Ductless airfoil assembly

Abstract

A ductless airfoil assembly includes an airfoil having spaced apart pressure and suction sides extending radially in a span from a root to a tip and axially in a chord between spaced apart leading and trailing edges. The airfoil defines a forward-most axial point and is disposed about the longitudinal axis and rotates about the longitudinal axis in a rotational direction. The tip leading edge of the airfoil is circumferentially offset in a direction opposite the direction of rotation relative to a circumferential position of the forward most axial point.

Description

Technical Field

[0001] The present subject matter relates generally to components of gas turbine engines, or more particularly, to ductless airfoil assemblies. Background Art

[0002] A gas turbine engine generally includes a fan and a turbine arranged in flow communication with each other. Additionally, the turbine of a gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section in a series flow sequence. In operation, air is provided from the fan to the inlet of the compressor section, where one or more axial compressors gradually compress the air until the air reaches the combustion section. Fuel is mixed with the compressed air and burned in the combustion section to provide combustion gases. The combustion gases are directed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then directed to, for example, the atmosphere through the exhaust section.

[0003] The fan is driven by the turbine. The fan includes a plurality of circumferentially spaced fan blades extending radially outward from the rotor disk. The rotation of the fan blades generates airflow through the inlet of the compressor section of the turbine, and airflow over the turbine. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] 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:

[0005] Figure 1 is a schematic cross-sectional illustration of an exemplary unducted gas turbine engine in accordance with various embodiments of the subject disclosure.

[0006] Figure 2 is a schematic diagram of an exemplary airfoil according to various embodiments of the present disclosure.

[0007] Figure 3 According to various embodiments of the present disclosure, Figure 2 Schematic cross-sectional view taken along line 3-3.

[0008] Figure 4 is a schematic diagram of an exemplary airfoil according to various embodiments of the present disclosure.

[0009] Figure 5 is a graph showing thickness ratio profiles of airfoil sections as a function of chord fraction according to an embodiment of the present disclosure.

[0010] Figure 6 is a schematic diagram of an exemplary airfoil according to another embodiment of the present disclosure.

[0011] Figure 7are schematic diagrams of exemplary airfoil cross sections taken at different radial locations of the airfoil according to an embodiment of the present disclosure.

[0012] Repeat use of reference characters in the present specification and drawings is intended to represent same or analogous features or elements of the present subject matter. DETAILED DESCRIPTION

[0013] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Like or similar reference numbers in the drawings and description have been used to refer to like or similar parts of the present disclosure.

[0014] As used herein, the word "exemplary" means "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0015] As used herein, the terms “first” and “second” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0016] The terms "front" and "rear" refer to relative positions within a turbine, gas turbine engine, or vehicle, and refer to the normal operating attitude of the turbine, gas turbine engine, or vehicle. For example, for a gas turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

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

[0018] Unless otherwise specified herein, the terms "coupled," "fixed," "attached to," and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features.

[0019] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0020] In a context such as “at least one of A, B, and C,” the term “at least one of” means only A, only B, only C, or any combination of A, B, and C.

[0021] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language indicates otherwise, such 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.

[0022] The terms “turbomachine” or “turbomachinery” refer to a machine that includes one or more compressors, a heat generating section (eg, a combustion section), and one or more turbines that together produce a torque output.

[0023] The term "gas turbine engine" refers to an engine having a turbine as all or part of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojets, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.

[0024] As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to the centerline of the gas turbine engine. Additionally, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to the centerline of the gas turbine engine. Additionally, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending in an arc around the centerline of the gas turbine engine.

[0025] In certain aspects of the present disclosure, a ductless airfoil assembly for a turbine is provided. The ductless airfoil assembly generally includes circumferentially spaced airfoils or blades. Each airfoil has a spaced pressure side and a suction side that extend radially from the root to the tip in the span and extend axially between the spaced leading edge and the trailing edge in the chord. The airfoils are arranged about a longitudinal axis and rotate in a rotational direction about the longitudinal axis. In some embodiments, an outer portion or tip portion of the airfoil is configured to have a defined inclination to minimize a cruise flight condition pressure signature. Embodiments of the present disclosure reduce noise by moving the maximum thickness of the airfoil closer to the leading edge in an acoustically sensitive portion of the airfoil.

[0026] Reference now Figure 1 , a schematic cross-sectional view of a gas turbine engine 100 according to an example embodiment of the present disclosure is provided. In particular, Figure 1 A turbofan engine 100 is provided, having a rotor assembly with a single-stage unducted rotor blade. In this manner, the rotor assembly may be referred to herein as a "unducted fan", or the entire gas turbine engine 100 may be referred to as a "unducted turbofan engine". In addition, Figure 1The gas turbine engine 100 includes a third flow path extending from the compressor section to a rotor assembly on the turbine, as will be explained in greater detail below.

[0027] For reference, the gas turbine engine 100 defines an axial direction A, a radial direction R, and a circumferential direction 113. In addition, the gas turbine engine 100 defines an axial centerline or longitudinal axis 112 extending along the axial direction A. Generally, the axial direction A extends parallel to the longitudinal axis 112, the radial direction R extends outwardly and inwardly to the longitudinal axis 112 in a direction orthogonal to the axial direction A, and the circumferential direction extends three hundred and sixty degrees (360°) around the longitudinal axis 112. The gas turbine engine 100 extends, for example, along the axial direction A between a forward end 114 and an aft end 116.

[0028] Gas turbine engine 100 includes a turbine 120 and a rotor assembly (also referred to as a fan section 150) positioned upstream thereof. Generally, turbine 120 includes a compressor section, a combustion section, a turbine section, and an exhaust section in series flow order. In particular, as shown in FIG. Figure 1 As shown, the turbine 120 includes a core shroud 122 defining an annular core inlet 124. The core shroud 122 further at least partially surrounds a low pressure system and a high pressure system. For example, the core shroud 122 shown at least partially surrounds and supports a supercharger or low pressure ("LP") compressor 126 for pressurizing air entering the turbine 120 through the core inlet 124. A high pressure ("HP") multi-stage axial compressor 128 receives the pressurized air from the LP compressor 126 and further increases the pressure of the air. The pressurized air flows downstream to the combustor 130 of the combustion section, where fuel is injected into the pressurized air flow and ignited to increase the temperature and energy level of the pressurized air.

[0029] It will be understood that, as used herein, the terms "high / low speed" and "high / low pressure" are interchangeably used for high pressure / high speed systems and low pressure / low speed systems. Further, it will be understood that the terms "high" and "low" are used in this same context to distinguish between the two systems and are not meant to imply any absolute speed and / or pressure values.

[0030] The high-energy combustion products flow downstream from the combustor 130 to the high-pressure turbine 132. The high-pressure turbine 132 drives the high-pressure compressor 128 via a high-pressure shaft 136. In this regard, the high-pressure turbine 132 is drivingly coupled to the high-pressure compressor 128. The high-energy combustion products then flow to the low-pressure turbine 134. The low-pressure turbine 134 drives the low-pressure compressor 126 and components of the fan section 150 via a low-pressure shaft 138. In this regard, the low-pressure turbine 134 is drivingly coupled to components of the low-pressure compressor 126 and the fan section 150. In this example embodiment, the LP shaft 138 is coaxial with the HP shaft 136. After driving each turbine 132, 134, the combustion products exit the turbine 120 through the turbine exhaust nozzle 140.

[0031] Thus, the turbine 120 defines a working gas flow path or core duct 142 extending between the core inlet 124 and the turbine exhaust nozzle 140. The core duct 142 is an annular duct positioned generally inwardly of the core shroud 122 in the radial direction R. The core duct 142 (e.g., the working gas flow path through the turbine 120) may be referred to as a secondary flow.

[0032] Fan section 150 includes fan 152, which is a primary fan in this example embodiment. Figure 1 In the illustrated embodiment, the fan 152 is an open rotor or unducted fan 152. In this manner, the gas turbine engine 100 may be referred to as an open rotor engine.

[0033] As shown, the fan 152 includes an array of airfoils arranged about the longitudinal axis 112 of the engine 100, and more specifically, includes an array of fan blades 154 arranged about the longitudinal axis 112 of the engine 100 ( Figure 1 Only one is shown in FIG. 1 ). Fan blades 154 are rotatable, for example, about longitudinal axis 112. As described above, fan 152 is drivingly coupled to low pressure turbine 134 via LP shaft 138. Figure 1 In the embodiment shown in FIG. 1 , the fan 152 is coupled to the LP shaft 138 via a reduction gearbox 155 , for example, in an indirect drive or gear drive configuration.

[0034] In addition, the array of fan blades 154 may be arranged equally spaced about the longitudinal axis 112. Each fan blade 154 has a proximal end or root, a distal end or tip, and a span defined therebetween. For purposes of this description, reference will be made to the "tip radius" of the fan blade 154, referred to as R tip Tip radius R tip is the radial distance from longitudinal axis 112 to the outermost radial coordinate, such as tip 157 of fan blade 154, generally where tip 157 intersects the leading edge of fan blade 154. The point at tip 157 will be referred to as tip radius Rtip 100% and the point at the longitudinal axis 112 will be referred to as the tip radius R tip Therefore, the position on the fan blade 154 can be determined based on R / R tip to define (e.g., the point at tip 157 would be defined as 1.0R / R tip , the point at the longitudinal axis 112 will be defined as 0.0R / R tip ). Each fan blade 154 defines a pitch or central blade axis 156. For this embodiment, each fan blade 154 of the fan 152 can be pitched about its central blade axis 156, for example, in unison with each other. One or more actuators 158 are provided to facilitate such rotation and can therefore be used to change the pitch of the fan blades 154 about their respective central blade axes 156.

[0035] Fan section 150 further includes an array of airfoils positioned behind fan blades 154 and also disposed about longitudinal axis 112, and more specifically, includes an array of fan guide vanes 160 including fan guide vanes 162 disposed about longitudinal axis 112 ( Figure 1 For this embodiment, the fan guide vanes 162 are not rotatable about the longitudinal axis 112. Each fan guide vane 162 has a proximal end or root, a distal end or tip, and a span defined therebetween. The fan guide vanes 162 may be as follows: Figure 1 It is shown unshielded, or alternatively, may be shielded, for example, by an annular shroud spaced outwardly from the tip of the fan guide vanes 162 in the radial direction R or attached to the fan guide vanes 162 .

[0036] Each fan guide vane 162 defines a central blade axis 164. For this embodiment, each fan guide vane 162 in the fan guide vane array 160 is capable of rotating about its respective central blade axis 164, e.g., rotating in unison with one another. One or more actuators 166 are provided to facilitate such rotation, and thus can be used to change the pitch of the fan guide vanes 162 about their respective central blade axes 164. However, in other embodiments, each fan guide vane 162 may be fixed or not pitchable about its central blade axis 164. The fan guide vanes 162 are mounted to a fan housing 170.

[0037] like Figure 1As shown, in addition to the unducted fan 152, a ducted fan 184 is also included behind the fan 152, so that the gas turbine engine 100 includes both ducted and unducted fans, both of which are used to generate thrust by the movement of air that does not pass through at least a portion of the turbine 120 (e.g., for the illustrated embodiment, does not pass through the HP compressor 128 and the combustion section). The ducted fan 184 is capable of rotating about the same axis as the fan blades 154 (e.g., the longitudinal axis 112). For the illustrated embodiment, the ducted fan 184 is driven by the low pressure turbine 134 (e.g., coupled to the LP shaft 138). In the illustrated embodiment, as described above, the fan 152 can be referred to as a primary fan and the ducted fan 184 can be referred to as a secondary fan. It will be understood that these terms "primary" and "secondary" are for convenience and do not imply any special importance, authority, etc.

[0038] Ducted fan 184 includes a plurality of fan blades arranged in a single stage ( Figure 1 112 ), so that the ducted fan 184 can be referred to as a single-stage fan. The fan blades of the ducted fan 184 can be arranged equally spaced around the longitudinal axis 112. Each blade of the ducted fan 184 has a proximal end or root, a distal end or tip, and a wingspan defined therebetween.

[0039] The fan cover 170 annularly encloses at least a portion of the core cover 122 and is generally positioned outside of at least a portion of the core cover 122 along the radial direction R. In particular, a downstream section of the fan cover 170 extends over a front portion of the core cover 122 to define a fan duct flow path or simply a fan duct 172. According to this embodiment, the fan flow path or the fan duct 172 may be understood to form at least a portion of the third flow of the engine 100.

[0040] Incoming air may enter through fan duct 172 through fan duct inlet 176 and may be discharged through fan exhaust nozzle 178 to generate propulsive thrust. Fan duct 172 is an annular duct positioned generally outside core duct 142 along radial direction R. Fan cowl 170 and core cowl 122 are connected together and are supported by a plurality of substantially radially extending, circumferentially spaced stationary struts 174 ( Figure 1The fan duct 172 and the core duct 142 may be supported by the fan duct 170 and the core duct 142. The fan duct 172 and the core duct 142 may be supported by the fan duct 170 and the core duct 12 ...42. Each stationary strut 174 may have an aerodynamic profile to guide air to flow therethrough. Other struts except the stationary strut 174 may be used to connect and support the fan duct 170 and / or the core duct 122. In many embodiments, the fan duct 172 and the core duct 142 may extend together (generally axially extend together) at least partially on the opposite sides (e.g., opposite radial sides) of the core duct 122. For example, the fan duct 172 and the core duct 142 may each extend directly from the leading edge 144 of the core duct 122, and may be partially on the opposite radial sides of the core duct 122 and generally extend together axially.

[0041] The gas turbine engine 100 further defines or includes an inlet duct 180. The inlet duct 180 extends between the engine inlet 182 and the core inlet 124 / fan duct inlet 176. The engine inlet 182 is generally defined at the front end of the fan casing 170 and is positioned between the fan 152 and the fan guide vane array 160 along the axial direction A. The inlet duct 180 is an annular duct positioned inside the fan casing 170 along the radial direction R. Air flowing downstream along the inlet duct 180 is not necessarily evenly divided into the core duct 142 and the fan duct 172 by the fan duct splitter or the leading edge 144 of the core casing 122. In the illustrated embodiment, along the radial direction R, the inlet duct 180 is wider than the core duct 142. Along the radial direction R, the inlet duct 180 is also wider than the fan duct 172.

[0042] It is noteworthy that for the illustrated embodiment, the engine 100 includes one or more features to increase the efficiency of the third flow thrust Fn3S (e.g., the thrust generated by the airflow exiting through the fan duct 172, through the fan exhaust nozzle 178, which is at least partially generated by the ducted fan 184). In particular, the engine 100 further includes an array of inlet guide vanes 186, which are positioned in the inlet duct 180, upstream of the ducted fan 184 and downstream of the engine inlet 182. The array of inlet guide vanes 186 is arranged around the longitudinal axis 112. For this embodiment, the inlet guide vanes 186 cannot rotate around the longitudinal axis 112. Each inlet guide vane 186 defines a central blade axis (not marked for clarity) and can be pitched around its corresponding central blade axis, for example, in unison with each other. In this way, the inlet guide vanes 186 can be considered as variable geometry components. One or more actuators 188 are provided to facilitate such rotation, and can therefore be used to change the pitch of the inlet guide vanes 186 around their respective central blade axes. However, in other embodiments, each inlet guide vane 186 may be fixed or non-pitchable about its central blade axis.

[0043] Further, at a location downstream of ducted fan 184 and upstream of fan duct inlet 176, gas turbine engine 100 includes an array of outlet guide vanes 190. Like array of inlet guide vanes 186, array of outlet guide vanes 190 is not rotatable about longitudinal axis 112. However, for the illustrated embodiment, unlike array of inlet guide vanes 186, array of outlet guide vanes 190 is configured as fixed pitch outlet guide vanes.

[0044] Further, it will be appreciated that for the illustrated embodiment, the fan exhaust nozzle 178 of the fan duct 172 is further configured as a variable geometry exhaust nozzle. In this manner, the engine 100 includes one or more actuators 192 for adjusting the variable geometry exhaust nozzle. For example, the variable geometry exhaust nozzle may be configured to vary the total cross-sectional area (e.g., the area of ​​the nozzle in a plane perpendicular to the longitudinal axis 112) to adjust the amount of thrust generated based on one or more engine operating conditions (e.g., temperature, pressure, mass flow, etc. of the airflow through the fan duct 172). Fixed geometry exhaust nozzles may also be employed.

[0045] In addition, still refer to Figure 1 In an exemplary embodiment, the air passing through the fan duct 172 may be relatively cooler (e.g., lower temperature) than one or more fluids used in the turbine 120. In this manner, one or more heat exchangers 200 may be positioned in thermal communication with the fan duct 172. For example, one or more heat exchangers 200 may be disposed within the fan duct 172 and used to cool one or more fluids from the core engine using the air passing through the fan duct 172 as a resource for removing heat from the fluids (e.g., compressor bleed air, oil, or fuel).

[0046] Reference now Figure 2 and Figure 3 , Figure 2 is a schematic partial view of an exemplary ductless airfoil assembly 210 according to various embodiments of the present disclosure, Figure 3 According to various embodiments of the present disclosure, Figure 2 3-3. The exemplary ductless airfoil assembly 210 may be configured to be used as Figure 1 The fan 152 or fan guide vane array 160 of the engine 100 is shown. The ductless airfoil assembly 210 includes an array of blades or airfoils 214 ( Figure 2 Only one is shown), which is surrounded by a fan 152 ( Figure 1) is centered on the longitudinal axis 112 of the rotor. The disk or hub 216 is regularly spaced circumferentially. Each airfoil 214 includes a leading edge 234, a trailing edge 236, a root or proximal end 250 (i.e., a portion extending in the radial direction R toward the longitudinal axis 112 ( Figure 1 ) and the tip 228. Moreover, the tip leading edge 238 of the airfoil 214 is defined as the intersection of the leading edge 234 and the tip 228. Each airfoil 214 extends radially outward from the root or proximal end 250 to the tip 228 along the span "S". For descriptive purposes, as described above, reference will also be made to the "tip radius" of the airfoil 214, referred to as R tip Tip radius R tip is the radial distance from the longitudinal axis 112 to the outermost radial coordinate of the airfoil 214, typically the tip leading edge 238. The point at the tip leading edge 238 will be referred to as the tip radius R tip 100% (or 1.0R / R tip ), and the point at the longitudinal axis 112 will be referred to as the tip radius R tip 0% (or 0.0R / R tip ). In different embodiments, different hub radius ratios may be used. For example, each airfoil 214 defines a tip radius R along the radial direction R from the longitudinal axis 112 to the outermost radial coordinate of the airfoil 214 (generally at the tip leading edge 238). tip , and the hub radius along the radial direction R from the longitudinal axis 112 to the outer radius of the hub 216 defined at the leading edge 234 of the airfoil 214. The hub radius ratio is generally the hub radius divided by the tip radius. As an example, for the outer radius of the hub 216 (in terms of the fan 152 ( Figure 1 ) of the longitudinal axis 112 ( Figure 1 ) is centered) radially at approximately thirty percent (30%) of the tip radius, 0.3R / R tip The value of corresponds to the zero percent (0%) span position. As mentioned above, an R / R of 0.0 tip The values ​​correspond to the longitudinal axis 112. Therefore, it should be appreciated that different hub radius ratios used in conjunction with the airfoil 214 may result in different R / R values ​​corresponding to various features of the airfoil 214 according to the present disclosure. tip Different spanwise coordinate values ​​for the coordinate values.

[0047] The airfoil 214 forms an aerodynamic surface extending in an axial direction A between a leading edge 234 and a trailing edge 236. The airfoil 214 extends outwardly from a root 250 in a radial direction R. In the depicted embodiment, the leading edge 234 includes an inner portion 242 extending outwardly in the radial direction R to a particular spanwise position, a middle portion 244 extending from the inner portion 242 to a tip portion 246, and a tip portion 246 extending from the middle portion 244 to a tip 228 and surrounding the tip 228 and the tip leading edge 238. As used herein, a "tip portion" of an airfoil is defined as a portion of an airfoil extending radially from a location of a forward-most axial point of the leading edge of the airfoil to a tip of the airfoil. For example, in an exemplary embodiment, the forward-most axial point of the leading edge 234 of the airfoil 214 may be radially located at fifty percent (50%) of the tip radius, or 0.5R / R tip 50% or more of the tip radius, or 0.5R / R tip The value of R / R causes the tip portion 246 of the airfoil 214 to change from 0.5R / R tip The radial value of R / R extends to the tip 228 of the airfoil 214. In an exemplary embodiment, the forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at fifty-five percent (55%) of the tip radius, or 0.55 R / R tip 55% (0.55%) of the tip radius, or 0.55R / R tip The value of R / R causes the tip portion 246 of the airfoil 214 to change from 0.55R / R tip The radial value of R / R extends to the tip 228 of the airfoil 214. In an exemplary embodiment, the forward-most axial point of the leading edge 234 of the airfoil 214 may be radially located at fifty-eight percent (58%) of the tip radius, or 0.58 R / R tip 58 percent (58%) of the tip radius, or 0.58R / R tip The value of R / R causes the tip portion 246 of the airfoil 214 to change from 0.58R / R tip The radial value of R / R extends to the tip 228 of the airfoil 214. In an exemplary embodiment, the forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at sixty percent (60%) of the tip radius, or 0.6R / R tip 60% or more of the tip radius, or 0.6R / R tip The value of R / R causes the tip portion 246 of the airfoil 214 to change from 0.6R / R tip20 to the tip 228 of the airfoil 214. In an exemplary embodiment, the forward-most axial point of the leading edge 234 of the airfoil 214 may be radially located at or greater than forty percent (40%) of the span of the airfoil 214, such that the tip portion 246 of the airfoil 214 extends from a radial position of forty percent (40%) of the span of the airfoil 214 to the tip 228 of the airfoil 214. In an exemplary embodiment, the forward-most axial point of the leading edge 234 of the airfoil 214 may be radially located at or greater than forty-five percent (45%) of the span of the airfoil 214, such that the tip portion 246 of the airfoil 214 extends from a radial position of forty-five percent (45%) of the span of the airfoil 214 to the tip 228 of the airfoil 214. In an exemplary embodiment, the forward-most axial point of the leading edge 234 of the airfoil 214 may be radially located at or greater than forty-eight percent (48%) of the span of the airfoil 214, such that the tip portion 246 of the airfoil 214 extends from the radial position of forty-eight percent (48%) of the span of the airfoil 214 to the tip 228 of the airfoil 214. In an exemplary embodiment, the forward-most axial point of the leading edge 234 of the airfoil 214 may be radially located at or greater than fifty percent (50%) of the span of the airfoil 214, such that the tip portion 246 of the airfoil 214 extends from the radial position of fifty percent (50%) of the span of the airfoil 214 to the tip 228 of the airfoil 214. In an exemplary embodiment, the forward-most axial point of the leading edge of the airfoil may vary based on the pitch angle of the airfoil (e.g., for a variable pitch fan). Thus, in an exemplary embodiment, the "tip portion" of the airfoil may be defined as a portion of the airfoil that extends radially from the position of the forward-most axial point of the leading edge of the airfoil to the tip of the airfoil when the airfoil is in its design orientation (e.g., in an orientation representing a subsonic cruise flight speed or operation). For example, cruise is a phase of flight that occurs when an aircraft reaches a set altitude after climbing and before beginning to descend. Thus, as used herein, cruise represents a continuous, high-speed, and stable flight condition in which an aircraft is intended to operate. This description is intended to distinguish cruise from certain abnormal or transient conditions (such as dives) in which an aircraft can reach high flight speeds, but the aircraft is not intended to experience this condition in most missions from takeoff to landing. Thus, a subsonic cruise flight speed may refer to subsonic operation at a flight Mach number of 0.4 or above, or 0.5 or above.Thus, in the exemplary embodiment, the tip portion 246 of the airfoil 214 is defined as a portion of the airfoil 214 that radially extends from the location of the forward-most axial point 240 of the leading edge 234 of the airfoil 214 to the tip 228 of the airfoil 214 when the airfoil 214 is in its design orientation (e.g., in an orientation representative of high subsonic cruise flight speeds or operations). In the illustrated embodiment, the leading edge 234 of the inner portion 242 sweeps forward in the axial direction A, and the leading edge 234 of the intermediate portion 244 begins sweeping rearward in the axial direction A outside of the inner portion 242. For example, the acoustically active spanwise portion of the airfoil 214 may be determined via a relationship between a source intensity distributed radially along the airfoil 214 and a radiation efficiency along the airfoil 214. The acoustically active portion of the airfoil 214 may be determined by multiplying the intensity of the sound source distributed radially along the airfoil 214 by the acoustic Green's function or radiation efficiency (e.g., the ability of a noise source to propagate acoustic energy to the surrounding medium) along the airfoil 214. The radiation efficiency may be any known relationship that describes the effective intensity of a noise source on an airfoil, fan, or propeller blade to an observer location of interest, and may depend on the airfoil shape, size, flow conditions, combinations thereof, etc. In some exemplary embodiments, the trailing edge 236 of the airfoil 214 is configured to have a smooth, curved profile (e.g., without steps or abrupt axial sweep changes / transitions).

[0048] Each airfoil 214 extends from a root or proximal end 250 at the hub 216 to a tip leading edge 238 and includes a generally concave pressure side 252 connected to a generally convex suction side 254 at the leading edge 234 and the trailing edge 236. The airfoils 214 may be shown as being stacked along a spanwise stacking line 256 (e.g., as shown in FIG. Figure 3 256 ). For each individual airfoil section of the airfoil 214, an imaginary straight line referred to as a "chord line" 258 connects the leading edge 234 and the trailing edge 236. In addition, for each individual airfoil section of the airfoil 214, a curve referred to as a "mean camber line" or "mean line" 260 represents a locus of points located midway between the concave pressure side 252 and the convex suction side 254. Typically, the airfoil 214 will include a "twist," which is a feature in which the stacked airfoil sections are rotated relative to each other about the spanwise stacking line 256. Although not shown in the example illustrated, it will be understood that the airfoil 214 may include a "twist," i.e., a twist in the circumferential direction 113 ( Figure 1 ), and “axial sweep”, i.e. the offset in the axial direction A.

[0049] As described above, each airfoil 214 extends radially outward from a root to a tip 228 along a span "S", with a chord (or chord dimension) "C" defined as the length of the chord line 258. The chord dimension may be constant over the span S, or may vary over the span S, as shown. The airfoil section of the airfoil 214 has a mean line angle 262, which refers to the angle between a tangent to the mean line 260 and the longitudinal axis 112. The mean line angle 262 may be measured at any position along the mean line 260. The value of the mean line angle 262 is a function of both the curvature of the mean line 260 and the pitch angle of the airfoil 214. Therefore, the absolute value of the mean line angle 262 will change as the pitch angle of the airfoil 214 changes. However, it will be understood that the overall mean line shape characteristics of the mean line 260 are constant and depend only on the curvature of the airfoil 214.

[0050] The airfoil 214 has a thickness 264, which is the distance measured orthogonally to the mean line 260 between the concave pressure side 252 and the convex suction side 254, which can be measured at any position along the mean line 260. The thickness 264 can be described using a chord fraction, and the value of the chord fraction can be expressed as a percentage. For reference purposes, the associated thickness "T" is measured at a distance "X" behind the leading edge 234, where the distance "X" represents or is defined as a percentage or fraction of the total chord length, referred to herein as a "chord position", "chord-wise position" or "chord fraction". For example, a "0.5" chord fraction represents a position behind the leading edge 234 that is equal to 50% of the total chord length. Therefore, as used herein, the chord fraction refers to the chord-wise distance from the leading edge 234 to the location of the point of interest divided by the chord C. Thus, for example, the leading edge 234 is located at a 0.0 chord fraction and the trailing edge 236 is located at a 1.0 chord fraction. Further, the thickness ratio can be expressed as the absolute value of the thickness T divided by the maximum thickness (T MAX ).

[0051] The thickness of the airfoil section of the airfoil 214 at a particular chordwise location is represented by the diameter of the inscribed circle 266 between the concave pressure side 252 and the convex suction side 254. In some embodiments, the chordwise fractional location (or chordwise fractional distance) of the maximum thickness of the airfoil 214 is the most forward in the tip portion 246 of the airfoil 214. As used herein, "most forward" refers to the fractional distance of the chord from the location of maximum thickness at a given radial location and chordwise cross section of the airfoil 214 from the leading edge 234. For example, in some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, the maximum thickness of at least a portion of the airfoil 214 in the tip portion 246 is between five and forty percent (between 0.05 and 0.40 chord fractions) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, the maximum thickness of at least a portion of the airfoil 214 in the tip portion 246 is between five and thirty percent (between 0.05 and 0.30 chord fractions) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, the maximum thickness of at least a portion of the airfoil 214 in the tip portion 246 is between five and twenty-five percent (between 0.05 and 0.25 chord fractions) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, the maximum thickness of at least a portion of the airfoil 214 in the tip portion 246 is between five and twenty percent (between 0.05 and 0.20 chord fractions) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from leading edge 234 to trailing edge 236, a maximum thickness of at least a portion of airfoil 214 in tip portion 246 is between leading edge 234 and forty percent (0.40 chord fraction) of the chord C measured from leading edge 234. In some embodiments, for a chord C extending from leading edge 234 to trailing edge 236, a maximum thickness of at least a portion of airfoil 214 in tip portion 246 is between leading edge 234 and thirty percent (0.30 chord fraction) of the chord C measured from leading edge 234. In some embodiments, for a chord C extending from leading edge 234 to trailing edge 236, a maximum thickness of at least a portion of airfoil 214 in tip portion 246 is between leading edge 234 and twenty-five percent (0.25 chord fraction) of the chord C measured from leading edge 234. In some embodiments, for a chord C extending from leading edge 234 to trailing edge 236 , a maximum thickness of at least a portion of airfoil 214 in tip portion 246 is between leading edge 234 and twenty percent (0.20 chord fraction) of the chord C measured from leading edge 234 .Thus, in the exemplary embodiment, for a chord-wise cross-section of airfoil 214 , the chord-wise fractional distance of the maximum thickness of airfoil 214 from leading edge 234 is smallest in tip portion 246 .

[0052] In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least twenty-five percent (25%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of forty percent (0.40 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least fifty percent (50%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of forty percent (0.40 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least sixty-seven percent (67%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of forty percent (0.40 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from leading edge 234 to trailing edge 236, a maximum thickness of at least seventy-five percent (75%) of the radial extent of the tip portion 246 of the airfoil 214 is located forward of forty percent (0.40 chord fraction) of the chord C measured from leading edge 234. In some embodiments, the above-recited percentage of the maximum thickness of the tip portion 246 is located adjacent to the tip 228.

[0053] In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least twenty-five percent (25%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of thirty percent (0.30 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least fifty percent (50%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of thirty percent (0.30 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least sixty-seven percent (67%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of thirty percent (0.30 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from leading edge 234 to trailing edge 236, a maximum thickness of at least seventy-five percent (75%) of the radial extent of the tip portion 246 of the airfoil 214 is located forward of thirty percent (0.30 chord fraction) of the chord C measured from leading edge 234. In some embodiments, the above-recited percentage of the maximum thickness of the tip portion 246 is located adjacent to the tip 228.

[0054] In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least twenty-five percent (25%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of twenty-five percent (0.25 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least fifty percent (50%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of twenty-five percent (0.25 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from the leading edge 234 to the trailing edge 236, at least sixty-seven percent (67%) of the maximum thickness of the radial extent of the tip portion 246 of the airfoil 214 is located forward of twenty-five percent (0.25 chord fraction) of the chord C measured from the leading edge 234. In some embodiments, for a chord C extending from leading edge 234 to trailing edge 236, a maximum thickness of at least seventy-five percent (75%) of the radial extent of the tip portion 246 of the airfoil 214 is located forward of twenty-five percent (0.25 chord fraction) of the chord C measured from leading edge 234. In some embodiments, the above-recited percentage of the maximum thickness of the tip portion 246 is located adjacent to the tip 228.

[0055] Further, as described above, the airfoil 214 depicted and described herein may be configured to function as a guide vane 162 ( Figure 1 ).

[0056] refer to Figure 4 , Figure 4 2 is a schematic diagram of an exemplary airfoil 272 of a ductless airfoil assembly 270 according to an embodiment of the present disclosure. In some embodiments, the airfoil 272 may be similar to the airfoil 214 ( Figure 2 and Figure 3 ) to construct. Figure 4 In FIG. 2 , the airfoil 272 is viewed from a rearward direction. The exemplary ductless airfoil assembly 270 may be configured for use as Figure 1 1 and 1. The fan 152 or fan guide vane array 160 of the engine 100 depicted in FIG. The ductless airfoil assembly 270 includes an array of airfoils 272 ( Figure 4 Only one is shown), the airfoil 272 surrounds the fan 152 ( Figure 1 ) of the longitudinal axis 112 ( Figure 1 ) is regularly spaced circumferentially (eg, in the circumferential direction 113) around a rotor disk or hub centered thereon. Each airfoil 272 includes a root or proximal end 274 (ie, extending in the radial direction R toward the longitudinal axis 112 ( Figure 1 ) and a tip portion 275 defining a tip 276 and a tip leading edge 278 (defined at the intersection of the tip 276 and the leading edge 280) such that the span or spanwise direction of the airfoil 272 is defined between the root 274 and the tip 276. As described above, the tip portion 275 is defined as a portion of the airfoil 272 extending radially from the forward-most axial point 292 of the leading edge 280 of the airfoil 272 to the radial position of the tip leading edge 278 of the airfoil 272. The airfoil 272 forms an aerodynamic surface extending in an axial direction between the leading edge 280 and the trailing edge 282. Each fan airfoil 272 defines a central airfoil axis 284. In some embodiments, each fan airfoil 272 is capable of pitching about its central airfoil axis 284.

[0057] In the illustrated embodiment, the airfoil 272 includes a pressure side 286 and a circumferentially or laterally opposed suction side 288. When the airfoil 272 rotates in a rotational direction 290, the pressure side 286 is generally concave and is in front of the generally convex suction side 288. In one aspect of the present disclosure, the airfoil 272 includes certain geometries having specific circumferential inclinations and axial sweep features for the leading edge 280, trailing edge 282, and tip leading edge 278. In an exemplary embodiment, the airfoil 272 includes certain geometries having specific circumferential inclinations and axial sweep features for the leading edge 280, trailing edge 282, and tip leading edge 278 when the airfoil 272 is positioned in its design orientation (e.g., such as in a variable pitch fan, where the airfoil 272 is positioned in an orientation representative of subsonic cruise operation). For example, in Figure 4 In the illustrated embodiment, the airfoil 272 includes a forward-most axial point 292 on the leading edge 280. In the illustrated embodiment, the circumferential coordinates of the tip leading edge 278 are located in a direction opposite to the rotational direction of the airfoil 272 (e.g., in a direction opposite to the rotational direction 290) relative to the circumferential coordinates of the forward-most axial point 292. Additionally, as shown Figure 4 As shown, the circumferential coordinates of the leading edge 280 and the trailing edge 282 of the tip portion 275 are inclined in a direction opposite to the rotational direction of the airfoil 272 (e.g., opposite to the rotational direction 290). In other words, the tip portion 275 is inclined toward the suction side 288 of the airfoil 272. Therefore, in the exemplary embodiment, the entire tip portion 275 is inclined in a direction opposite to the rotational direction 290. Therefore, in the exemplary embodiment, the circumferential coordinates of the tip leading edge 278 relative to the forward-most axial point 292 are circumferentially offset in a direction opposite to the rotational direction 290.

[0058] Further, in some embodiments of the present disclosure, the circumferential coordinates of the leading edge 280 in the radial direction are relative to the circumferential coordinates of the forward most axial point 292 of the leading edge 280 at a distance from the rotor assembly (e.g., the rotor assembly 150 ( Figure 1 )) increases monotonically in the direction of rotation (for example, in a direction away from the rotation direction 290 or in the direction opposite to the rotation direction 290) exceeding a certain R / R tip For example, in some embodiments, the circumferential coordinates of the leading edge 280 are located in the radial direction away from the rotor assembly (e.g., the rotor assembly 150 ( Figure 1 )) in the direction of rotation (e.g., in a direction away from the rotation direction 290 or in the direction opposite to the rotation direction 290) that monotonically increases by more than 0.6 R / R tip In some embodiments, the circumferential coordinates of the leading edge 280 are located in a radial direction away from the rotor assembly (eg, the rotor assembly 150 ( Figure 1 )) in the direction of rotation (e.g., in a direction away from the rotation direction 290 or opposite to the rotation direction 290) that monotonically increases by more than 0.65 R / R tip In some embodiments, the circumferential coordinates of the leading edge 280 are located in a radial direction away from the rotor assembly (eg, the rotor assembly 150 ( Figure 1 )) increases monotonically by more than 0.68 R / R in the direction of rotation (e.g., away from the rotation direction 290 or in the direction opposite to the rotation direction 290). tipIn some embodiments, the circumferential coordinates of the leading edge 280 are located in a radial direction away from the rotor assembly (eg, the rotor assembly 150 ( Figure 1 )) increases monotonically by more than 0.72 R / R in the direction of rotation (e.g., away from the rotation direction 290 or in the direction opposite to the rotation direction 290). tip It should also be appreciated that in exemplary embodiments where airfoil 272 comprises guide vanes (eg, fan guide vanes 162 ), the tilt of airfoil 272 will be in the direction of rotation.

[0059] In some embodiments, relative to the circumferential coordinates of the foremost axial point 292 of the leading edge 280, the circumferential coordinates of the first sub-portion 294 of the leading edge 280 in the tip portion 275 immediately outside the foremost axial point 292 are within the rotor assembly (e.g., rotor assembly 150 ( Figure 1 )) is inclined in the direction of rotation, and for a second sub-portion 296 of a leading edge 280 extending into a tip portion 275 of a tip 276 immediately outside the first sub-portion 294, the circumferential coordinate of the leading edge 280 is in a radial direction, in a direction away from a rotor assembly (e.g., rotor assembly 150 ( Figure 1 )) increases monotonically in the direction of rotation (e.g., in a direction away from or opposite to the direction of rotation 290). In some embodiments, the first subportion 294 comprises less than twenty-five percent (25%) of the tip portion 275. In some embodiments, the first subportion 294 comprises less than fifteen percent (15%) of the tip portion 275. In some embodiments, the first subportion 294 comprises less than ten percent (10%) of the tip portion 275.

[0060] Figure 5 3 is a graph 300 showing a thickness ratio profile of an airfoil section as a function of chord fraction according to an embodiment of the present disclosure. Graph 300 may represent an airfoil 214 ( Figure 2 and Figure 3 ) and airfoil 272 ( Figure 4 In the depicted embodiment, the diagram 300 shows the airfoil 214 / 272 ( Figure 2-4 )'s tip 228 / 276( Figure 2-4 It should be understood that the thickness ratio may correspond to a thickness ratio near the tip of the airfoil (eg, near the tip of the airfoil 214 / 272 ( Figure 4) or elsewhere. The graph 300 includes a horizontal axis 302 graded in units of chord fraction and a vertical axis 304 expressed as a thickness ratio (i.e., the thickness of the airfoil section at the point of interest divided by the maximum thickness of the airfoil section). The airfoil section of the exemplary embodiment is designed for low noise and high efficiency (i.e., at least a portion of the tip portion 246 / 275 of the airfoil 214 ( Figure 2 and Figure 3 ) or airfoil 272 ( Figure 4 ) having a thickness ratio 306 that increases significantly between 0.0 and 0.16 chord fractions. In one embodiment, the airfoil 214 / 272 ( Figure 2-4 ) has a peak or maximum thickness 308 in at least a portion of the tip portion 246 / 275 at approximately the 0.16 chord fraction. Thus, in this example, the airfoil 214 / 272 ( Figure 2-4 ) MAX 0.05 and 0.2 chord fractions. Further, in the depicted embodiment, the thickness ratio 306 is equal to or greater than 0.8 at a chord fraction of 0.05. In other words, in some embodiments, the thickness ratio 306 is equal to or greater than 0.8 at a chord fraction between 0.05 and 0.16. In some embodiments, the thickness ratio 306 is equal to or greater than 0.8 at a chord fraction between 0.05 and 0.15. In some embodiments, the thickness ratio 306 is equal to or greater than 0.8 at a chord fraction between 0.05 and 0.10.

[0061] In some embodiments, the airfoil cross-section (eg, at the airfoil 214 / 272 ( Figure 4 ) in at least a portion of the tip portion 246 / 275 of the wing relative to T within a specific chord fraction range. MAX The chord fraction remains substantially flat. For example, using airfoil 272 ( Figure 4 ) As an example, in some embodiments, in the case of MAX The chord fraction position extends to the position at the airfoil 272 ( Figure 4 )'s leading edge 280( Figure 4 ) or trailing edge 282( Figure 4 ) between the chord fractions of the intermediate chord fractions, the thickness T will not change from T MAX For example, in some embodiments, T MAX Located at 0.16 chord fraction. T of 0.16 MAX Chord fraction and leading edge 280( Figure 4 ) is between 0.08 chord fraction and 0.16 T MAX Chord fraction and trailing edge 282( Figure 4) is 0.58 chord fraction. Therefore, in this embodiment, airfoil 272 ( Figure 4 ) will not be between 0.08 and 0.58 chord fractions from T MAX Therefore, in this embodiment, T / T MAX Between chord fractions 0.08 and 0.58, equal to or greater than 0.9. Thus, in this embodiment, for a particular T MAX The chord fraction, at 280° from the leading edge ( Figure 4 ) and T MAX The intermediate chord fraction positions between the chord fractions extend to the trailing edge 282 ( Figure 4 ) and T MAX The chord score range between the chord scores of the middle chord scores, T / T MAX Equal to or greater than 0.90.

[0062] In some embodiments (e.g., in airfoil 214 / 272 ( Figure 4 ) in at least a portion of the tip portion 246 / 275, T / T MAX Between chord fractions 0.08 and 0.58, equal to or greater than 0.85. For example, in this embodiment and as Figure 5 As shown, the airfoil 272 ( Figure 4 ) has a thickness T between 0.08 and 0.58 chord fractions and does not vary from T MAX Thus, in some embodiments, for a particular T MAX The chord fraction, at 280° from the leading edge ( Figure 4 ) and T MAX The intermediate chord fractions between the chord fractions extend to the trailing edge 282 ( Figure 4 ) and T MAX The chord scores between the middle chord scores within the chord score range, T / T MAX Equal to or greater than 0.85. In some embodiments, the airfoil 272 ( Figure 4 ) has a thickness T between 0.08 and 0.58 chord fractions and does not vary from T MAX Therefore, in this embodiment, for a specific T MAX The chord fraction, at 280° from the leading edge ( Figure 4 ) and T MAX The intermediate chord fractions between the chord fractions extend to the trailing edge 282 ( Figure 4 ) and T MAX The chord scores between the middle chord scores within the chord score range, T / T MAX Equal to or greater than 0.80.

[0063] refer to Figure 6 , Figure 6 4 is a schematic diagram of an exemplary airfoil 422 of a ductless airfoil assembly 420 according to another embodiment of the present disclosure. The airfoil 422 may be similar to the airfoil 214 ( Figure 2 and Figure 3 ) and / or airfoil 272 ( Figure 4 ) to construct. The airfoil 422 may be constructed to be used as Figure 1 A fan 152 or fan guide vane array 160 of engine 100 is depicted.

[0064] exist Figure 6 4, the airfoil 422 includes a sculpted trailing edge feature 436. For example, in the illustrated embodiment, the airfoil 422 includes a leading edge 428, a trailing edge 430, a proximal end or root 424 (ie, a portion of the airfoil 422 extending in the radial direction R toward the longitudinal axis 112 ( Figure 1 ) and a tip 425. Moreover, the intersection of the tip 425 and the leading edge 428 is defined as a tip leading edge 426, so that the span or spanwise direction of the airfoil 422 is defined between the root 424 and the tip 425. The airfoil 422 forms an aerodynamic surface extending along the axial direction A between the leading edge 428 and the trailing edge 430. In the illustrated embodiment, the airfoil 422 includes a shaped trailing edge feature 436 (e.g., a wave-shaped feature or multiple features) at its trailing edge 430, which is configured to promote wake mixing to reduce interaction noise caused by the wake of the airfoil 422 impinging on a downstream stationary airfoil or stator (or stator vane), as described in U.S. Patent No. 8,083,487 B2, which is incorporated herein by reference in its entirety. A baseline 434 trailing edge having a smooth profile is depicted to further illustrate the shaped trailing edge feature 436. Alternatively or additionally, the shaped trailing edge feature 436 may be applied to the engine 100 ( Figure 1 ) on the guide vanes 162 to reduce the broadband noise generated by turbulence in the blade boundary layer convected through its trailing edge.

[0065] refer to Figure 7 , Figure 7 is a schematic diagram of an exemplary cross section of an airfoil 500 taken at different radial positions of the airfoil 500 according to an embodiment of the present disclosure. For example, in an exemplary embodiment, Figure 7 The exemplary cross-sections of airfoil 500 taken at different radial locations of airfoil 500 depicted in FIG. 5 may correspond to airfoil 500 positioned at its design orientation (e.g., in a variable pitch fan, where airfoil 500 is positioned at an orientation representing high subsonic cruise flight speeds or operations). Airfoil 500 may be similar to airfoil 214 ( Figure 2 and Figure 3 )、airfoil 272( Figure 4 ) or airfoil 422 ( Figure 6 ) structure. Figure 7 , the horizontal axis 502 represents the axial direction A (e.g., Figure 7 The vertical axis 504 represents the circumferential direction, the rotation direction of the airfoil 500 is represented by arrow 506, and the intersection of the horizontal axis 502 and the vertical axis 504 represents the pitch change axis 508 of the airfoil 500.

[0066] exist Figure 7 In the embodiment, the airfoil section 510 of the airfoil 500 is at the hub position of the airfoil 500 (eg, the airfoil 214 ( Figure 2 ) wheel hub 216( Figure 2 )), the airfoil section 512 of the airfoil 500 is taken at the forwardmost axial point 514 of the airfoil 500 (eg, the airfoil 214 ( Figure 2 )'s forward axial point 240( Figure 2 )), and the airfoil section 516 of the airfoil 500 is at the tip of the airfoil 500 (eg, the airfoil 214 ( Figure 2 ) tip 228( Figure 2 )) is intercepted. Figure 7 As depicted, in the exemplary embodiment, the tip of the airfoil 500 (eg, the tip of the airfoil 214 ( Figure 2 ) tip 228( Figure 2 The point on the airfoil cross section 516 at )) is located such that its circumferential offset from the pitch change axis 508 is greater than its axial offset from the pitch change axis 508, and wherein the point is located axially rearward of the pitch change axis 508. For example, at Figure 7 , line 520 drawn at an angle of forty-five degrees (45°) relative to horizontal axis 502 represents points having equal circumferential and axial offsets relative to pitch change axis 508. Figure 7 In the portion of the airfoil 500 that is above the horizontal axis 502 and to the right of the vertical axis 504, points that are above and to the left of the line 520 represent a greater circumferential offset relative to the pitch change axis 508 than an axial offset relative to the pitch change axis 508. Additionally, in some embodiments, at least a portion 522 of the airfoil 500 in the tip portion (e.g., as depicted by at least the airfoil cross section 512) is axially forward relative to the pitch change axis 508 and circumferentially away from the rotational direction 506 (e.g., at Figure 7 in the portion above the horizontal axis 502 and to the left of the vertical axis 504).

[0067] Therefore, embodiments of the present disclosure include circumferentially spaced airfoils or blades, wherein the blades in the outer portion or tip portion are configured to have a defined inclination to minimize the cruise flight condition pressure characteristics. Additionally, in some embodiments, the thickness distribution in the outer portion or tip portion of the airfoil is configured to minimize the wake under landing and take-off (LTO) flight conditions. Further, embodiments of the present disclosure provide greater mechanical stability while reducing the noise radiated by the blades. For example, embodiments of the present disclosure have an increased leading edge thickness, which improves the incidence tolerance under non-design flight conditions, thereby reducing noise. In addition, embodiments of the present disclosure move the thickness distribution forward without changing the maximum thickness value in the desired area while maintaining blade weight and reducing noise.

[0068] As can be understood from the description herein, various embodiments of gas turbine engines are provided. Certain of these embodiments may be ductless, single-rotor gas turbine engines or ducted turbofan engines. An example of a ducted turbofan engine can be found in U.S. Patent Application No. 16 / 811,368 filed on March 6, 2020 (published as U.S. Patent Application Publication No. 2021 / 0108597) (FIG. 10, paragraph

[0062] , etc.; an annular fan casing 13 including airfoil blades 21 around a rotating element 20 and blades 31 around a stationary element 30; and including a third flow / fan duct 73 (shown in FIG. 10 and described extensively throughout the application)). Various additional aspects of one or more of these embodiments are discussed below. These exemplary aspects may be combined with one or more exemplary gas turbine engines discussed above with respect to the accompanying drawings.

[0069] For example, in some embodiments of the present disclosure, the engine may include a heat exchanger located in an annular duct, such as a heat exchanger located in the third flow. The heat exchanger may extend substantially continuously in the circumferential direction of the gas turbine engine (e.g., at least about 300 degrees, such as at least about 330 degrees).

[0070] In one or more of these embodiments, during the cruise operating mode, at cruising altitude, the threshold power or disk load of the fan (e.g., a non-ducted single rotor or main forward fan) can be in the range of 25 horsepower per square foot (hp / ft2) or greater. In a particular embodiment of the engine, the structures and methods provided herein produce a power load of between 80 hp / ft2 and 160 hp / ft2 or more at cruising altitude during the cruise operating mode, depending on whether the engine is an open rotor or ducted engine.

[0071] In various embodiments, the engine of the present disclosure is applied to a carrier with a cruising altitude of up to about 65,000ft. In certain embodiments, the cruising altitude is between about 28,000ft and about 45,000ft. In certain embodiments, the cruising altitude is expressed as a flight altitude based on the standard air pressure at sea level, wherein the cruising flight condition is between FL280 and FL650. In another embodiment, the cruising flight condition is between FL280 and FL450. In certain embodiments, the cruising altitude is defined at least based on atmospheric pressure, wherein based on a sea level pressure of about 14.70psia and a sea level temperature of about 59 degrees Fahrenheit, the cruising altitude is between about 4.85psia and about 0.82psia. In another embodiment, the cruising altitude is between about 4.85psia and about 2.14psia. It should be understood that in certain embodiments, the range of the cruising altitude defined by pressure can be adjusted based on different reference sea level pressures and / or sea level temperatures.

[0072] It will be appreciated, therefore, that an engine of such a configuration may be configured to produce at least about 20,000 pounds and less than about 80,000 pounds of thrust during operation at rated speed, such as between about 20,000 pounds and 50,000 pounds of thrust during operation at rated speed, such as between about 25,000 pounds and 40,000 pounds of thrust during operation at rated speed.

[0073] In various exemplary embodiments, the fan may include twelve (12) fan blades. From a load perspective, such a number of blades can allow the span of each blade to be reduced so that the overall diameter of the primary fan can also be reduced (e.g., in one exemplary embodiment, to about twelve feet). That is, in other embodiments, the fan can have any suitable number of blades and any suitable diameter. In certain suitable embodiments, the fan includes at least eight (8) blades. In another suitable embodiment, the fan can have at least twelve (12) blades. In yet another suitable embodiment, the fan can have at least fifteen (15) blades. In yet another suitable embodiment, the fan can have at least eighteen (18) blades. In one or more of these embodiments, the fan includes twenty-six (26) or fewer blades, such as twenty (20) or fewer blades.

[0074] Further, in certain exemplary embodiments, the rotor assembly can define a rotor diameter (or fan diameter) of at least 10 feet, such as at least 11 feet, such as at least 12 feet, such as at least 13 feet, such as at least 15 feet, such as at least 17 feet, such as up to 28 feet, such as up to 26 feet, such as up to 24 feet, such as up to 18 feet.

[0075] In various embodiments, it will be appreciated that the engine includes a ratio of the number of blades to the number of blades that may be less than, equal to, or greater than 1:1. For example, in certain embodiments, the engine includes twelve (12) fan blades and ten (10) blades. In other embodiments, the blade assembly includes a greater number of blades than fan blades. For example, in certain embodiments, the engine includes ten (10) fan blades and twenty-three (23) blades. For example, in certain embodiments, the engine may include a ratio of the number of blades to the number of blades between 2:5 and 2:1, or between 2:4 and 3:2, or between 0.5 and 1.5. The ratio may be adjusted based on a variety of factors, including blade size, to ensure that a desired amount of vortex is removed from the primary fan's airflow. In various embodiments, the number of blades is twenty (20) or less. In still other certain embodiments, the sum of the number of blades and the number of blades is between twenty (20) and thirty (30), or between twenty-four (24) and twenty-eight (28), or between twenty-five (25) and twenty-seven (27). In one embodiment, the engine includes between eleven (11) and sixteen (16) blades. In another embodiment, the engine includes twelve (12) blades and ten (10) vanes. In yet another embodiment, the engine includes between three (3) and twenty (20) blades and between three (3) and twenty (20) vanes. In yet another embodiment, the engine includes an equal number of blades and vanes. In yet another embodiment, the engine includes an equal number of blades and vanes, wherein the number of blades is equal to or less than twenty (20). In various embodiments, the engine includes a combination of a ratio of the number of blades to the number of vanes between 2:5 and 2:1, a difference between the number of blades and the number of vanes between two (2) and negative two (-2), and a number of blades between eleven (11) and sixteen (16). For example, the difference between the number of blades and the number of vanes may correspond to an engine having fourteen (14) blades and sixteen (16) vanes, or fourteen (14) blades and twelve (12) vanes, or sixteen (16) blades and eighteen (18) vanes, or sixteen (16) blades and fourteen (14) vanes, or eleven (11) blades and thirteen (13) vanes, or eleven (11) blades and nine (9) vanes, etc.

[0076] Additionally, in certain exemplary embodiments where the engine includes a tertiary flow and an intermediate fan (a ducted fan behind the main forward fan), the ratio R1 / R2 can be between about 1 and 10, or between 2 and 7, or at least about 3.3, at least about 3.5, at least about 4 and less than or equal to about 7, where R1 is the radius of the primary fan and R2 is the radius of the intermediate fan.

[0077] It will be appreciated that various embodiments of the engine, such as the unducted single rotor engine depicted and described herein, can allow operation at normal subsonic aircraft cruise altitudes of Mach 0.5 or above. In certain embodiments, the engine allows normal aircraft operation between Mach 0.55 and Mach 0.85 at cruise altitude. In yet other specific embodiments, the engine allows normal aircraft operation between Mach 0.75 and Mach 0.85. In certain embodiments, the engine allows rotor blade tip speeds equal to or less than 750 feet per second (fps). In other embodiments, the rotor blade tip speed under cruise flight conditions can be 650 to 900 fps, or 700 to 800 fps.

[0078] The fan of the fan assembly may have a fan pressure ratio (FPR) of 1.04 to 1.20, or in some embodiments 1.05 to 1.1, or in some embodiments less than 1.08, as measured at the fan blades at cruise flight conditions.

[0079] In order to operate a gas turbine engine with a fan having the above characteristics to define the above FPR, a gear assembly can be provided to reduce the rotational speed of the fan assembly relative to a drive shaft (e.g., a low-pressure shaft coupled to a low-pressure turbine). In some embodiments, the gear ratio of the input rotational speed to the output rotational speed is greater than 4.1. For example, in particular embodiments, the gear ratio is in the range of 4.1 to 14.0, in the range of 4.5 to 14.0, or in the range of 6.0 to 14.0. In certain embodiments, the gear ratio is in the range of 4.5 to 12 or in the range of 6.0 to 11.0. Thus, in some embodiments, the fan may be configured to rotate at a speed of 700 to 1500 rpm under cruise flight conditions, while the power turbine (e.g., a low-pressure turbine) is configured to rotate at a speed of 2,500 to 15,000 rpm under cruise flight conditions. In certain embodiments, the fan may be configured to rotate at a speed of 850 to 1,350 rpm under cruise flight conditions, while the power turbine is configured to rotate at a speed of 5,000 to 10,000 rpm under cruise flight conditions.

[0080] For the turbine of the gas turbine engine, the compressor and / or turbine may include various numbers of stages. As disclosed herein, the number of stages includes the number of rotor or blade stages in a particular component (e.g., a compressor or turbine). For example, in some embodiments, the low-pressure compressor may include 1 to 8 stages, the high-pressure compressor may include 8 to 15 stages, the high-pressure turbine may include 1 to 2 stages, and / or the low-pressure turbine (LPT) may include 3 to 7 stages. In particular, the LPT may have 4 stages, or between 4 and 7 stages. For example, in certain embodiments, the engine may include a one-stage low-pressure compressor, an 11-stage high-pressure compressor, a two-stage high-pressure turbine, and a 4-stage or 4-stage LPT between 7 stages. As another example, the engine may include a three-stage low-pressure compressor, a 10-stage high-pressure compressor, a two-stage high-pressure turbine, and a 7-stage low-pressure turbine.

[0081] The core engine is generally enclosed in an outer casing that defines half of the core diameter (Dcore), which can be considered to be the maximum range from the centerline axis (reference of R). In certain embodiments, the engine includes a length (L) from the longitudinal (or axial) front end to the longitudinal rear end. In various embodiments, the engine defines a ratio of L / Dcore that provides reduced installation resistance. In one embodiment, L / Dcore is at least 2. In another embodiment, L / Dcore is at least 2.5. In some embodiments, L / Dcore is less than 5, less than 4, or less than 3. In various embodiments, it should be understood that L / Dcore is used for a non-ducted single-rotor engine.

[0082] Reduced installation drag can further provide improved efficiency, such as improved specific fuel consumption. Additionally or alternatively, reduced drag can provide cruise altitude engine and aircraft operation at the above Mach number at cruise altitude. Still other specific embodiments can provide such benefits by utilizing reduced interaction noise between blade assemblies and vane assemblies and / or reduced overall noise generated by the engine through structures located in the annular duct of the engine.

[0083] Additionally, it should be understood that the range of power loads and / or rotor blade tip speeds may correspond to certain configurations, core sizes, thrust outputs, etc., or other configurations at the core engine. However, as previously mentioned, to the extent that one or more configurations provided herein may be known in the art, it should be understood that the present disclosure may include combinations of configurations not previously known to be combined, at least in part for reasons based on conflicting benefits and penalties, desired modes of operation, or other forms of teaching in the art.

[0084] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the present disclosure, including making and using any device or system and performing any incorporated method. The patentable scope of the present disclosure is defined by the claims, and may include other examples that occur to those 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.

[0085] Although the present disclosure is described as having an exemplary design, the present disclosure may be further modified within the scope of the present disclosure. Therefore, the present application is intended to use its general principles to cover any variation, use or adaptation of the present disclosure. Further, the present application is intended to cover such deviations from the present disclosure within the known or customary practice of the field to which the present disclosure belongs and within the limits of the appended claims.

[0086] Further aspects are provided by the subject matter of the following clauses:

[0087] A ductless airfoil assembly comprising: an airfoil having a spaced apart pressure side and a suction side, the pressure side and the suction side extending radially from a root to a tip in the span and extending axially between spaced apart leading and trailing edges in the chord, and wherein the airfoil comprises a forward-most axial point; and wherein the airfoil is arranged about a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein a tip leading edge of the airfoil is circumferentially offset relative to a circumferential position of the forward-most axial point in a direction opposite to the rotational direction.

[0088] A ductless airfoil assembly as described in the preceding clause, wherein said forward-most axial point is defined when said airfoil is oriented in a design orientation for subsonic cruise operation.

[0089] A ductless airfoil assembly according to any preceding clause, wherein the forward-most axial point is radially located at or greater than fifty percent of a tip radius of the airfoil.

[0090] A ductless airfoil assembly according to any preceding clause, wherein the forward-most axial point is radially located at or greater than fifty-five percent of a tip radius of the airfoil.

[0091] A ductless airfoil assembly according to any preceding clause, wherein the forward-most axial point is radially located at or greater than fifty-eight percent of a tip radius of the airfoil.

[0092] A ductless airfoil assembly according to any preceding clause, wherein the forward-most axial point is radially located at or greater than sixty percent of a tip radius of the airfoil.

[0093] A ductless airfoil assembly according to any preceding clause, wherein the forward-most axial point is radially located at or greater than forty percent of the span of the airfoil.

[0094] A ductless airfoil assembly according to any preceding clause, wherein the forward-most axial point is radially located at or greater than forty-five percent of the span of the airfoil.

[0095] A ductless airfoil assembly according to any preceding clause, wherein the forward-most axial point is radially located at or greater than forty-eight percent of the span of the airfoil.

[0096] A ductless airfoil assembly according to any preceding clause, wherein the forward-most axial point is radially located at or greater than fifty percent of the span of the airfoil.

[0097] A ductless airfoil assembly according to any preceding clause, wherein a thickness of the airfoil is defined as a distance measured between the pressure side and the suction side, and wherein the airfoil comprises a maximum thickness, and wherein a chord-wise fractional distance is defined from the leading edge, and wherein the chord-wise fractional distance to the maximum thickness is smallest in a tip portion of the airfoil.

[0098] A ductless airfoil assembly according to any preceding clause, wherein the airfoil comprises a tip portion extending radially from the foremost axial point to the tip, and wherein, for a chord extending from the leading edge to the trailing edge, the maximum thickness of the airfoil in at least a portion of the tip portion is between five percent and thirty percent of the chord measured from the leading edge.

[0099] A ductless airfoil assembly according to any preceding clause, wherein the trailing edge comprises a shaped trailing edge feature.

[0100] A ductless airfoil assembly according to any preceding clause, wherein the airfoil comprises a tip portion extending radially from the foremost axial point to the tip, and wherein a thickness ratio of an airfoil section is defined as the thickness of the airfoil section divided by the maximum thickness of the airfoil section, and wherein at chord fractions between 0.05 and 0.16, the thickness ratio of any airfoil section in the tip portion is greater than 0.8.

[0101] A ductless airfoil assembly according to any preceding clause, wherein the airfoil comprises a tip portion extending radially from the foremost axial point to the tip, and wherein a maximum thickness of at least a portion of the airfoil in the tip portion is between 0.05 and 0.40 chord fraction.

[0102] A ductless airfoil assembly according to any preceding clause, wherein the maximum thickness of at least a portion of the airfoil in the tip portion is between 0.05 and 0.30 chord fraction.

[0103] A ductless airfoil assembly according to any preceding clause, wherein a maximum thickness of at least a portion of the airfoil in the tip portion is between 0.05 and 0.25 chord fraction.

[0104] A ductless airfoil assembly according to any preceding clause, wherein a maximum thickness of at least a portion of the airfoil in the tip portion is between 0.05 and 0.20 chord fraction.

[0105] A ductless airfoil assembly according to any preceding clause, wherein a maximum thickness of at least a portion of the airfoil in the tip portion is located between the leading edge and a chord fraction of 0.40.

[0106] A ductless airfoil assembly according to any preceding clause, wherein a maximum thickness of at least a portion of the airfoil in the tip portion is located between the leading edge and a chord fraction of 0.30.

[0107] A ductless airfoil assembly according to any preceding clause, wherein a maximum thickness of at least a portion of the airfoil in the tip portion is located between the leading edge and 0.25 chord fraction.

[0108] A ductless airfoil assembly according to any preceding clause, wherein a maximum thickness of at least a portion of the airfoil in the tip portion is located between the leading edge and a chord fraction of 0.20.

[0109] A ductless airfoil assembly according to any preceding clause, wherein at least twenty-five percent (25%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.40 chord fraction measured from the leading edge.

[0110] A ductless airfoil assembly according to any preceding clause, wherein at least fifty percent (50%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.40 chord fraction measured from the leading edge.

[0111] A ductless airfoil assembly according to any preceding clause, wherein at least sixty-seven percent (67%) of a maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.40 chord fraction measured from the leading edge.

[0112] A ductless airfoil assembly according to any preceding clause, wherein at least seventy-five percent (75%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.40 chord fraction measured from the leading edge.

[0113] A ductless airfoil assembly according to any preceding clause, wherein at least twenty-five percent (25%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.30 chord fraction measured from the leading edge.

[0114] A ductless airfoil assembly according to any preceding clause, wherein at least fifty percent (50%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.30 chord fraction measured from the leading edge.

[0115] A ductless airfoil assembly according to any preceding clause, wherein at least sixty-seven percent (67%) of a maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.30 chord fraction measured from the leading edge.

[0116] A ductless airfoil assembly according to any preceding clause, wherein at least seventy-five percent (75%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.30 chord fraction measured from the leading edge.

[0117] A ductless airfoil assembly according to any preceding clause, wherein at least twenty-five percent (25%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.25 chord fraction measured from the leading edge.

[0118] A ductless airfoil assembly according to any preceding clause, wherein at least fifty percent (50%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.25 chord fraction measured from the leading edge.

[0119] A ductless airfoil assembly according to any preceding clause, wherein at least sixty-seven percent (67%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.25 chord fraction measured from the leading edge.

[0120] A ductless airfoil assembly according to any preceding clause, wherein at least seventy-five percent (75%) of the maximum thickness of the radial extent of the tip portion of the airfoil is located forward of 0.25 chord fraction measured from the leading edge.

[0121] A ductless airfoil assembly according to any preceding clause, wherein the airfoil is a guide vane.

[0122] A ductless airfoil assembly according to any preceding clause, wherein the airfoil is arranged about a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein the leading edge in the tip portion is inclined in a direction opposite to the rotational direction.

[0123] A ductless airfoil assembly according to any preceding clause, wherein a chord-wise fractional distance from the leading edge of a maximum thickness of an airfoil section of the airfoil is smallest in a tip portion of the airfoil.

[0124] A ductless airfoil assembly according to any preceding clause, wherein the chord-wise fractional location of maximum thickness of the airfoil is located most forwardly in the tip portion as measured from the leading edge.

[0125] A ductless airfoil assembly according to any preceding clause, wherein the airfoil is arranged about a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein the airfoil comprises a tip portion, and wherein the leading edge in the tip portion is inclined in a direction opposite to the rotational direction.

[0126] A ductless airfoil assembly according to any preceding clause, wherein the airfoil is arranged about a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein the airfoil comprises a tip portion, and wherein a circumferential coordinate of the leading edge in the tip portion is inclined in a direction opposite to the rotational direction.

[0127] A ductless airfoil assembly according to any preceding clause, wherein the circumferential coordinates of the leading edge increase monotonically in a radial direction away from the direction of rotation of the blade relative to the circumferential coordinates of the foremost axial point of the leading edge above an R / Rtip value of 0.6.

[0128] A ductless airfoil assembly according to any preceding clause, wherein the circumferential coordinates of the leading edge increase monotonically in a radial direction away from the direction of rotation of the blade relative to the circumferential coordinates of the foremost axial point of the leading edge above an R / Rtip value of 0.65.

[0129] A ductless airfoil assembly according to any preceding clause, wherein the circumferential coordinates of the leading edge increase monotonically in a radial direction away from the direction of rotation of the blade relative to the circumferential coordinates of the foremost axial point of the leading edge above an R / Rtip value of 0.68.

[0130] A ductless airfoil assembly according to any preceding clause, wherein the circumferential coordinates of the leading edge increase monotonically in a radial direction away from the direction of rotation of the blade relative to the circumferential coordinates of the foremost axial point of the leading edge above an R / Rtip value of 0.72.

[0131] A ductless airfoil assembly according to any preceding clause, wherein for at least a sub-portion of the tip portion, the circumferential coordinates of the leading edge increase monotonically in a radial direction relative to the circumferential coordinates of the foremost axial point of the leading edge in a direction away from the direction of rotation of the airfoil beyond the value of the tip radius of the foremost axial point.

[0132] A ductless airfoil assembly according to any preceding clause, wherein the circumferential coordinates of a first sub-portion of the leading edge in the tip portion of the airfoil immediately outside the foremost axial point are inclined in the rotational direction of the airfoil, and for a second sub-portion of the leading edge in the tip portion extending to the tip immediately outside the first sub-portion, the circumferential coordinates of the leading edge increase monotonically in a radial direction, in a direction away from the rotational direction.

[0133] The ductless airfoil assembly of any preceding clause, wherein the first sub-portion comprises less than twenty-five percent (25%) of the tip portion.

[0134] The ductless airfoil assembly of any preceding clause, wherein the first sub-portion comprises less than fifteen percent (15%) of the tip portion.

[0135] The ductless airfoil assembly of any preceding clause, wherein the first sub-portion comprises less than ten percent (10%) of the tip portion.

[0136] A ductless airfoil assembly according to any preceding clause, wherein the airfoil comprises a tip portion, and wherein a thickness ratio of an airfoil section is defined as the thickness of the airfoil section divided by a maximum thickness of the airfoil section, and wherein at a chord fraction of 0.16, the thickness ratio of the tip portion is greater than 0.8.

[0137] A ductless airfoil assembly according to any preceding clause, wherein the maximum thickness of an airfoil section of the airfoil is between 0.05 and 0.2 chord fraction measured from the leading edge.

[0138] A ductless airfoil assembly according to any preceding clause, wherein at a chord fraction of 0.05 measured from the leading edge, the thickness ratio is equal to or greater than 0.8.

[0139] A ductless airfoil assembly according to any preceding clause, wherein at a chord fraction between 0.05 and 0.16 measured from the leading edge, the thickness ratio is equal to or greater than 0.8.

[0140] A ductless airfoil assembly according to any preceding clause, wherein at a chord fraction between 0.05 and 0.15 measured from the leading edge, the thickness ratio is equal to or greater than 0.8.

[0141] A ductless airfoil assembly according to any preceding clause, wherein at a chord fraction between 0.05 and 0.10 measured from the leading edge, the thickness ratio is equal to or greater than 0.8.

[0142] A ductless airfoil assembly comprising: an airfoil having a root, a mid-section and a tip portion, the airfoil having spaced apart pressure and suction sides, the pressure and suction sides extending radially from the root in span to a tip defined in the tip portion and extending axially between spaced apart leading and trailing edges in chord; and wherein a thickness of the airfoil is defined as a distance measured between the pressure and suction sides, and wherein a thickness ratio is defined as the thickness of an airfoil cross section divided by a maximum thickness at the airfoil cross section, and wherein the thickness ratio in the tip portion is equal to or greater than 0.85 over a chord fraction range extending from a chord fraction position midway between the leading edge and the chord fraction position of maximum thickness to a chord fraction position midway between the trailing edge and the chord fraction position of maximum thickness.

[0143] A ductless airfoil assembly according to any preceding clause, wherein the thickness ratio of the airfoil section in the tip portion is equal to or greater than 0.8 over a range of chord fractions extending from the chord fraction location of the maximum thickness to a chord fraction intermediate between the leading edge or the trailing edge.

[0144] A ductless airfoil assembly according to any preceding clause, wherein the thickness of the blade in the tip portion remains within ten percent of the maximum thickness between 0.08 and 0.58 chord fractions measured from the leading edge.

[0145] A ductless airfoil assembly according to any preceding clause, wherein the thickness ratio is equal to or greater than 0.9 between chord fractions of 0.08 and 0.58 measured from the leading edge.

[0146] A ductless airfoil assembly according to any preceding clause, wherein over a chord fraction range extending from a chord fraction position midway between the leading edge and the chord fraction position of maximum thickness to a chord fraction position midway between the trailing edge and the chord fraction position of maximum thickness, the thickness ratio in the tip portion is equal to or greater than 0.90.

[0147] A ductless airfoil assembly according to any preceding clause, wherein the thickness ratio in the tip portion is equal to or greater than 0.85 between chord fractions of 0.08 and 0.58 measured from the leading edge.

[0148] A ductless airfoil assembly according to any preceding clause, wherein the thickness of the airfoil in the tip portion remains within twenty percent of the maximum thickness of the airfoil in the tip portion between 0.08 and 0.58 chord fractions measured from the leading edge.

[0149] A ductless airfoil assembly according to any preceding clause, wherein the position of a point on the airfoil cross section at the tip at a twenty-five percent (25%) chord fraction on a mean line of the airfoil cross section is such that the circumferential offset of the point from a pitch change axis of the airfoil is greater than the axial offset of the point from the pitch change axis, and wherein the point is located axially rearward of the pitch change axis.

[0150] A ductless airfoil assembly according to any preceding clause, wherein at least a portion of the airfoil in the tip portion is axially forward relative to a pitch change axis of the airfoil and circumferentially away from the direction of rotation of the airfoil.

Claims

1. A ductless airfoil assembly, characterized in that: include: an airfoil having spaced apart pressure and suction sides extending radially from a root to a tip in the span and axially between spaced apart leading and trailing edges in the chord, and wherein the airfoil includes a forward-most axial point; and wherein the airfoil is arranged about a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein a tip leading edge of the airfoil is circumferentially offset in a direction opposite to the rotational direction relative to a circumferential position of the foremost axial point.

2. The ductless airfoil assembly according to claim 1, wherein: wherein a thickness of the airfoil is defined as a distance measured between the pressure side and the suction side, and wherein the airfoil comprises a maximum thickness, and wherein a chord-wise fractional distance is defined from the leading edge, and wherein the chord-wise fractional distance to the maximum thickness is smallest in a tip portion of the airfoil.

3. The ductless airfoil assembly according to claim 1, wherein: wherein the airfoil includes a tip portion extending radially from the foremost axial point to the tip, and wherein, for a chord extending from the leading edge to the trailing edge, a maximum thickness of at least a portion of the airfoil in the tip portion is between five percent and forty percent of the chord measured from the leading edge.

4. The ductless airfoil assembly according to claim 1, wherein: Wherein the trailing edge includes a shaped trailing edge feature.

5. The ductless airfoil assembly according to claim 1, wherein: wherein the airfoil includes a tip portion extending radially from the foremost axial point to the tip, and wherein a thickness ratio of an airfoil section is defined as a thickness of the airfoil section divided by a maximum thickness at the airfoil section, and wherein at a chord fraction between 0.05 and 0.16, the thickness ratio of any airfoil section in the tip portion is greater than 0.

8.

6. The ductless airfoil assembly according to claim 1, wherein: wherein the airfoil includes a tip portion extending radially from the foremost axial point to the tip, and wherein a maximum thickness of at least a portion of the airfoil in the tip portion is between 0.05 and 0.30 chord fractions.

7. The ductless airfoil assembly according to claim 1, wherein: wherein the airfoil includes a tip portion extending radially from the foremost axial point to the tip, wherein a thickness ratio of an airfoil section is defined as a thickness of the airfoil section divided by a maximum thickness at the airfoil section, and wherein the thickness ratio of the airfoil section of at least a portion of the airfoil in the tip portion is equal to or greater than 0.8 over a chord fraction range extending from a chord fraction location of the maximum thickness to a chord fraction intermediate between the leading edge or the trailing edge.

8. A ductless airfoil assembly, characterized in that: include: an airfoil having spaced apart pressure and suction sides extending radially from a root to a tip in a span and extending axially between spaced apart leading and trailing edges in a chord, and wherein the airfoil includes a foremost axial point, and wherein the airfoil defines a tip portion extending radially from the foremost axial point to the tip; and Wherein a thickness of the airfoil is defined as a distance measured between the pressure side and the suction side, and wherein a maximum thickness of at least a portion of the airfoil in the tip portion is between 0.05 and 0.40 chord fractions.

9. The ductless airfoil assembly according to claim 8, wherein: wherein the airfoil is arranged about a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein the leading edge in the tip portion is inclined in a direction opposite to the rotational direction.

10. The ductless airfoil assembly according to claim 8, wherein: wherein a thickness ratio of an airfoil section is defined as the thickness of the airfoil section divided by a maximum thickness of the airfoil section, and wherein at a chord fraction of 0.05, the thickness ratio of any airfoil section in the tip portion is 0.8 or greater.

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