold, the composite part may require a post-curing process.

[0029] It is contemplated that RTM may be a vacuum-assisted process. That is, the air in the cavity or mold may be removed and replaced with resin prior to heating or curing. It is further contemplated that placement of dry fiber or matrix material may be manual or automated.

[0030] The dry fiber or matrix material may be shaped to form a composite part or to introduce resin. Optionally, additional layers or reinforcements of a material different from the dry fiber or matrix material may also be included or added prior to heating or curing.

[0031] As used herein, CMC refers to a class of materials having reinforcing fibers in a ceramic matrix. Typically, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates (e.g., mullite), or mixtures thereof), or mixtures thereof.

[0032] Some examples of ceramic matrix materials may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (A12O3), silicon dioxide (SiO2), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, the ceramic matrix may also include ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite).

[0033] In general, a particular CMC may be referred to as a combination of its fiber type / matrix type. For example, C / SiC represents carbon fiber reinforced silicon carbide, SiC / SiC represents silicon carbide fiber reinforced silicon carbide, SiC / SiN represents silicon carbide fiber reinforced silicon nitride, SiC / SiC-SiN represents silicon carbide fiber reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, a CMC may be composed of a matrix and reinforcing fibers, the reinforcing fibers including oxide-based materials such as aluminum oxide (A12O3), silicon dioxide (SiO2), aluminosilicates, and mixtures thereof. Aluminosilicates may include crystalline materials such as mullite (3A12O3·2SiO2) as well as glassy aluminosilicates.

[0034] In certain non-limiting examples, the reinforcing fibers may be bundled and / or coated before being incorporated into the matrix. For example, the fiber bundles may be formed into reinforcing tapes, such as unidirectional reinforcing tapes. Multiple tapes may be laid together to form a preform component. The fiber bundles may be impregnated with a slurry composition before or after forming the preform. The preform may then be heat treated and subsequently chemically treated to obtain a component formed of a CMC material having a desired chemical composition. For example, the preform may be cured or burned out to produce a high carbon residue in the preform, followed by melt infiltration with silicon, or cured or pyrolyzed to produce a silicon carbide matrix in the preform, followed by chemical vapor infiltration with silicon carbide. Before or after chemical vapor infiltration, additional steps may be taken to improve the densification of the preform by injecting a liquid resin or polymer into the preform and then performing a heat treatment step to fill the voids with silicon carbide. CMC materials as used herein may be formed using any known or later developed method including, but not limited to, melt infiltration, chemical vapor infiltration, polymer infiltration pyrolysis (PIP), or any combination thereof.

[0035] Such materials, as well as certain monolithic ceramics (i.e., ceramic materials that do not contain reinforcement materials), are particularly suitable for higher temperature applications. In addition, these ceramic materials are lightweight compared to superalloys, but can still provide strength and durability to the parts made from them. As a result, such materials are currently being considered for use in many gas turbine components used in the high temperature sections of gas turbine engines, such as airfoils (e.g., turbines and buckets), combustors, shrouds, and other similar components that would benefit from the lighter weight and higher temperature capabilities that these materials can provide.

[0036] The term "metal" as used herein refers to materials including metals, such as but not limited to titanium, iron, aluminum, stainless steel, and nickel alloys. A metal material or alloy may be a combination of at least two or more elements or materials, at least one of which is a metal.

[0037] As used herein, the term “leading length” or “LL” refers to the length between the leading edge of an airfoil and a seam between a leading edge protector and a portion of the airfoil.

[0038] The term “primary lead-off length” or “PLL” as used herein refers to the lead-off length of the primary airfoil.

[0039] As used herein, the term “secondary leading length” or “SLL” refers to the leading length of the secondary stage airfoil immediately upstream of the primary airfoil.

[0040] As used herein, the term “chord length” or “CL” refers to the length between the leading edge of an airfoil and the trailing edge of the airfoil.

[0041] As used herein, the term “primary chord length” or “PCL” refers to the chord length of the primary airfoil.

[0042] As used herein, the term “secondary chord length” or “SCL” refers to the chord length of the secondary stage airfoil.

[0043] As used herein, the term "airfoil protection factor" or "APF" refers to a relationship in the form of a ratio of the leading length to the chord length of an airfoil. As more protection is provided to any given airfoil, the leading length increases, and thus the APF also increases.

[0044] As used herein, the term “stage protection factor” or “SPF” refers to the relationship in the form of a ratio of the airfoil protection factor of a primary airfoil or “APF1 ” to the airfoil protection factor of a secondary-stage airfoil or “APF2 ”.

[0045] As used herein, the term “fan pressure ratio” refers to the ratio of the pressure immediately downstream of a plurality of fan blades to the pressure immediately upstream of the plurality of fan blades.

[0046] The term "fan tip speed" as used herein is defined by a plurality of fan blades and refers to a linear speed of an outer tip of each of the plurality of fan blades in a circumferential direction.

[0047] As used herein, the term "rated speed" refers to the maximum operating speed of a turbine engine wherein the plurality of fan blades are producing a maximum amount of power.

[0048] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language indicates otherwise, these ranges are identified and include all sub-ranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0049] In certain exemplary embodiments of the present disclosure, a turbine engine defining a centerline and a circumferential direction is provided. A gas turbine engine generally includes a rotor assembly and a stator assembly. The rotor assembly and the stator assembly together define a substantially annular flow path through the gas turbine engine. It is contemplated that the fan blades described herein include composite fan blades located within a high-speed fan inlet.

[0050] The rotor assembly includes a set of blades. Each blade in the set of blades is mounted to the rotor, such as to a disk, by a dovetail of the blade. The set of blades is distributed circumferentially around the centerline of the engine. It is further envisioned that the set of blades is mounted to a disk.

[0051] The stator assembly includes a set of vanes. The set of vanes extends between the inner band and the outer band and is distributed circumferentially about the centerline. The set of vanes also defines a set of nozzles. It is further contemplated that the set of vanes may include a pair of vanes that define a single nozzle. The number of vanes and the number of nozzles in the stage are both selected to control flow across each vane and through the nozzle.

[0052] Each of the set of blades and the set of vanes includes at least one airfoil having a protective cover on the leading edge of the airfoil. The protective cover may be any suitable material, including a composite cover, a metal cover, or a non-metal cover. The protective cover is referred to herein as a leading edge protector.

[0053] The amount of protective covering on the leading edge of an airfoil can be designed for various flight conditions, including takeoff, descent, and idle. An important consideration when designing airfoils, particularly partial span IGVs and fan blades, is balancing the added weight of the protective covering or shroud on the leading edge with the amount of adequate protection required for the leading edge of a particular airfoil. This balance between added weight and leading edge protection is particularly important in fan modules of high speed gas turbine engines. Effective designs achieve the correct balance between leading edge dominance in response to bird ingestion or similar events and airfoil dominance in blade aerodynamic properties, such as flutter.

[0054] There is a tradeoff between the percentage of the airfoil chord length that is covered by the leading edge protector and the performance of the airfoil. The protective cover provides stiffness to the airfoil to prevent bird ingestion, while the remainder of the blade is more flexible for aerodynamic purposes. Additionally, the factors that influence the optimal design decision are not necessarily the same for a partial span IGV. A partial span IGV is stationary. The dynamics of a bird ingestion event on a partial span IGV are different than those of a rotating airfoil. Additionally, the large number of partial span IGVs compared to the number of fan blades must also be taken into account.

[0055] The inventors determined that the leading edge protector should overlap the airfoil sufficiently to provide a strong enough bond, but preferably minimize the overlap to allow the blade to flex. The leading edge protector also provides erosion protection and is required for both static and rotating airfoils. The leading edge protector characteristics were derived from multiple tests and simulation analyses covering the ingestion of birds of different sizes at different span locations, as well as analysis of blades returned for repair after bird strikes in operational service. In addition, partial span IGVs do not rotate and are subject to different stresses when impacted by birds. Multiple simulations and analyses were considered depending on the manner in which the bird strikes the fan. Direct strikes centered on the leading edge protector or strikes between two adjacent blades, where the blades are more likely to pass through without breaking into smaller pieces, were simulated and analyzed. The inventors needed to perform various empirical tests to determine what percentage of the airfoil needed to be covered with the leading edge protector for a particular engine architecture. The differences in the forces acting on the airfoil when struck by a bird or similar object make determining the airfoil coverage percentage particularly challenging.

[0056] The inventor's practice includes designing an airfoil stage, modifying the airfoil stage by adding a leading edge protector, and redesigning the airfoil stage, wherein the leading edge protector meets the protection requirements associated with the airfoil stage. After calculating and checking the amount of protection provided and the amount of weight increase or decrease associated with the leading edge protector, the process is repeated until a specific architecture and performance requirements are met, and then the process is repeated for different architectures and performance requirements. Examples of these architectures and performance requirements are described below. Figure 1 An example of an architecture is shown in It has been found that an airfoil design that meets the performance requirements of one location in an engine or class or type of engine may not necessarily meet the requirements of another location.

[0057] Specific implementation methods of the main embodiments

[0058] Referring now to the drawings, wherein like numerals represent like elements throughout the several views, Figure 1 1 is a schematic cross-sectional view of a turbine engine 10 for an aircraft. More specifically, the turbine engine 10 is a high bypass gas turbofan jet engine. Figure 1 As shown, a turbine engine 10 has an engine core 11 that defines a generally longitudinally extending axis or engine centerline 12 defining an axial direction (designated “A”), a radial direction (designated “R”), and a circumferential direction (i.e., a direction extending about the axial direction A; see e.g. Figure 6 ). Generally, the turbine engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14 .

[0059] The illustrated turbine 16 generally includes a substantially tubular outer casing 18 defining an annular inlet 20. The outer casing 18 surrounds a compressor section, including a supercharger or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24, in series flow relationship; a combustion section 26; a turbine section, including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and an ejection exhaust nozzle section 32. A high pressure (HP) shaft or shafts 34 drive-connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or shafts 36 drive-connects the LP turbine 30 to the LP compressor 22. The LP turbine 30 may also be referred to as a "drive turbine". The turbine engine 10 as described herein is provided as a non-limiting example only, and other architectures are possible, such as, but not limited to, a steam turbine engine, a supercritical carbon dioxide turbine engine, or any other suitable turbine engine.

[0060] The fan section 14 includes a variable pitch fan 38, referred to herein as "fan 38", having a plurality of fan blades 40 coupled to a disk 42 in a spaced manner. More specifically, for the illustrated embodiment, the fan section 14 includes a single-stage fan 38 that accommodates a single-stage fan blade 40. As shown, the plurality of fan blades 40 extend outwardly from the disk 42 generally in a radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by means of the plurality of fan blades 40 being operably coupled to an appropriate actuating member 44, the actuating member being configured to uniformly and collectively change the pitch of the plurality of fan blades 40. The fan 38 is mechanically coupled to the LP turbine 30 and can rotate with the LP turbine 30, or drive the turbine. More specifically, the plurality of fan blades 40, the disk 42, and the actuating member 44 can rotate together about the engine centerline 12 through the LP shaft or spool 36 in a "direct drive" configuration. Therefore, the fan 38 is coupled to the LP turbine 30 such that the fan 38 may be rotated by the LP turbine 30 at the same rotational speed as the LP turbine 30 .

[0061] Furthermore, it should be appreciated that fan 38 defines a fan pressure ratio, while plurality of fan blades 40 each define a fan tip speed. As will be described in greater detail below, the illustrated exemplary turbofan engine 10 defines a relatively high fan tip speed and a relatively low fan pressure ratio during turbofan engine operation at rated speed.

[0062] The disk 42 is covered by a rotatable front hub 48 having an aerodynamic profile to facilitate airflow 58 through the plurality of fan blades 40. In addition, the exemplary fan section 14 includes an annular fan casing or nacelle 50 that circumferentially surrounds the plurality of fan blades 40 of the fan 38 and / or at least a portion of the turbine 16. More specifically, the nacelle 50 includes an inner wall 52, and a downstream section 54 of the inner wall 52 of the nacelle 50 extends over an outer portion of the turbine 16 to define a bypass airflow passage 56 therebetween. In addition, the nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 55.

[0063] The rotor assembly 92 includes, but is not limited to, the fan blades 40, the HP turbine rotor blades 70, and the LP turbine rotor blades 74. The disk 42, the HP shaft 34, and the LP shaft 36 are also considered part of the rotor assembly 92. The stator assembly 94 includes, but is not limited to, the outlet guide vanes 55 of the blade set, the HP turbine stator vanes 68, and the LP turbine stator vanes 72.

[0064] During operation of turbine engine 10, a volume of airflow 58 enters turbine engine 10 through nacelle 50 and associated first inlet 60a of fan section 14. As volume of airflow 58 passes through plurality of fan blades 40, a first portion of airflow 58, as indicated by arrow 62, is directed or directed into bypass airflow passage 56, while a second portion of airflow 58, as indicated by arrow 64, is directed or directed into LP compressor 22. The ratio between first portion 62 of air and second portion 64 of air is generally referred to as a bypass ratio. The bypass ratio may range from about 7:1 to about 20:1, such as from about 10:1 to about 18:1. The pressure of second portion 64 of air increases as it is directed through HP compressor 24 and into combustion section 26, where second portion 64 of air is mixed with fuel and combusted to provide combustion gases 66.

[0065] The combustion gases 66 are directed through the HP turbine 28, wherein a portion of the heat and / or kinetic energy from the combustion gases 66 is extracted via successive stages of HP turbine stator blades 68 coupled to the outer casing 18 and HP turbine rotor blades 70 coupled to the HP shaft or spool 34, thereby rotating the HP shaft or spool 34 to support operation of the HP compressor 24. The combustion gases 66 are then directed through the LP turbine 30, wherein a second portion of the heat and kinetic energy is extracted from the combustion gases 66 via successive stages of LP turbine stator blades 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft or spool 36, thereby rotating the LP shaft or spool 36 to support operation of the LP compressor 22 and / or rotation of the fan 38.

[0066] The combustion gases 66 are then directed through the jet exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of the air 62 is significantly increased as it is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the turbine engine 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbine 16.

[0067] It should be understood that Figure 1 The turbine engine 10 shown and described above is only an example, and the turbine engine 10 may have any other suitable configuration. For example, in other exemplary embodiments, the turbine 16 may include any other suitable number of compressors, turbines, and / or shafts or spools. In addition, the turbine engine 10 may not include each of the features described herein, or alternatively, may include one or more features not described herein. For example, in other exemplary embodiments, the fan 38 may not be a variable pitch fan.

[0068] To increase the efficiency of turbine 16, LP turbine 30 rotates at a relatively high rotational speed, thereby causing the plurality of fan blades 40 of fan 38 to rotate at a relatively high rotational speed. For example, during operation of turbine engine 10 at rated speed, the fan tip speed of each of the plurality of fan blades 40 may be greater than 1,250 feet per second. In another example, in certain exemplary embodiments, during operation of turbine engine 10 at rated speed, the fan tip speed of each of the plurality of fan blades 40 may be greater than about 1,350 feet per second, such as greater than about 1,450 feet per second, such as greater than about 1,550 feet per second, such as up to about 2,200 feet per second.

[0069] Despite these relatively high fan tip speeds, fan 38 is configured to produce a relatively low fan pressure ratio. For example, during operation of turbine engine 10 at rated speed, fan 38 has a fan pressure ratio of less than 1.5. For example, during operation of turbine engine 10 at rated speed, the fan pressure ratio may be between about 1.15 and about 1.5, such as between about 1.25 and about 1.4.

[0070] As will be appreciated, operating the direct drive turbine engine 10 in this manner generally results in a loss of efficiency of the fan 38 due to impact losses and flow separation of the airflow 58 on the plurality of fan blades 40, particularly at the radially outer tips of the plurality of fan blades 40 of the fan 38. Therefore, a set of pre-swirl features 80 are located upstream of the plurality of fan blades 40 of the fan 38 to offset or minimize such efficiency losses of the fan 38. By including a set of pre-swirl features 80, the efficiency gains of the turbine 16 (due to, for example, the increased rotational speed of the LP turbine 30) outweigh the potential efficiency losses discussed above.

[0071] The set of pre-swirl features 80 may be a plurality of partial span IGVs 100 arranged circumferentially about the engine centerline 12. The plurality of fan blades 40 may each include a first leading edge protector 90a, and the plurality of partial span IGVs 100 may each include a second leading edge protector 90b. Although two partial span IGVs 100 are shown, it should be understood that any number of partial span IGVs 100 are contemplated.

[0072] Figure 2 yes Figure 1 1 is an enlarged cross-sectional view of the front end of the fan section 14 and the turbine 16 of the turbine engine 10. Each of the plurality of partial span IGVs 100 is attached to or integrated with the nacelle 50. The plurality of partial span IGVs 100 are cantilevered from the nacelle 50 (e.g., from the inner wall 52 of the nacelle 50) at an upstream position of the plurality of fan blades 40 of the fan 38 in the axial direction A and behind the first inlet 60a of the nacelle 50.

[0073] Each of the plurality of partial span IGVs 100 extends between an IGV root 102 and an IGV tip 104 in a radial direction R. In addition, each of the plurality of partial span IGVs 100 extends between an IGV leading edge 108 and an IGV trailing edge 110 in an axial direction A. Each of the plurality of partial span IGVs 100 is attached / connected to the nacelle 50 at the IGV root 102 by a suitable connection means (not shown). For example, each of the plurality of partial span IGVs 100 may be bolted to the inner wall 52 of the nacelle 50 at the IGV root 102, welded to the inner wall 52 of the nacelle 50 at the IGV root 102, or attached to the nacelle 50 at the IGV root 102 in any other suitable manner.

[0074] The IGV tip 104 defines a cantilever end of each of the plurality of partial span IGVs 100, which results in a cantilever IGV. Thus, the IGV tip 104 is not connected to adjacent partial span IGVs 100, i.e., adjacent partial span IGVs 100 do not contact each other at the IGV tip 104, and does not include any intermediate connection members, such as, but not limited to, connection rings or struts, at the IGV tip 104. More specifically, each partial span IGV 100 is completely supported by a connection to the nacelle 50 at the corresponding IGV root 102.

[0075] The tip cap 103 can cover at least a portion of the IGV tip 104. In some embodiments, the tip cap 103 can extend the entire axial length of the IGV tip 104, as shown.

[0076] Furthermore, each of the plurality of partial span IGVs 100 only partially extends into the first inlet 60a of the turbine engine 10. More specifically, each of the plurality of partial span IGVs 100 defines an IGV span 106 along the radial direction R. The IGV span 106 refers to a measurement along the radial direction R between an IGV root 102 and an IGV tip 104 of the partial span IGV 100 at an IGV leading edge 108 of the partial span IGV 100.

[0077] Similarly, the plurality of fan blades 40 of the fan 38 extend along the radial direction R between a fan tip 114 and a fan root 116 to define a fan blade span 112 along the radial direction R. More specifically, each of the plurality of fan blades 40 of the fan 38 further defines a fan leading edge 115 and a fan trailing edge 117 , and the fan blade span 112 refers to a measurement along the radial direction R between the fan tip 114 and the fan root 116 of the fan blade 40 at the fan leading edge 115 of the respective fan blade 40 .

[0078] In one example, IGV span 106 is at least about 5 percent of fan blade span 112 and at most about 55 percent of fan blade span 112. For example, in certain exemplary embodiments, IGV span 106 may be between 15 percent of fan blade span 112 and about 45 percent of fan blade span 112, such as between about 30 percent of fan blade span 112 and about 40 percent of fan blade span 112.

[0079] Each of the plurality of partial span IGVs 100 is configured to pre-swirl the airflow 58 provided through the first inlet 60a of the nacelle 50, upstream of the plurality of fan blades 40 of the fan 38. Pre-swirl the airflow 58 provided through the first inlet 60a of the nacelle 50 before it reaches the plurality of fan blades 40 of the fan 38 may reduce separation losses and / or impulse losses, thereby allowing the fan 38 to operate at the relatively high fan tip speeds described above with less loss in efficiency.

[0080] Figure 1 and Figure 2 The plurality of partial span IGVs 100 shown in FIG. 1 are generally configured to pre-swirl a portion of the airflow 58 passing through the first inlet 60a of the outer nacelle 50 in the same rotational direction as the rotational direction of the plurality of fan blades 40 of the fan 38. For example, the plurality of fan blades 40 of the fan 38 are configured to rotate clockwise when viewed from front to rear, and the plurality of partial span IGVs 100 are configured to pre-swirl a portion of the airflow 58 passing through the first inlet 60a of the outer nacelle 50 in the same direction. However, in other exemplary embodiments, the gas turbine engine 10 may include a fan 38 in which the fan blades 40 are configured to rotate counterclockwise when viewed from front to rear, in which case the plurality of partial span IGVs 100 may instead be mirrored such that they are configured to pre-swirl the airflow in the counterclockwise direction. In addition, it is also contemplated that the plurality of partial span IGVs 100 are configured to pre-swirl the airflow 58 in a rotational direction opposite to the plurality of fan blades 40 of the fan 38.

[0081] Furthermore, it should be appreciated that the illustrated exemplary partial span IGV 100 is provided by way of example only. The plurality of partial span IGVs 100 may have any suitable configuration for providing a desired amount of pre-swirl upstream of a plurality of fan blades of a fan of a gas turbine engine.

[0082] Figure 3 1 is a schematic illustration of an airfoil 130 in the form of a fan blade 131. As a non-limiting example, fan blade 131 may be a single blade from a plurality of fan blades 40. Although illustrated as fan blade 131, similar properties of airfoil 130 may be attributed to a single partial-span IGV from a plurality of partial-span IGVs 100 described herein. It is contemplated that airfoil 130 may be a blade, bucket, airfoil, or other component of any turbine engine, such as, but not limited to, a gas turbine engine, a turboprop engine, a turboshaft engine, or a turbofan engine.

[0083] The airfoil 130 may include a wall 132 defining an interior 133. The wall 132 may define an exterior surface 134 that extends radially between a leading edge 135 and a trailing edge 136 to define a chord-wise direction (denoted as “CD”). The chord-wise direction CD may generally extend along the axial direction A, however, as the curvature of the airfoil 130 increases, it should be understood that the chord-wise direction CD may extend into and out of the page and across the page from left to right. The airfoil 130 has a chord length (denoted as “CL”) measured along the chord-wise direction CD between the leading edge 135 and the trailing edge 136.

[0084] The outer surface 134 may further extend between a root 137 and a tip 138 to define a span direction (denoted as "SD") parallel to the radial direction R. The airfoil 130 has a span length (denoted as "SL") measured between the root 137 and the tip 138 along the span direction SD, wherein the root 137 is considered to be 0% of the span length SL and the tip 138 is considered to be 100% of the span length SL. The span length SL is a measurement along the span direction SD between the tip 138 and the root 137 of the airfoil 130 at the leading edge 135. It should be understood that the airfoil 130 may take any suitable shape, profile or form, including that the leading edge 135 does not need to be curved. The leading edge 135 may be the fan leading edge 115, and the trailing edge 136 may be the fan trailing edge 117. The root 137 may be the fan root 116, and the tip 138 may be the fan tip 114. It is also conceivable that the leading edge 135 is the IGV leading edge 108, the trailing edge is the IGV trailing edge 110, and the root 137 is the IGV root 102 ( Figure 2 ), and the tip 138 is the IGV tip 104 ( Figure 2 ).

[0085] The outer surface 134 is defined by the leading edge protector 90 and the main body portion 150. A seam 139 separates the leading edge protector 90 from the main body portion 150 along the outer surface 134. The leading edge protector 90 extends in a chordwise direction CD between the leading edge 135 and the seam 139 to define a leading length (denoted as 'LL').

[0086] The leading edge protector 90 is made of a more durable material than the main body portion 150. Typically, the leading edge protector 90 is made of, but not limited to, steel, aluminum, a refractory metal such as titanium, or a superalloy based on nickel, cobalt, or iron. It should be understood that the leading edge protector 90 can be the first leading edge protector 90a of the fan blade 131 in the plurality of fan blades 40. In addition, the leading edge protector 90 can be the second leading edge protector 90b of a single partial span IGV in the plurality of partial span IGVs 100. Either or both of the first leading edge protector 90a and the second leading edge protector 90b can be a metal leading edge protector. Either or both of the first leading edge protector 90a and the second leading edge protector 90b can be made of a polyurethane material. In addition, the leading edge protectors 90, 90a, 90b described herein can be any suitable material, such as a metal, a thermoplastic, or a polyurethane, wherein the two are the same or different.

[0087] The main body portion 150 may include a body leading edge 152 spaced a distance (indicated as "D") from the leading edge 135. The body leading edge 152 may define at least a portion or all of the seam 139. It is further contemplated that at least a portion of the leading edge protector 90 overlaps the main body portion 150 such that at least a portion (indicated in phantom) or all of the body leading edge 152 is located upstream of the seam 139. In other words, the leading edge protector 90 may define a sheath 144 or protective covering over the body leading edge 152.

[0088] The body portion 150 may be a composite component made of one or more layers of material. The one or more layers of material may be applied at the same level or at different levels of manufacturing the airfoil 130. As a non-limiting example, the body portion 150 may include at least one polymer matrix composite (PMC) portion or polymer portion. The polymer matrix composite may include, but is not limited to, a matrix of thermosetting material (epoxy, phenolic) or thermoplastic material (polycarbonate, polyvinyl chloride, nylon, acrylic) and embedded glass, carbon, steel or Kevlar fibers.

[0089] The leading edge protector 90 and the main body portion 150 can be formed by a variety of methods, including additive manufacturing, casting, electroforming, or direct metal laser melting, in non-limiting examples. As used herein, an "additively manufactured" part refers to a part formed by an additive manufacturing (AM) process, in which the part is built layer by layer by continuously depositing material. AM is a suitable name for describing the technology of building 3D objects by adding materials layer by layer (whether the materials are plastic, ceramic, or metal). AM technology can utilize computers, 3D modeling software (computer-aided design or CAD), machine equipment, and layered materials. Once a CAD sketch is generated, the AM equipment can read the data from the CAD file and lay or add continuous layers of liquid, powder, sheet material, or other materials in a layer-by-layer manner to manufacture a 3D object. It should be understood that the term "additive manufacturing" covers many technologies, including subsets such as 3D printing, rapid prototyping (RP), direct digital manufacturing (DDM), layered manufacturing, and additive manufacturing. Non-limiting examples of additive manufacturing that can be used to form additively manufactured parts include powder bed fusion, vat photopolymerization, binder jetting, material extrusion, directed energy deposition, material jetting, or sheet lamination. It is also contemplated that the process used can include printing a negative of the part through a refractory metal, ceramic, or printed plastic, and then casting the part using the negative.

[0090] The relationship between the leading length LL and the chord length CL is referred to herein as the airfoil protection factor or simply “APF.” In other words, for any given airfoil 130 having a predetermined chord length CL, as the amount of coverage provided by the leading edge protector 90 increases, the leading length LL, and thus the APF, also increases.

[0091] Figure 4 is along Figure 3 The leading edge protector 90 is a schematic cross-section taken along line IV-IV. The leading edge protector 90 is a sheath 144 having a first wall 146, a second wall 147, and a third wall 148, the third wall 148 interconnecting the first wall 146 and the second wall 147. The first wall 146, the second wall 147, and the third wall 148 of the leading edge protector 90 are oriented and shaped so that they define a generally U-shaped (or C-shaped) channel 154 therebetween. Figure 4 As shown, and as will be discussed below, the channel 154 is sized and shaped to receive at least a portion of the body leading edge 152 of the main body portion 150. It is important to note that the shape of the channel 154 is shown by way of example only, and the channel 154 is not limited to this particular shape and is not necessarily drawn to scale.

[0092] The airfoil 130 may extend between the first side 156 and the second side 158. The seam 139 ( Figure 3) may include two seams 139c, 139d at corresponding ends of the channel 154. The leading length LL is measured from the leading edge 135 to the seam 139d farthest from the leading edge 135. Although shown in two different locations, it should be understood that the seams 139c, 139d can be located at the same leading length LL. Although shown as having a rectangular blunt end at the seam 139, the leading edge protector 90 can be tapered so that the leading edge protector 90 and the main body portion 150 are flush at the seam 90 to define the outer surface 134.

[0093] It should be understood that Figure 4 The leading edge protector 90 may be the first leading edge protector 90a, the second leading edge protector 90b, or both the first leading edge protector 90a and the second leading edge protector 90b.

[0094] Figure 5 1 is a schematic enlarged view of a portion of the fan section 14. A plurality of fan blades 40 define a primary airfoil 200a within the fan section 14. The first airfoil 230a in the primary airfoil 200a is similar to the airfoil 130 ( Figure 3 ), and thus like parts of first airfoil 230a will be identified with like numerals increased by 100 and with the symbol "a", it being understood that the description of like parts of airfoil 130 applies to first airfoil 230a unless otherwise noted. Although only a single fan blade is shown in cross-section, it is understood that a plurality of fan blades 40 are included and spaced apart about fan section 14.

[0095] The first airfoil 230a has a first span length (denoted as "SL1") measured along the span direction SD between the first root 237a and the first tip 238a, wherein the first root 237a is considered to be 0% of the first span length SL1 and the first tip 238a is considered to be 100% of the first span length SL1. The first span length SL1 is a value measured along the span direction SD between the first root 237a and the first tip 238a of the first airfoil 230a at the first leading edge 235a.

[0096] The first leading edge protector 240a extends in the chord-wise direction CD between the first leading edge 235a and the first seam 239a to define a primary leading length (denoted as "PLL"). The first airfoil 230a has a primary chord length (denoted as "PCL") measured in the chord-wise direction CD between the first leading edge 235a and the first trailing edge 236a.

[0097] The relationship between the primary preamble length PLL and the primary chord length PCL is expressed here by the first expression of the APF:

[0098]

[0099] A plurality of partial span IGVs 100 define a secondary-stage airfoil 200b located upstream of the primary airfoil 200a. A second airfoil 230b in the secondary-stage airfoil 200b is similar to the airfoil 130 described previously, and thus similar portions of the second airfoil 230b will be identified with similar numbers increased by 100 and with the symbol "b", with the understanding that the description of similar portions of the airfoil 130 applies to the second airfoil 230b unless otherwise noted. The second airfoil 230b is located upstream of the first airfoil 230a. Although only a single partial span IGV 100 is shown in cross section, it is understood that a plurality of partial span IGVs 100 are included and spaced apart around the fan section 14.

[0100] The second leading edge protector 240b extends in the chord-wise direction CD between the second leading edge 235b and the second seam 239b to define a secondary leading length (denoted as "SLL"). The second airfoil 230b has a secondary chord length (denoted as "SCL") measured in the chord-wise direction CD between the second leading edge 235b and the second trailing edge 236b.

[0101] The second airfoil 230b has a second span length (denoted as "SL2") measured along the span direction SD between the second root 237b and the second tip 238b, wherein the second root 237b is considered to be 0% of the second span length SL2, and the second tip 238b is considered to be 100% of the second span length SL2. The second span length SL2 is the maximum distance between the second root 237b and the second tip 238b of the second airfoil 230b.

[0102] The first leading edge protector 240a and the second leading edge protector 240b may each define a first sheath 244a and a second sheath 244b. The outer surface 234a, 234b of each airfoil 230a, 230b is defined by the corresponding first leading edge protector 240a and the second leading edge protector 240b and the corresponding first body portion 250a and the second body portion 250b. The first body portion 250a and the second body portion 250b may each include a corresponding first body leading edge 252a and a second body leading edge 252b, which may define at least a portion or all of the corresponding seams 239a, 239b. The first body portion 250a and the second body portion 250b may both be composite components. It is also conceivable that the first body portion 250a of the primary airfoil 200a is a composite component, while the second body portion 250b of the secondary airfoil 200b is metal, thereby defining a secondary metal airfoil.

[0103] The relationship between the secondary leader length SLL and the secondary chord length SCL is expressed here by the second expression of the APF:

[0104]

[0105] As will be discussed further herein, the APF describes the amount of protection coverage of the leading edge protector of any of the airfoils 130, 230a, 230b described herein. The balanced trade-off between the amount of protection and the weight gain / loss associated with any leading edge protector described herein can be represented by an APF value from 0.1 to 0.3, inclusive. In other words, to meet the protection requirements, the leading edge protector described herein should protect at least 10% and up to 30% (including 30%) of the width of the airfoil measured from the leading edge of the airfoil described herein before becoming too heavy.

[0106] The primary airfoil 200a has a first number of airfoils and the secondary stage airfoil 200b has a second number of airfoils, which may be different from the first number. In other words, the size and number of airfoils of successive stages of airfoils may vary.

[0107] Figure 6 yes Figure 1 and Figure 2 1 , a first exemplary axial view of a first inlet 60a of a turbine engine 10 of the embodiment of the present invention is shown, showing a plurality of partial span IGVs 100 of the turbine engine 10, and as a non-limiting example, a secondary stage airfoil 200b is shown. A relatively large number of partial span IGVs 100 are shown, and more specifically, thirty-two partial span IGVs 100 are shown. Any number of the plurality of partial span IGVs 100 is contemplated, including from twenty partial span IGVs 100 to approximately fifty partial span IGVs 100, from thirty partial span IGVs 100 to approximately forty-five partial span IGVs 100. In addition, each of the plurality of partial span IGVs 100 is substantially evenly spaced along the circumferential direction CD. More specifically, each of the plurality of partial span IGVs 100 defines a circumferential spacing 118 with an adjacent partial span IGV 100, wherein the circumferential spacing 118 between each adjacent partial span IGV 100 is substantially equal. Although depicted as substantially equal circumferential spacing, any spacing is contemplated.

[0108] For example, see Figure 7 , showing Figure 1 and Figure 21. A second exemplary axial view of a second inlet 60b arrangement of a turbine engine 10 of FIG. 1. Likewise, the plurality of partial span IGVs 100 of the turbine engine 10 are secondary-stage airfoils 200b of a non-limiting example. In the second exemplary inlet 60b, the turbofan engine 10 may include less than twenty partial span IGVs 100. More specifically, the turbofan engine 10 includes at least eight partial span IGVs 100, or more specifically, includes exactly eight partial span IGVs 100. In addition, the plurality of partial span IGVs 100 are not substantially evenly spaced along the circumferential direction CD. For example, some of the plurality of partial span IGVs 100 define a first circumferential spacing 118A, while other partial span IGVs 100 define a second circumferential spacing 118B.

[0109] The number of partial span IGVs 100 may be greater than the number of fan blades 40 of the fan 38 of the turbine engine 10. In one non-limiting example, the number of partial span IGVs 100 is two to three times the number of fan blades 40. It is also contemplated that the number of partial span IGVs 100 is substantially equal to the number of fan blades 40 or less than the number of fan blades 40 of the fan 38 of the turbine engine 10. Furthermore, it should be understood that in other exemplary embodiments, the turbine engine 10 may include any other suitable number of partial span IGVs 100 and / or circumferential spacing 118 of the partial span IGVs 100.

[0110] It should be understood that the number, size, and configuration of airfoils described herein are provided by way of example only, and in other exemplary embodiments, the airfoils may have any other suitable configuration, including that multiple airfoils may be at multiple rotor stages, etc.

[0111] Finding a viable solution that balances the amount of protective cover for the airfoil described herein while maintaining weight requirements and maintaining the positive effects of the pre-swirl features 80 described herein is a labor intensive and time consuming process because the process is iterative and involves selecting multiple airfoils for the partial span IGV 100 and fan blades with various protector leading edge protector lengths and chord lengths. The design process typically requires that the airfoil 130 ( Figure 3) into a turbine engine designed for a first flight operating condition and demonstrates protection effects and acceptable weight gain / loss for that first flight operating condition. It is time consuming to evaluate whether the same selected airfoil 130 maintains thermal efficiency and acceptable protection effects for other operating conditions under a second, third, or other flight operating condition. In some cases, this may even result in redesigning the airfoil and turbine engine if the conditions are not met. Adding to the process of implementing the pre-swirl feature 80 as a partial span IGV 100 adds considerable variables and consequences.

[0112] It would be desirable to have the ability to implement an improved airfoil, such as the airfoil 130 described herein, rather than relying on chance. It would be desirable to have a limited or reduced range of possible airfoil configurations to meet mission requirements, including protection, weight limitations, heat transfer, pressure ratio and noise transmission level requirements, and the ability to withstand bird strikes when selecting and positioning the airfoil 130 within the engine.

[0113] The inventors have attempted to find a trade-off balance between leading edge protection and weight gain / loss while meeting all design requirements, including the benefits of implementing the pre-swirl feature 80. Considering all variables will result in a more ideal airfoil that is tailored to the specific needs of the engine, as described above. Understanding these trade-offs is also a great way to save time.

[0114] Table 1 below shows example configurations of a primary airfoil 200a and a secondary airfoil 200b located upstream of the primary airfoil 200a, which provide feasible solutions to the trade-off problem. It should be understood that the primary airfoil 200a can be from a plurality of fan blades 40, while the secondary airfoil 200b can be a plurality of partial span IGVs. For this particular set of examples, SL2 is 25% of SL1.

[0115] Table 1 Example

[0116]

[0117] During the engine design process and the time-consuming iterative process described above, the inventor unexpectedly discovered that there is a relationship between APF1 and APF2. It has been found that the amount of protective coverage required for the primary airfoil 200a is within a specific range based on the amount of protective coverage on the secondary-stage airfoil 200b. In addition, the amount of protective coverage on the primary airfoil 200a is affected by the second span length SL2 of the secondary-stage airfoil 200b. In other words, if the second span length SL2 is 25% of the first span length SL1 as shown in Table 1, the amount of protective coverage required for the primary airfoil 200a will be much less than when the second span length SL2 is 50% of the first span length SL1.

[0118] The inventors have found that the required modification of the primary airfoil relative to the structure and the modification of the secondary airfoil cannot be predicted simply, consistently or reliably by a model. Previously this could only be achieved by a previously time consuming and iterative process. The ratio of the primary leader length and / or the secondary leader length to the primary chord length and / or the secondary chord length is a carefully chosen balance between the leading edge, which dominates the response to a bird ingestion or similar event, and the airfoil material, which dominates the properties of the airfoil aerodynamics in normal operation. The primary airfoil rotates, while the secondary airfoil is stationary, therefore the dynamics of a bird ingestion event are different for the two airfoils, resulting in two ratios of leader length to corresponding chord lengths.

[0119] Table 2 below shows an example continuous airfoil stage with a feasible solution to the trade-off problem. Different span percentages are shown in Table 2. It was found that CL and LL should be taken from anywhere between 20% and 80% (including the end points) of the span length SL of the fan blade. The span % shown in Table 2 is measured as a percentage relative to the root of the airfoil. Therefore, it can be understood that the 20% position of the plurality of fan blades 40 is closest to the hub 48, while the 80% position is farther from the hub 48 and closer to the fan tip 114. The 20% position of the partial span IGV 100 is closest to the nacelle 50, while the 80% position is farther from the nacelle 50 and closer to the IGV tip 104. A specific range of percentages of the span length SL is selected because the airfoil may have different properties, profiles, etc. at its far end. In a non-limiting example, the size of the plurality of fan blades 40 determines APF1, while the size of the partial span IGV 100 determines APF2.

[0120] Table 2

[0121]

[0122] Furthermore, utilizing this relationship, the inventors have discovered that the number of suitable or feasible airfoil possibilities that can be placed in a turbine engine that meet the design requirements can be significantly reduced, thereby facilitating faster selection of airfoils for consideration during engine development. This benefit can provide a deeper understanding of the requirements for a given engine and the requirements for a specific airfoil stage within the engine before specific technical, integration, or system requirements are fully developed. The discovered relationship also avoids or prevents late-stage redesigns while still providing the desired protection effect within given weight parameters for both airfoil stages.

[0123] In addition, the inventors have discovered that the relationship between a first expression of the APF, APF1, and a second expression of the APF, APF2, improves the amount of protection for successive airfoil stages. This relationship was discovered unexpectedly during the engine design process, namely designing multiple stages of airfoil sections, such as fan sections, fan blades, and outlet guide vanes as non-limiting examples, and evaluating the impact of the amount of protection on the fan blades on the desired amount of protection on the outlet guide vanes, and vice versa. The improved airfoil defined as described herein is based in part on the surrounding airfoil properties, such that the APF of the downstream airfoil is affected by the APF of the upstream airfoil. Furthermore, in light of this unexpected result, narrowing the range of options based on the surrounding airfoil stages can significantly reduce material and time costs.

[0124] For example, refer again to Figure 5 , the amount of protection provided by the first leading edge protector 240a on the first airfoil 230a affects the amount of protection required by the second airfoil 230b upstream of the first airfoil 230a. This relationship between multiple-stage airfoils or continuous airfoils (such as 230a and 230b) can be described by a stage protection factor (denoted as "SPF") determined based on the relationship between APF1 and APF2 described above. The SPF is represented by the ratio of the first airfoil protection factor APF1 to the second airfoil protection factor APF2, and expression (3) is as follows:

[0125]

[0126] More specifically, it was found that for any position from 20% to 80% (including the end points) of the span length SL, the desired SPF is greater than or equal to 0.30 and less than or equal to 5.0 (0.30≤SPF≤5.0). The specific range of the span length SL from 20% to 80% reflects the inventor's consideration of the unique properties of the airfoil with respect to its relationship to the inner surface of the fan case near the root, such as the hub at the root and the tip. It was found that for any position from 20% to 80% (including the end points) of the span length SL, the desired results can be achieved when the airfoil SPF is relatively more uniform along the span direction within this range, so the determined ratio is applicable.

[0127] It will be appreciated that due to their size and motion, a rotating fan blade may require more coverage from a leading edge protector than a static airfoil, which would push the relationship ratio into the 0.30 to 5.0 range. This is because the rotating blade has a higher impact kinetic energy and is driven by the rotational speed of the airfoil.

[0128] Using this relationship, the inventors were able to obtain airfoils that perform better in terms of the amount of protection at an acceptable weight increase. The inventors discovered that the SPF of the first group of airfoils and the second group of airfoils located upstream of the first group of airfoils can be narrowed to an SPF range greater than or equal to 0.70 and less than or equal to 4.0 (0.70≤SPF≤4.0). Narrowing the SPF range can provide a deeper understanding of the requirements of a given engine before specific technology, integration, and system requirements are fully developed. For example, as the fan speed decreases, the coverage of the first leading edge 235a by the first leading edge protector 240a can be reduced, so that APF1 is also reduced. In addition, understanding the range of SPFs can prevent or minimize late-stage redesigns, reduce material costs, and save time.

[0129] The SPF value indicates how the amount of protection on a primary airfoil, such as primary airfoil 200a, affects the amount of protection required for any upstream airfoil stages relative to the first set of airfoil stages.

[0130] In one example, Figure 5 The set of fan blades 40 shown in can have the dimensions of the fan blades at 20% in Table 2, and the plurality of partial span IGVs 100 can have the dimensions of the inlet guide vanes at 20% in Table 2. This results in an APF1 value of (11.2 / 46.9) or 0.24 and an APF2 value of (1.3 / 12.7) or 0.10. Using expression (3), the SPF is (0.24 / 0.10) or 2.40.

[0131] In another example, Figure 5 The set of fan blades 40 shown in Table 2 may have the dimensions of the fan blades at 68% of Table 2, while the plurality of partial span IGVs 100 may have the dimensions of the inlet guide vanes at 70% of Table 2. This results in an APF1 value of (15.4 / 61.9) or 0.25 and an APF2 value of (3.8 / 12.7) or 0.30. Using expression (3), the SPF is (0.25 / 0.30) or 0.83.

[0132] Table 3 below provides some lower and upper limit values ​​for each design parameter used to determine expression (3):

[0133]

[0134] There are many trade-offs when developing and designing a turbine engine and all of its corresponding parts. One parameter that may change due to other parameters is the number of fan blades. The number of blades per stage of the engine is a carefully selected value that trades off propulsive efficiency against acoustic characteristics and many other factors, including turbine speed and optional reduction gearbox ratios. As a non-limiting example, a turbine engine with a 35k 1bs thrust may have 16 fan blades with a wider chord than a 70k 1bs thrust turbine engine with 24 blades. The blades with the widest chord will be on a high thrust engine with a small number of fan blades, and the blades with the smallest chord will be on a small engine with many blades, with a lot of potential overlap in the middle. Therefore, it is desirable to have a range that is not necessarily determined by the number of blades, but rather by the interrelationships between the stages.

[0135] The first and second airfoil pairs having dimensions within the ranges listed below in Table 4 were found to conform to the airfoil dimensions described herein above. These ranges minimize weight gain for compact and well protected airfoils.

[0136] Table 4

[0137]

[0138] To the extent provided, first and second airfoil pairs, with the second airfoil located downstream of the first airfoil, may be assembled to conform to a fan section, or other downstream stage relationships within an engine core of blades / vanes and blades / blades.

[0139] The SPF can be used to make tradeoffs when determining the amount of protection on a first airfoil versus the amount of protection on an airfoil located downstream of the first airfoil. For example, when the space available for fan blades in a fan section is limited, knowing these dimensions and the downstream airfoil dimensions can determine an acceptable protector whose leading edge protector length allows for adequate leading edge protection.

[0140] The benefits associated with the SPF described herein include rapid evaluation of the design parameters of the airfoil in the downstream relationship. In addition, the SPF described herein can quickly visualize the trade-offs of geometries, which are defined by the constraints imposed by the materials used, the available space where the airfoil is located, the type of turbine engine or system casing, and the configuration of surrounding components or any other design constraints. The SPF can utilize the available factors to manufacture a high-performance airfoil stage with peak performance. Although it can save time, money and resources to narrow these multiple factors to the possibility area, the greatest benefit is at the system level, where the airfoil stage described herein can improve system performance. The previously developed airfoil stage may reach a peak in one performance field in design, but will lose efficiency or life benefits in another performance field. In other words, the SPF can develop and produce a higher performance airfoil stage across multiple performance indicators within a given set of constraints. The improved airfoil design defined by the APF and SPF relationship takes these factors and the desired results into account.

[0141] To the extent that one or more of the structures provided herein may be known in the art, it should be understood that the present disclosure may include combinations of previously unknown structures to be combined, at least in part for reasons based on conflicts of interest and loss, desired modes of operation, or other forms of teaching in the art.

[0142] This written description uses examples to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and may include other examples that occur to one skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims.

[0143] Further aspects are provided by the subject matter of the following clauses:

[0144] A turbine engine, comprising: an engine core, the engine core defining an engine centerline and including a rotor assembly and a stator assembly; a primary airfoil, the primary airfoil being arranged circumferentially around the engine centerline and defining at least a portion of the rotor assembly, a first airfoil in the primary airfoil comprising: a first main body portion, the first main body portion extending in a chordwise direction between a first body leading edge and a first trailing edge; and a first leading edge protector, the first leading edge protector receiving at least a portion of the first body leading edge of the first main body portion, the first leading edge protector extending a primary leading length (PLL) in the chordwise direction from the first leading edge to the first trailing edge; wherein the first main body portion and the first leading edge protector together define an outer surface of the first airfoil, and extend in a chordwise direction between the first leading edge and the first trailing edge to define a primary chord length (PCL); a secondary airfoil, the secondary airfoil The second airfoil of the secondary stage airfoil is located upstream of the primary airfoil and is arranged circumferentially around the centerline of the engine, the second airfoil in the secondary stage airfoil comprises: a second main body portion, the second main body portion extending in a chordwise direction between a second body leading edge and a second trailing edge; and a second leading edge protector, the second leading edge protector receiving at least a portion of the second body leading edge of the second main body portion, the second leading edge protector extending a secondary leading length (SLL) in the chordwise direction from the second leading edge to the second trailing edge; wherein the second main body portion and the second leading edge protector together define an outer surface of the second airfoil, and extend in a chordwise direction between the second leading edge and the second trailing edge to define a secondary chord length (SCL); wherein the primary leading length (PLL) and the primary chord length (PCL) are associated with the secondary leading length (SLL) and the secondary chord length (SCL) by a stage protection factor (SPF), and wherein, And SPF is greater than or equal to 0.3 and less than or equal to 5 (0.3 <SPF<5)。

[0145] A turbine engine according to any of the preceding clauses, wherein the primary airfoil and the secondary airfoil each comprise a composite airfoil.

[0146] A turbine engine according to any of the preceding clauses, wherein the composite airfoil comprises a polymer matrix composite (PMC).

[0147] A turbine engine according to any of the preceding clauses, wherein at least one of the first leading edge protector or the second leading edge protector is a metallic leading edge protector.

[0148] A turbine engine according to any of the preceding clauses, wherein at least one of the first leading edge protector or the second leading edge protector is a metallic leading edge protector.

[0149] A turbine engine according to any of the preceding clauses, wherein the secondary-stage airfoil is a secondary-stage metal airfoil.

[0150] A turbine engine according to any of the preceding clauses, wherein the primary airfoil comprises a plurality of fan blades.

[0151] A turbine engine according to any preceding clause, wherein the airfoils of the secondary stage airfoils are partial span inlet guide vanes.

[0152] The turbine engine according to any of the preceding clauses, further comprising a nacelle surrounding and at least partially enclosing the plurality of fan blades.

[0153] A turbine engine according to any of the preceding clauses, wherein the partial span inlet guide vanes are integral with the nacelle.

[0154] A turbine engine as described in any preceding clause, wherein the partial span inlet guide vanes are cantilevered.

[0155] A turbine engine as described in any preceding clause, wherein the PLL ranges from 5.0 cm to 20 cm.

[0156] A turbine engine as claimed in any preceding clause, wherein the PCL ranges from 30 cm to 70 cm and the SCL ranges from 0.75 cm to 20 cm.

[0157] A turbine engine as in any preceding clause, wherein said SLL ranges from 0.75 cm to 37.5 cm.

[0158] A turbine engine according to any of the preceding clauses, wherein the first airfoil extends in a span direction between a first root and a first tip to define a first span length, and wherein the second airfoil extends in a span direction between a second root and a second tip to define a second span length, and wherein the SPF is measured from 20% to 80% of the first span length and the second span length.

[0159] A turbine engine according to any of the preceding clauses, wherein at least one of the first leading edge protector or the second leading edge protector is made of polyurethane or a thermoplastic material.

[0160] A turbine engine according to any of the preceding clauses, wherein at least one of the first leading edge protector or the second leading edge protector is made of a composite material.

[0161] A turbine engine according to any of the preceding clauses, wherein the primary airfoil has a first number of airfoils and the secondary airfoil has a second number of airfoils, and wherein the first number is less than the second number.

[0162] The turbine engine of claim 18, wherein the second number is two to three times the first number.

[0163] The turbine engine according to claim 1, wherein the SPF is greater than or equal to 0.7 and less than or equal to 4.0 (0.7≤SPF≤4.0).

[0164] Turbine engine according to any of the preceding clauses, wherein the first leading edge protector and the second leading edge protector each comprise a jacket.

[0165] The turbine engine of any of the preceding clauses, further comprising a tip cap receiving at least a portion of the second tip.

[0166] A turbine engine according to any of the preceding clauses, wherein the tip cap extends the entire axial length of the second tip.

[0167] Turbine engine according to any of the preceding clauses, wherein the first leading edge protector is metallic and the second leading edge protector is made of composite material.

[0168] Turbine engine according to any of the preceding clauses, wherein the first leading edge protector is made of composite material and the second leading edge protector is metallic.

[0169] Turbine engine according to any of the preceding clauses, wherein the first leading edge protector and the second leading edge protector are both made of composite material.

[0170] A turbine engine according to any of the preceding clauses, wherein the first leading edge protector and the second leading edge protector are both metallic.

[0171] A turbine engine according to any of the preceding clauses, wherein the secondary-stage airfoil is a metal airfoil.

[0172] A turbine engine according to any of the preceding clauses, wherein the primary airfoil is a composite airfoil.

Claims

1. A turbine engine, characterized in that: include: an engine core defining an engine centerline and including a rotor assembly and a stator assembly; Primary airfoils arranged circumferentially about the engine centerline and defining at least a portion of the rotor assembly, a first one of the primary airfoils comprising: a first main body portion extending in a chordwise direction between a first body leading edge and a first trailing edge; and a first leading edge protector receiving at least a portion of the first body leading edge of the first main body portion, the first leading edge protector extending a primary leading length (PLL) in a chordwise direction from the first leading edge to the first trailing edge; wherein the first body portion and the first leading edge protector together define an outer surface of the first airfoil and extend in a chordwise direction between the first leading edge and the first trailing edge to define a primary chord length (PCL); A secondary airfoil, the secondary airfoil being located upstream of the primary airfoil and arranged circumferentially around the centerline of the engine, wherein a second airfoil in the secondary airfoil comprises: a second main body portion extending in a chordwise direction between a second body leading edge and a second body trailing edge; and a second leading edge protector receiving at least a portion of the second body leading edge of the second main body portion, the second leading edge protector extending a secondary leading length (SLL) in a chordwise direction from the second leading edge to the second trailing edge; wherein the second body portion and the second leading edge protector together define an outer surface of the second airfoil and extend in a chordwise direction between the second leading edge and the second trailing edge to define a secondary chord length (SCL); wherein the primary preamble length (PLL) and the primary chord length (PCL) are associated with the secondary preamble length (SLL) and the secondary chord length (SCL) by a level protection factor (SPF), and in, And the SPF is greater than or equal to 0.3 and less than or equal to 5 (0.3≤SPF≤5).

2. The turbine engine according to claim 1, characterized in that: in, The primary airfoil and the secondary airfoil each include a composite airfoil.

3. The turbine engine according to claim 2, characterized in that: in, The composite airfoil includes a polymer matrix composite (PMC), and at least one of the first leading edge protector or the second leading edge protector is a metallic leading edge protector.

4. The turbine engine according to claim 1, characterized in that: in, At least one of the first leading edge protector or the second leading edge protector is a metal leading edge protector.

5. The turbine engine according to claim 4, characterized in that in, The secondary-stage airfoil is a secondary-stage metal airfoil.

6. The turbine engine according to claim 1, characterized in that in, The primary airfoil includes a plurality of fan blades.

7. The turbine engine according to claim 6, characterized in that in, The airfoil in the secondary stage airfoil is a partial span inlet guide vane.

8. The turbine engine according to claim 6, characterized in that Further included is a nacelle surrounding and at least partially enclosing the plurality of fan blades.

9. The turbine engine according to claim 1, characterized in that in, The PLL ranges from 5.0 cm to 20 cm, the PCL ranges from 30 cm to 71 cm, the SCL ranges from 0.75 cm to 20 cm, and the SLL ranges from 0.75 cm to 7.5 cm.

10. The turbine engine according to claim 1, characterized in that in, The SPF is greater than or equal to 0.7 and less than or equal to 4.0 (0.7≤SPF≤4.0).

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