UNDUCTED AIRFOIL ASSEMBLY

Abstract

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

Description

FIELD

The present subject matter relates generally to components of a gas turbine engine, or more particularly to an unducted airfoil assembly.

BACKGROUND

A gas turbine engine generally includes a fan and a turbomachine arranged in flow communication with one another. Additionally, the turbomachine of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.

The fan is driven by the turbomachine. The fan includes a plurality of circumferentially spaced fan blades extending radially outward from a rotor disk. Rotation of the fan blades creates an airflow through the inlet to the compressor section of the turbomachine, as well as an airflow over the turbomachine.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present disclosure, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 is a schematic, cross-sectional view of an exemplary, unducted gas turbine engine according to various embodiments of the present subject disclosure.

FIG. 2 is a schematic view of an exemplary airfoil according to various embodiments of the present disclosure.

FIG. 3 is a schematic sectional view taken along line 3-3 of FIG. 2 in accordance with various embodiments of the present disclosure.

FIG. 4 is a schematic view of an exemplary airfoil according to various embodiments of the present disclosure.

FIG. 5 is a graph illustrating a thickness ratio profile of an airfoil section as a function of a chord fraction according to embodiments of the present disclosure.

FIG. 6 is a schematic view of an exemplary airfoil according to another embodiment of the present disclosure.

FIG. 7 is a schematic view of exemplary airfoil sections taken at different radial locations of an airfoil according to embodiments of the present disclosure.

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

DETAILED DESCRIPTION

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

The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless specifically identified otherwise, all embodiments described herein should be considered exemplary.

As used herein, the terms “first” and “second” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a turbomachine, gas turbine engine, or vehicle and refer to the normal operational attitude of the same. For example, with regard to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.

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

The terms “coupled”, “fixed”, “attached to”, and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.

The term “at least one of” in the context of, e.g., “at least one of A, B, and C” refers to only A, only B, only C, or any combination of A, B, and C.

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

The term “turbomachine” or “turbomachinery” refers to a machine including one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output.

The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.

As used herein, the terms “axial” and “axially” refer to directions and orientations that extend substantially parallel to a centerline of the gas turbine engine. Moreover, the terms “radial” and “radially” refer to directions and orientations that extend substantially perpendicular to the centerline of the gas turbine engine. In addition, as used herein, the terms “circumferential” and “circumferentially” refer to directions and orientations that extend arcuately about the centerline of the gas turbine engine.

In certain aspects of the present disclosure, an unducted airfoil assembly for a turbomachine is provided. The unducted airfoil assembly generally includes circumferentially spaced airfoils or blades. Each airfoil has spaced-apart pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced apart leading and trailing edges. The airfoil is arranged around a longitudinal axis and rotates about the longitudinal axis in a rotational direction. In some embodiments, the outer or tip portion of the airfoil is configured having a defined lean to minimize cruise flight condition pressure signatures. Embodiments of the present disclosure reduce noise by moving a maximum thickness of the airfoil closer to the leading edge in the acoustically sensitive portions of the airfoil.

Referring now to FIG. 1, a schematic cross-sectional view of a gas turbine engine 100 is provided according to an example embodiment of the present disclosure. Particularly, FIG. 1 provides a turbofan engine 100 having a rotor assembly with a single stage of unducted rotor blades. In such a manner, the rotor assembly may be referred to herein as an “unducted fan,” or the entire gas turbine engine 100 may be referred to as an “unducted turbofan engine.” In addition, the gas turbine engine 100 of FIG. 1 includes a third stream extending from the compressor section to a rotor assembly flowpath over a turbomachine, as will be explained in more detail below.

For reference, the gas turbine engine 100 defines an axial direction A, a radial direction R, and a circumferential direction 113. Moreover, the gas turbine engine 100 defines an axial centerline or longitudinal axis 112 that extends along the axial direction A. In general, the axial direction A extends parallel to the longitudinal axis 112, the radial direction R extends outward from and inward to the longitudinal axis 112 in a direction orthogonal to the axial direction A, and the circumferential direction extends three hundred sixty degrees (360°) around the longitudinal axis 112. The gas turbine engine 100 extends between a forward end 114 and an aft end 116, e.g., along the axial direction A.

The gas turbine engine 100 includes a turbomachine 120 and a rotor assembly, also referred to as a fan section 150, positioned upstream thereof. Generally, the turbomachine 120 includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. Particularly, as shown in FIG. 1, the turbomachine 120 includes a core cowl 122 that defines an annular core inlet 124. The core cowl 122 further encloses at least in part a low pressure system and a high pressure system. For example, the core cowl 122 depicted encloses and supports at least in part a booster or low pressure (“LP”) compressor 126 for pressurizing the air that enters the turbomachine 120 through core inlet 124. A high pressure (“HP”), multi-stage, axial-flow compressor 128 receives pressurized air from the LP compressor 126 and further increases the pressure of the air. The pressurized air stream flows downstream to a combustor 130 of the combustion section where fuel is injected into the pressurized air stream and ignited to raise the temperature and energy level of the pressurized air.

It will be appreciated that as used herein, the terms “high/low speed” and “high/low pressure” are used with respect to the high pressure/high speed system and low pressure/low speed system interchangeably. Further, it will be appreciated that the terms “high” and “low” are used in this same context to distinguish the two systems, and are not meant to imply any absolute speed and/or pressure values.

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

Accordingly, the turbomachine 120 defines a working gas flowpath or core duct 142 that extends between the core inlet 124 and the turbomachine exhaust nozzle 140. The core duct 142 is an annular duct positioned generally inward of the core cowl 122 along the radial direction R. The core duct 142 (e.g., the working gas flowpath through the turbomachine 120) may be referred to as a second stream.

The fan section 150 includes a fan 152, which is the primary fan in this example embodiment. For the depicted embodiment of FIG. 1, the fan 152 is an open rotor or unducted fan 152. In such a manner, the gas turbine engine 100 may be referred to as an open rotor engine.

As depicted, the fan 152 includes an array of airfoils arranged around the longitudinal axis 112 of engine 100, and more particularly includes an array of fan blades 154 (only one shown in FIG. 1) arranged around the longitudinal axis 112 of engine 100. The fan blades 154 are rotatable, e.g., about the longitudinal axis 112. As noted above, the fan 152 is drivingly coupled with the low pressure turbine 134 via the LP shaft 138. For the embodiments shown in FIG. 1, the fan 152 is coupled with the LP shaft 138 via a speed reduction gearbox 155, e.g., in an indirect-drive or geared-drive configuration.

Moreover, the array of fan blades 154 can be arranged in equal spacing around the longitudinal axis 112. Each fan blade 154 has a proximal end or root and a distal end or tip and a span defined therebetween. For descriptive purposes, reference will be made to a “tip radius”, referred to as R_tip, of the fan blade 154. The tip radius R_tip is the radial distance from the longitudinal axis 112 to the outermost radial coordinate, such as a tip 157 of the fan blade 154, typically where the tip 157 intersects the leading edge of the fan blade 154. A point located at the tip 157 would be referred to as 100% of tip radius R_tip, and a point at the longitudinal axis 112 would be referred to as 0% of tip radius R_tip. Thus, a location on the fan blade 154 may be defined in terms of R/R_tip (e.g., a point at the tip 157 would be defined as 1.0 R/R_tip and a point at the longitudinal axis 112 would be defined as 0.0 R/R_tip). Each fan blade 154 defines a pitch change or central blade axis 156. For this embodiment, each fan blade 154 of the fan 152 is pitchable about its central blade axis 156, e.g., in unison with one another. One or more actuators 158 are provided to facilitate such rotation and therefore may be used to change a pitch of the fan blades 154 about their respective central blade axes 156.

The fan section 150 further includes an array of airfoils positioned aft of the fan blades 154 and also disposed around longitudinal axis 112, and more particularly includes a fan guide vane array 160 that includes fan guide vanes 162 (only one shown in FIG. 1) disposed around the longitudinal axis 112. 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 and a distal end or tip and a span defined therebetween. The fan guide vanes 162 may be unshrouded as shown in FIG. 1 or, alternatively, may be shrouded, e.g., by an annular shroud spaced outward from the tips of the fan guide vanes 162 along the radial direction R or attached to the fan guide vanes 162.

Each fan guide vane 162 defines a central blade axis 164. For this embodiment, each fan guide vane 162 of the fan guide vane array 160 is rotatable about its respective central blade axis 164, e.g., in unison with one another. One or more actuators 166 are provided to facilitate such rotation and therefore may be used to change a pitch of the fan guide vane 162 about its respective central blade axis 164. However, in other embodiments, each fan guide vane 162 may be fixed or unable to be pitched about its central blade axis 164. The fan guide vanes 162 are mounted to a fan cowl 170.

As shown in FIG. 1, in addition to the fan 152, which is unducted, a ducted fan 184 is included aft of the fan 152, such that the gas turbine engine 100 includes both a ducted and an unducted fan which both serve to generate thrust through the movement of air without passage through at least a portion of the turbomachine 120 (e.g., without passage through the HP compressor 128 and combustion section for the embodiment depicted). The ducted fan 184 is rotatable about the same axis (e.g., the longitudinal axis 112) as the fan blade 154. The ducted fan 184 is, for the embodiment depicted, driven by the low pressure turbine 134 (e.g. coupled to the LP shaft 138). In the embodiment depicted, as noted above, the fan 152 may be referred to as the primary fan, and the ducted fan 184 may be referred to as a secondary fan. It will be appreciated that these terms “primary” and “secondary” are terms of convenience, and do not imply any particular importance, power, or the like.

The ducted fan 184 includes a plurality of fan blades (not separately labeled in FIG. 1) arranged in a single stage, such that the ducted fan 184 may be referred to as a single stage fan. The fan blades of the ducted fan 184 can be arranged in equal spacing around the longitudinal axis 112. Each blade of the ducted fan 184 has a proximal end or root and a distal end or tip and a span defined therebetween.

The fan cowl 170 annularly encases at least a portion of the core cowl 122 and is generally positioned outward of at least a portion of the core cowl 122 along the radial direction R. Particularly, a downstream section of the fan cowl 170 extends over a forward portion of the core cowl 122 to define a fan duct flowpath, or simply a fan duct 172. According to this embodiment, the fan flowpath or fan duct 172 may be understood as forming at least a portion of the third stream of the engine 100.

Incoming air may enter through the fan duct 172 through a fan duct inlet 176 and may exit through a fan exhaust nozzle 178 to produce propulsive thrust. The fan duct 172 is an annular duct positioned generally outward of the core duct 142 along the radial direction R. The fan cowl 170 and the core cowl 122 are connected together and supported by a plurality of substantially radially-extending. circumferentially-spaced stationary struts 174 (only one shown in FIG. 1). The stationary struts 174 may each be aerodynamically contoured to direct air flowing thereby. Other struts in addition to the stationary struts 174 may be used to connect and support the fan cowl 170 and/or core cowl 122. In many embodiments, the fan duct 172 and the core duct 142 may at least partially co-extend (generally axially) on opposite sides (e.g., opposite radial sides) of the core cowl 122. For example, the fan duct 172 and the core duct 142 may each extend directly from a leading edge 144 of the core cowl 122 and may partially co-extend generally axially on opposite radial sides of the core cowl 122.

The gas turbine engine 100 also defines or includes an inlet duct 180. The inlet duct 180 extends between an engine inlet 182 and the core inlet 124/fan duct inlet 176. The engine inlet 182 is defined generally at the forward end of the fan cowl 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 that is positioned inward of the fan cowl 170 along the radial direction R. Air flowing downstream along the inlet duct 180 is split, not necessarily evenly, into the core duct 142 and the fan duct 172 by a fan duct splitter or leading edge 144 of the core cowl 122. In the embodiment depicted, 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 along the radial direction R.

Notably, for the embodiment depicted, the engine 100 includes one or more features to increase an efficiency of a third stream thrust, Fn3S (e.g., a thrust generated by an airflow through the fan duct 172 exiting through the fan exhaust nozzle 178, generated at least in part by the ducted fan 184). In particular, the engine 100 further includes an array of inlet guide vanes 186 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 are arranged around the longitudinal axis 112. For this embodiment, the inlet guide vanes 186 are not rotatable about the longitudinal axis 112. Each inlet guide vanes 186 defines a central blade axis (not labeled for clarity), and is pitchable about its respective central blade axis, e.g., in unison with one another. In such a manner, the inlet guide vanes 186 may be considered a variable geometry component. One or more actuators 188 are provided to facilitate such rotation and therefore may be used to change a pitch of the inlet guide vanes 186 about their respective central blade axes. However, in other embodiments, each inlet guide vanes 186 may be fixed or unable to be pitched about its central blade axis.

Further, located downstream of the ducted fan 184 and upstream of the fan duct inlet 176, the gas turbine engine 100 includes an array of outlet guide vanes 190. As with the array of inlet guide vanes 186, the array of outlet guide vanes 190 are not rotatable about the longitudinal axis 112. However, for the embodiment depicted, unlike the array of inlet guide vanes 186, the array of outlet guide vanes 190 are configured as fixed-pitch outlet guide vanes.

Further, it will be appreciated that for the embodiment depicted, the fan exhaust nozzle 178 of the fan duct 172 is further configured as a variable geometry exhaust nozzle. In such a manner, the engine 100 includes one or more actuators 192 for modulating the variable geometry exhaust nozzle. For example, the variable geometry exhaust nozzle may be configured to vary a total cross-sectional area (e.g., an area of the nozzle in a plane perpendicular to the longitudinal axis 112) to modulate an amount of thrust generated based on one or more engine operating conditions (e.g., temperature, pressure, mass flowrate, etc. of an airflow through the fan duct 172). A fixed geometry exhaust nozzle may also be adopted.

Moreover, referring still to FIG. 1, in exemplary embodiments, air passing through the fan duct 172 may be relatively cooler (e.g., lower temperature) than one or more fluids utilized in the turbomachine 120. In this way, 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 utilized to cool one or more fluids from the core engine with the air passing through the fan duct 172, as a resource for removing heat from a fluid, e.g., compressor bleed air, oil or fuel.

Referring now to FIGS. 2 and 3, FIG. 2 is a schematic and fragmentary view of an exemplary unducted airfoil assembly 210 in accordance with various embodiments of the present disclosure, and FIG. 3 is a schematic sectional view taken along line 3-3 of FIG. 2 in accordance with various embodiments of the present disclosure. The exemplary unducted airfoil assembly 210 may be configured for use as the fan 152 or the fan guide vane array 160 of the engine 100 as depicted in FIG. 1. The unducted airfoil assembly 210 includes an array of blades or airfoils 214 (only one shown in FIG. 2) that are regularly spaced apart circumferentially around a disk or hub 216 of a rotor centered on the longitudinal axis 112 of the fan 152 (FIG. 1). Each airfoil 214 includes a leading edge 234, a trailing edge 236, a root or proximal end 250 (i.e., an inboard end in the radial direction R toward the longitudinal axis 112 (FIG. 1)), and a tip 228. Also, a tip leading edge 238 of the airfoil 214 is defined as an intersection of the leading edge 234 with the tip 228. Each airfoil 214 extends radially outward along a span “S” from the root or proximal end 250 to the tip 228. For descriptive purposes, and as described above, reference will also be made to a “tip radius”, referred to as R_tip, of the airfoil 214. The tip radius R_tip is the radial distance from the longitudinal axis 112 to the outermost radial coordinate (typically the tip leading edge 238) of the airfoil 214. A point located at the tip leading edge 238 would be referred to as 100% of tip radius R_tip (or 1.0 R/R_tip), and a point at the longitudinal axis 112 would be referred to as 0% of tip radius R_tip (or 0.0 R/R_tip). In different embodiments, different hub radius ratios may be used. For example, each airfoil 214 defines a tip radius R_tip along the radial direction R from the longitudinal axis 112 to the outermost radial coordinate of the airfoil 214 (typically at the tip leading edge 238), and a 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 typically the hub radius divided by the tip radius. As an example, for an exemplary embodiment where an outer radius of the hub 216 (centered on the longitudinal axis 112 (FIG. 1) of the fan 152 (FIG. 1)) is located radially at approximately thirty percent (30%) of the tip radius, a value of 0.3 R/R_tip corresponds to a zero percent (0%) span location. As indicated above, an R/R_tip value of 0.0 corresponds to the longitudinal axis 112. Thus, it should be understood that different hub radius ratios used in connection with the airfoil 214 may result in different span coordinate values for different R/R_tip coordinate values corresponding to various features of the airfoil 214 according to the present disclosure.

The airfoil 214 forms an aerodynamic surface extending along the axial direction A between the leading edge 234 and the trailing edge 236. The airfoil 214 extends outward from the root 250 in the radial direction R. In the illustrated embodiment, the leading edge 234 includes an inboard portion 242 that extends outward in the radial direction R to a particular span location, a medial portion 244 that extends from the inboard portion 242 to a tip portion 246, and the tip portion 246 that extends from the medial portion 244 to the tip 228 and encompasses the tip 228 and the tip leading edge 238. As used herein, a “tip portion” of an airfoil is defined as a portion of the airfoil extending radially from a location of a forward-most axial point of the leading edge of the airfoil to the tip of the airfoil. For example, in exemplary embodiments, a forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at or greater than fifty percent (50%) of the tip radius, or a value of 0.5 R/R_tip, such that the tip portion 246 of the airfoil 214 extends from a radial value of 0.5 R/R_tip to the tip 228 of the airfoil 214. In exemplary embodiments, a forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at or greater than fifty-five percent (55%) of the tip radius, or a value of 0.55 R/R_tip, such that the tip portion 246 of the airfoil 214 extends from a radial value of 0.55R/R_tip to the tip 228 of the airfoil 214. In exemplary embodiments, a forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at or greater than fifty-eight percent (58%) of the tip radius, or a value of 0.58 R/R_tip, such that the tip portion 246 of the airfoil 214 extends from a radial value of 0.58R/R_tip to the tip 228 of the airfoil 214. In exemplary embodiments, a forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at or greater than sixty percent (60%) of the tip radius, or a value of 0.6 R/R_tip, such that the tip portion 246 of the airfoil 214 extends from a radial value of 0.6 R/R_tip to the tip 228 of the airfoil 214. In exemplary embodiments, a forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at or greater than forty percent (40%) of a span of the airfoil 214, such that the tip portion 246 of the airfoil 214 extends from a radial location of forty percent (40%) of the span of the airfoil 214 to the tip 228 of the airfoil 214. In exemplary embodiments, a forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at or greater than forty-five percent (45%) of a span of the airfoil 214, such that the tip portion 246 of the airfoil 214 extends from a radial location of forty-five percent (45%) of the span of the airfoil 214 to the tip 228 of the airfoil 214. In exemplary embodiments, a forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at or greater than forty-eight percent (48%) of a span of the airfoil 214, such that the tip portion 246 of the airfoil 214 extends from a radial location of forty-eight percent (48%) of the span of the airfoil 214 to the tip 228 of the airfoil 214. In exemplary embodiments, a forward-most axial point of the leading edge 234 of the airfoil 214 may be located radially at or greater than fifty percent (50%) of a span of the airfoil 214, such that the tip portion 246 of the airfoil 214 extends from a radial location of fifty percent (50%) of the span of the airfoil 214 to the tip 228 of the airfoil 214. In exemplary embodiments, a forward-most axial point of a leading edge of an airfoil may vary based on a pitch angle of the airfoil (e.g., for a variable pitch fan). Accordingly, in exemplary embodiments, a “tip portion” of an airfoil may be defined as a portion of the airfoil extending radially from a location of a forward-most axial point of the leading edge of the airfoil to the tip of the airfoil when the airfoil is at its design orientation (e.g., at an orientation representative of subsonic cruise flight speed or operation). For example, cruise is a phase of the flight that occurs when an aircraft levels to a set altitude after a climb and before it begins to descend. Thus, as used herein, cruise represents a continuous, high speed, and stable condition of flight for which an aircraft is intended to operate. This description is to distinguish cruise from certain conditions that are abnormal or transient, such as dive, in which the aircraft can reach high flight speeds, but the aircraft is not intended to experience for a substantial portion of the mission from takeoff to landing. Thus, a subsonic cruise flight speed may refer to subsonic operation at a flight Mach number at or above 0.4, or at or above 0.5. Thus, in exemplary embodiments, the tip portion 246 of the airfoil 214 is defined as a portion of the airfoil 214 extending radially from the location of a 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 at its design orientation (e.g., at an orientation representative of high subsonic cruise flight speed or operation). In the illustrated embodiment, the leading edge 234 of the inboard portion 242 sweeps forward in the axial direction A, and the leading edge 234 of the medial portion 244 begins sweeping aft in the axial direction A outboard of the inboard portion 242. An acoustically active spanwise portion of the airfoil 214 may be determined, for example, via a relationship between a source strength 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 an acoustic source strength distributed radially along the airfoil 214 by an acoustic Green's function or radiation efficiency (e.g., the ability of noise sources to propagate acoustic energy to surrounding media) along the airfoil 214. The radiation efficiency may be any known relation describing the effective strength of a noise source on the airfoil, fan or propeller blade to an observer location of interest, and may be dependent on the airfoil shape, size, flow conditions, combinations thereof, or the like. In some exemplary embodiments, the trailing edge 236 of the airfoil 214 is configured having a smooth, curved profile (e.g., without steps or abrupt axial sweep changes/transitions).

Each airfoil 214 extends from the root or proximal end 250 at the hub 216 to the tip leading edge 238 and includes a generally concave pressure side 252 joined to a generally convex suction side 254 at the leading edge 234 and the trailing edge 236. The airfoil 214 may be represented as an array or “stack” of individual airfoil sections arrayed along a spanwise stacking line 256 (e.g., in-and-out of the page as depicted in FIG. 3). 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. Also, for each individual airfoil section of the airfoil 214, a curve called the “mean camber line” or “meanline” 260 represents the locus of points lying halfway between the concave pressure side 252 and the convex suction side 254. Typically, the airfoil 214 would incorporate “twist”, 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 illustrated example, it will be understood that the airfoil 214 may incorporate “lean”, a shift in the circumferential direction 113 (FIG. 1), and “axial sweep”, a shift in the axial direction A.

As indicated above, each airfoil 214 extends radially outward along a span “S” from the root to the tip 228, and 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 it may vary over the span S, as shown. An airfoil section of the airfoil 214 has a meanline angle 262, which refers to the angle between the tangent to the meanline 260 and the longitudinal axis 112. The meanline angle 262 can be measured at any location along the meanline 260. The value of the meanline angle 262 is a function of both the curvature of the meanline 260 and the pitch angle of the airfoil 214. Thus, the absolute value of the meanline angle 262 will change as the pitch angle of the airfoil 214 changes. However, it will be understood that the overall meanline shape characteristic of the meanline 260 is unchanging and depends solely on the curvature of the airfoil 214.

The airfoil 214 has a thickness 264 which is a distance measured normal to the meanline 260 between the concave pressure side 252 and the convex suction side 254, which can be measured at any location along the meanline 260. The thickness 264 may be described using a chord fraction, the value of which may be expressed as a percentage. For reference purposes, a relevant thickness “T” is measured at a distance “X” aft of the leading edge 234 where the distance “X” is represented or defined as a percentage or fraction of the total chord length, referred to herein as a percentage or fraction of “chord location,” “chordwise location,” or “chord fraction.” For example, a “0.5” chord fraction” represents a location aft of the leading edge 234 equal to 50% of the total chord length. Thus, as used herein, the chord fraction refers to a chordwise distance of the location from leading edge 234 to a point of interest divided by the chord C. So, for example, the leading edge 234 is located at 0.0 chord fraction, and the trailing edge 236 is located at 1.0 chord fraction. Further, a thickness ratio may be represented as the absolute value of the thickness T divided by a maximum thickness (T_MAX) of a particular airfoil section.

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

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

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

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

Further, as indicated above, the airfoil 214 as depicted and described herein may be configured for use as a guide vane 162 (FIG. 1).

Referring to FIG. 4, FIG. 4 is a schematic view of an exemplary airfoil 272 of an unducted airfoil assembly 270 according to an embodiment of the present disclosure. In some embodiments, airfoil 272 may be configured similarly to airfoil 214 (FIGS. 2 and 3). In FIG. 4, the airfoil 272 is viewed from the aft direction looking in the forward direction. The exemplary unducted airfoil assembly 270 may be configured for use as the fan 152 or the fan guide vane array 160 of the engine 100 as depicted in FIG. 1. The unducted airfoil assembly 270 includes an array of the airfoils 272 (only one shown in FIG. 4) that are regularly spaced apart circumferentially (e.g., in the circumferential direction 113) around a disk or hub of a rotor centered on the longitudinal axis 112 (FIG. 1) of the fan 152 (FIG. 1). Each airfoil 272 includes a root or proximal end 274 (i.e., an inboard end in the radial direction R toward the longitudinal axis 112 (FIG. 1)) and a tip portion 275 defining a tip 276 and a tip leading edge 278 (defined at an intersection of the tip 276 with a leading edge 280) such that a span or spanwise direction of the airfoil 272 is defined between the root 274 and the tip 276. As indicated above, the tip portion 275 is defined as a portion of the airfoil 272 extending radially from a forward-most axial point 292 of the leading edge 280 of the airfoil 272 to the radial location of the tip leading edge 278 of the airfoil 272. The airfoil 272 forms an aerodynamic surface extending along the axial direction between a 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 pitchable about its central airfoil axis 284.

In the illustrated embodiment, the airfoil 272 includes a pressure side 286 and a circumferentially or laterally opposite suction side 288. The pressure side 286 is generally concave and precedes the generally convex suction side 288 as the airfoil 272 rotates in the rotational direction 290. In one aspect of the present disclosure, the airfoil 272 includes certain geometries having specific circumferential lean and axial sweep features for the leading edge 280, the trailing edge 282, and the tip leading edge 278. In exemplary embodiments, the airfoil 272 includes certain geometries having specific circumferential lean and axial sweep features for the leading edge 280, the trailing edge 282, and the tip leading edge 278 when the airfoil 272 is positioned at its design orientation (e.g., as in a variable pitch fan with the airfoil 272 positioned at an orientation representative of subsonic cruise operation). For example, in the embodiment illustrated in FIG. 4, the airfoil 272 includes the forward-most axial point 292 on the leading edge 280. In the illustrated embodiment, a circumferential coordinate of the tip leading edge 278 is located in a direction opposite a direction of rotation of the airfoil 272 (e.g., a direction opposite the rotational direction 290) with respect to a circumferential coordinate of the forward-most axial point 292. Additionally, as illustrated in FIG. 4, circumferential coordinates of the leading edge 280 and the trailing edge 282 of the tip portion 275 lean in a direction opposite a direction of rotation of the airfoil 272 (e.g., a direction opposite the rotational direction 290). In other words, the tip portion 275 leans toward the suction side 288 of the airfoil 272. Thus, in exemplary embodiments, the entire tip portion 275 leans in a direction opposite the rotational direction 290. Thus, in exemplary embodiments, the tip leading edge 278 is circumferentially offset in a direction opposite the rotational direction 290 relative to a circumferential coordinate of the forward-most axial point 292.

Further, in some embodiments of the present disclosure, circumferential coordinates of the leading edge 280 in a radial direction monotonically increase relative to a circumferential coordinate of the forward-most axial point 292 of the leading edge 280 in a direction away from a rotation of a rotor assembly (e.g., rotor assembly 150 (FIG. 1)) containing the airfoil 272 (e.g., in a direction away or opposite the rotational direction 290) beyond certain R/R_tip values. For example, in some embodiments, circumferential coordinates of the leading edge 280 monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point 292 in a direction away from a rotation of a rotor assembly (e.g., rotor assembly 150 (FIG. 1)) containing the airfoil 272 (e.g., in a direction away or opposite the rotational direction 290) beyond a R/R_tip value of 0.6. In some embodiments, circumferential coordinates of the leading edge 280 monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point 292 in a direction away from a rotation of a rotor assembly (e.g., rotor assembly 150 (FIG. 1)) containing the airfoil 272 (e.g., in a direction away or opposite the rotational direction 290) beyond a R/R_tip value of 0.65. In some embodiments, circumferential coordinates of the leading edge 280 monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point 292 in a direction away from a rotation of a rotor assembly (e.g., rotor assembly 150 (FIG. 1)) containing the airfoil 272 (e.g., in a direction away or opposite the rotational direction 290) beyond a R/R_tip value of 0.68. In some embodiments, circumferential coordinates of the leading edge 280 monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point 292 in a direction away from a rotation of a rotor assembly (e.g., rotor assembly 150 (FIG. 1)) containing the airfoil 272 (e.g., in a direction away or opposite the rotational direction 290) beyond a R/R_tip value of 0.72. It should also be understood that in exemplary embodiments where the airfoil 272 comprises a guide vane (e.g., the fan guide vane 162), the lean of the airfoil 272 will be in the direction of rotation.

In some embodiments, relative to the circumferential coordinate of the forward-most axial point 292 of the leading edge 280, circumferential coordinates of a first sub-portion 294 of the leading edge 280 in the tip portion 275 immediately outboard of the forward-most axial point 292 lean in the direction of rotation of a rotor assembly (e.g., rotor assembly 150 (FIG. 1)) containing the airfoil 272, and for a second sub-portion 296 of the leading edge 280 in the tip portion 275 immediately outboard of the first sub-portion 294 extending to the tip 276, circumferential coordinates of the leading edge 280 monotonically increase, in a radial direction, in a direction away from a rotation of a rotor assembly (e.g., rotor assembly 150 (FIG. 1)) containing the airfoil 272 (e.g., in a direction away or opposite the rotational direction 290). In some embodiments, the first sub-portion 294 comprises less than twenty-five percent (25%) of the tip portion 275. In some embodiments, the first sub-portion 294 comprises less than fifteen percent (15%) of the tip portion 275. In some embodiments, the first sub-portion 294 comprises less than ten percent (10%) of the tip portion 275.

FIG. 5 is a graph 300 illustrating a thickness ratio profile of an airfoil section as a function of a chord fraction according to embodiments of the present disclosure. The graph 300 may be representative of the airfoil 214 (FIGS. 2 and 3) and the airfoil 272 (FIG. 4) as described herein. In the illustrated embodiment, the graph 300 illustrates a thickness ratio near the tip 228/276 (FIGS. 2-4) of the airfoil 214/272 (FIGS. 2-4). It should be understood that the thickness ratio may correspond to different airfoil sections near the tip of the airfoil (e.g., in at least a portion of the tip portions 246/275 of the airfoils 214/272 (FIG. 4)) or elsewhere. Graph 300 includes a horizontal axis 302 graduated in units of chord fraction and a vertical axis 304 expressed as thickness ratio (i.e., airfoil section thickness at a point of interest divided by a maximum thickness of the airfoil section). The airfoil section of the exemplary embodiment is designed for low noise and high efficiency (i.e., within airfoil 214 (FIGS. 2 and 3) or airfoil 272 (FIG. 4)) having a thickness ratio 306 substantially increased between 0.0 to 0.16 chord fraction. In one embodiment, a peak or maximum thickness 308 of the airfoil section of the airfoil 214/272 (FIGS. 2-4) in at least a portion of the tip portion 246/275 is at approximately 0.16 chord fraction. Thus, in this example, T_MAX of the airfoil 214/272 (FIGS. 2-4) is between 0.05 and 0.2 chord fraction. Further, in the illustrated 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.

In some embodiments, the thickness profile of the airfoil section (e.g., in at least a portion of the tip portion 246/275 of the airfoil 214/272 (FIG. 4)) remains substantially flat over a particular chord fraction range relative to a T_MAX chord fraction. For example, using the airfoil 272 (FIG. 4) as an example, in some embodiments, the thickness T does not decrease from T_MAX by more than ten percent (10%) over a chord fraction range extending from the T_MAX chord fraction location to a chord fraction located midway between either the leading edge 280 (FIG. 4) or trailing edge 282 (FIG. 4) of the airfoil 272 (FIG. 4). For example, in some embodiments, T_MAX is located at 0.16 chord fraction. Midway between the T_MAX chord fraction of 0.16 and the leading edge 280 (FIG. 4) is 0.08 chord fraction, and midway between the T_MAX chord fraction of 0.16 and the trailing edge 282 (FIG. 4) is 0.58 chord fraction. Thus, in this embodiment, the thickness T of the airfoil 272 (FIG. 4) does not decrease from T_MAX by more than ten percent (10%) between 0.08 and 0.58 chord fractions. Thus, in this embodiment, T/T_MAX is equal to or greater than 0.9 between chord fractions 0.08 and 0.58. Thus, in this embodiment, for a particular T_MAX chord fraction, T/T_MAX is equal to or greater than 0.90 over a chord fraction range extending from a chord fraction location midway between the leading edge 280 (FIG. 4) and the T_MAX chord fraction to a chord fraction location midway between the trailing edge 282 (FIG. 4) and the T_MAX chord fraction.

In some embodiments (e.g., in at least a portion of the tip portion 246/275 of the airfoil 214/272 (FIG. 4)). T/T_MAX is equal to or greater than 0.85 between chord fractions 0.08 and 0.58. For example, in this embodiment and as shown in FIG. 5, the thickness T of the airfoil 272 (FIG. 4) does not decrease from T_MAX by more than fifteen percent (15%) between 0.08 and 0.58 chord fractions. Accordingly, in some embodiments, for a particular T_MAX chord fraction, T/T_MAX is equal to or greater than 0.85 over a chord fraction range extending from a chord fraction midway between the leading edge 280 (FIG. 4) and the T_MAX chord fraction to a chord fraction midway between the trailing edge 282 (FIG. 4) and the T_MAX chord fraction. In some embodiments, the thickness T of the airfoil 272 (FIG. 4) does not decrease from T_MAX by more than twenty percent (20%) between 0.08 and 0.58 chord fractions. Accordingly, in this embodiment, for a particular T_MAX chord fraction, T/T_MAX is equal to or greater than 0.80 over a chord fraction range extending from a chord fraction midway between the leading edge 280 (FIG. 4) and the T_MAX chord fraction to a chord fraction midway between the trailing edge 282 (FIG. 4) and the T_MAX chord fraction.

Referring to FIG. 6, FIG. 6 is a schematic view of an exemplary airfoil 422 of an unducted airfoil assembly 420 according to another embodiment of the present disclosure. The airfoil 422 may be configured similarly to the airfoil 214 (FIGS. 2 and 3) and/or the airfoil 272 (FIG. 4). The airfoil 422 may be configured for use as the fan 152 or the fan guide vane array 160 of the engine 100 as depicted in FIG. 1.

In FIG. 6, 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 (i.e., an inboard end in the radial direction R toward the longitudinal axis 112 (FIG. 1)) and a tip 425. Also, an intersection of the tip 425 and the leading edge 428 is defined as a tip leading edge 426 such that a 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 at its trailing edge 430 the sculpted trailing edge feature 436 (e.g., a wavy feature or plurality of features) configured to facilitate wake mixing to reduce interaction noise caused by the airfoil 422 wakes impinging on downstream stationary airfoils or stators (or stator vanes), as described in U.S. Pat. No. 8,083,487 B2 which is hereby incorporated by reference in its entirety. A baseline 434 trailing edge having a smooth profile is depicted to further illustrate the sculpted trailing edge feature 436. Alternatively or additionally, the sculpted trailing edge feature 436 may be applied on the guide vanes 162 of the engine 100 (FIG. 1) to reduce the broadband noise generated by the turbulence in the vane boundary layer convecting past its trailing edge.

Referring to FIG. 7, FIG. 7 is a schematic view of exemplary sections of an airfoil 500 taken at different radial locations of the airfoil 500 according to embodiments of the present disclosure. For example, in exemplary embodiments, exemplary sections of the airfoil 500 taken at different radial locations of the airfoil 500 depicted in FIG. 7 may correspond to the airfoil 500 being positioned at its design orientation (e.g., in a variable pitch fan with the airfoil 500 positioned at an orientation representative of high subsonic cruise flight speed or operation). The airfoil 500 may be configured similarly to the airfoil 214 (FIGS. 2 and 3), the airfoil 272 (FIG. 4), or the airfoil 422 (FIG. 6). In FIG. 7, a horizontal axis 502 represents an axial direction A (e.g., left-to-right in FIG. 7 representing the aft direction), a vertical axis 504 represents a circumferential direction, the direction of rotation of the airfoil 500 is represented by an arrow 506, and the intersection of the horizontal axis 502 with the vertical axis 504 represents a pitch change axis 508 of the airfoil 500.

In FIG. 7, an airfoil section 510 of the airfoil 500 is taken at a hub location (e.g., the hub 216 (FIG. 2) for the airfoil 214 (FIG. 2)) of the airfoil 500, an airfoil section 512 of the airfoil 500 is taken at a forward-most axial point 514 of the airfoil 500 (e.g., the forward-most axial point 240 (FIG. 2) for the airfoil 214 (FIG. 2), and an airfoil section 516 of the airfoil 500 is taken at a tip of the airfoil 500 (e.g., the tip 228 (FIG. 2) for the airfoil 214 (FIG. 2)). As depicted in FIG. 7, in exemplary embodiments, the position of a point on the airfoil section 516 at the tip of the airfoil 500 (e.g., the tip 228 (FIG. 2) for the airfoil 214 (FIG. 2)) at a twenty-five percent (25%) chord fraction on the meanline of the airfoil section 516 is such that the magnitude of its circumferential offset from the pitch change axis 508 is greater than the magnitude of the axial offset from the pitch change axis 508, and wherein the point is located axially aft of the pitch change axis 508. For example, in FIG. 7, a line 520 depicted at a forty-five degree (45°) angle relative to the horizontal axis 502 represents points of equal magnitude of circumferential offset and axial offset relative to the pitch change axis 508. Thus, in the portion of FIG. 7 above the horizontal axis 502 and to the right of the vertical axis 504, points above and to the left of the line 520 represent a greater magnitude of circumferential offset relative to the pitch change axis 508 than the magnitude of 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 at least by the airfoil section 512) lies axially forward and circumferentially away from the direction of rotation 506 relative to the pitch change axis 508 (e.g., in the portion of FIG. 7 above the horizontal axis 502 and to the left of the vertical axis 504).

Thus, embodiments of the present disclosure include circumferentially spaced airfoils or blades where the blades in the outer or tip portion are configured having a defined lean to minimize cruise flight condition pressure signatures. Additionally, in some embodiments, a thickness distribution in the outer or tip portion of the airfoil is configured to minimize wakes at landing and takeoff (LTO) flight conditions. Further, embodiments of the present disclosure provide greater mechanical stability while reducing noise radiated by the blade. For example, embodiments of the present disclosure have an increased leading edge thickness that improves incidence tolerance at off-design flight conditions, thereby reducing noise. Moreover, embodiments of the present disclosure move the thickness distribution forward without changing a maximum thickness value in regions where desired while maintaining blade weight and reducing noise.

As will be appreciated from the description herein, various embodiments of a gas turbine engine are provided. Certain of these embodiments may be an unducted, single rotor gas turbine engine, or a ducted turbofan engine. An example of a ducted turbofan engine can be found in U.S. patent application Ser. No. 16/811,368 (Published as U.S. Patent Application Publication No. 2021/0108597), filed Mar. 6, 2020 (FIG. 10, Paragraph [0062], et al.; including an annular fan case 13 surrounding the airfoil blades 21 of rotating element 20 and surrounding vanes 31 of stationary element 30; and including a third stream/fan duct 73 (shown in FIG. 10, 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 of the exemplary gas turbine engine(s) discussed above with respect to the figures.

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

In one or more of these embodiments, a threshold power or disk loading for a fan (e.g., an unducted single rotor or primary forward fan) may range from 25 horsepower per square foot (hp/ft2) or greater at cruise altitude during a cruise operating mode. In particular embodiments of the engine, structures and methods provided herein generate power loading between 80 hp/ft2 and 160 hp/ft2 or higher at cruise altitude during a cruise operating mode, depending on whether the engine is an open rotor or ducted engine.

In various embodiments, an engine of the present disclosure is applied to a vehicle with a cruise altitude up to approximately 65,000 ft. In certain embodiments, cruise altitude is between approximately 28,000 ft and approximately 45,000 ft. In still certain embodiments, cruise altitude is expressed in flight levels based on a standard air pressure at sea level, in which a cruise flight condition is between FL280 and FL650. In another embodiment, cruise flight condition is between FL280 and FL450. In still certain embodiments, cruise altitude is defined based at least on a barometric pressure, in which cruise altitude is between approximately 4.85 psia and approximately 0.82 psia based on a sea level pressure of approximately 14.70 psia and sea level temperature at approximately 59 degrees fahrenheit. In another embodiment, cruise altitude is between approximately 4.85 psia and approximately 2.14 psia. It should be appreciated that in certain embodiments, the ranges of cruise altitude defined by pressure may be adjusted based on a different reference sea level pressure and/or sea level temperature.

As such, it will be appreciated that an engine of such a configuration may be configured to generate at least about 20,000 pounds and less than about 80,000 of thrust during operation at a rated speed, such as between about 20,000 and 50,000 pounds of thrust during operation at a rated speed, such as between about 25,000 and 40,000 pounds of thrust during operation at a rated speed.

In various exemplary embodiments, the fan may include twelve (12) fan blades. From a loading standpoint, such a blade count may allow a span of each blade to be reduced such that the overall diameter of the primary fan may also be reduced (e.g., to about twelve feet in one exemplary embodiment). That said, in other embodiments, the fan may have any suitable blade count and any suitable diameter. In certain suitable embodiments, the fan includes at least eight (8) blades. In another suitable embodiment, the fan may have at least twelve (12) blades. In yet another suitable embodiment, the fan may have at least fifteen (15) blades. In yet another suitable embodiment, the fan may 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.

Further, in certain exemplary embodiments, the rotor assembly may 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.

In various embodiments, it will be appreciated that the engine includes a ratio of a quantity of vanes to a quantity of blades that could be less than, equal to, or greater than 1:1. For example, in particular embodiments, the engine includes twelve (12) fan blades and ten (10) vanes. In other embodiments, the vane assembly includes a greater quantity of vanes to fan blades. For example, in particular embodiments, the engine includes ten (10) fan blades and twenty-three (23) vanes. For example, in certain embodiments, the engine may include a ratio of a quantity of blades to a quantity of vanes between 2:5 and 2:1, or between 2:4 and 3:2, or between 0.5 and 1.5. The ratio may be tuned based on a variety of factors including a size of the vanes to ensure a desired amount of swirl is removed for an airflow from the primary fan. In various embodiments, the quantity of blades is twenty (20) or fewer. In still certain embodiments, a sum of the quantity of blades and the quantity of vanes 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 a quantity of blades between eleven (11) and sixteen (16). In another embodiment, the engine includes twelve (12) blades and ten (10) vanes. In still 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 quantity of blades and vanes. In still yet another embodiment, the engine includes an equal quantity of blades and vanes, in which the quantity of blades is equal to or fewer than twenty (20). In various embodiments, the engine includes a combination of the quantity of blades to the quantity of vanes between 2:5 and 2:1, the difference between the quantity of blades and the quantity of vanes between two (2) and negative two (−2), and the quantity of blades between eleven (11) and sixteen (16). For example, a difference between the quantity of blades and the quantity 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.

Additionally, in certain exemplary embodiments, where the engine includes the third stream and a mid-fan (a ducted fan aft of the primary, forward fan), a ratio R1/R2 may be between about 1 and 10, or 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 mid-fan.

It should be appreciated that various embodiments of the engine, such as the single unducted rotor engine depicted and described herein, may allow for normal subsonic aircraft cruise altitude operation at or above Mach 0.5. In certain embodiments, the engine allows for normal aircraft operation between Mach 0.55 and Mach 0.85 at cruise altitude. In still particular embodiments, the engine allows for normal aircraft operation between Mach 0.75 and Mach 0.85. In certain embodiments, the engine allows for rotor blade tip speeds at or less than 750 feet per second (fps). In other embodiments, the rotor blade tip speed at a cruise flight condition can be 650 to 900 fps, or 700 to 800 fps.

A fan pressure ratio (FPR) for the fan of the fan assembly can be 1.04 to 1.20, or in some embodiments 1.05 to 1.1, or in some embodiments less than 1.08, as measured across the fan blades at a cruise flight condition.

In order for the gas turbine engine to operate with a fan having the above characteristics to define the above FPR, a gear assembly may be provided to reduce a rotational speed of the fan assembly relative to a driving shaft (such as a low pressure shaft coupled to a low pressure turbine). In some embodiments, a 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 within a range of 4.1 to 14.0, within a range of 4.5 to 14.0, or within a range of 6.0 to 14.0. In certain embodiments, the gear ratio is within a range of 4.5 to 12 or within a range of 6.0 to 11.0. As such, in some embodiments, the fan can be configured to rotate at a rotational speed of 700 to 1500 rpm at a cruise flight condition, while the power turbine (e.g., the low-pressure turbine) is configured to rotate at a rotational speed of 2,500 to 15,000 rpm at a cruise flight condition. In particular embodiments, the fan can be configured to rotate at a rotational speed of 850 to 1,350 rpm at a cruise flight condition, while the power turbine is configured to rotate at a rotational speed of 5,000 to 10,000 rpm at a cruise flight condition.

With respect to a turbomachine of the gas turbine engine, the compressors and/or turbines can include various stage counts. As disclosed herein, the stage count includes the number of rotors or blade stages in a particular component (e.g., a compressor or turbine). For example, in some embodiments, a low pressure compressor may include 1 to 8 stages, a high-pressure compressor may include 8 to 15 stages, a high-pressure turbine may include 1 to 2 stages, and/or a 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, an engine may include a one stage low pressure compressor, an 11 stage high pressure compressor, a two stage high pressure turbine, and 4 stages, or between 4 and 7 stages for the LPT. As another example, an engine can 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.

A core engine is generally encased in an outer casing defining one half of a core diameter (Dcore), which may be thought of as the maximum extent from a centerline axis (datum for R). In certain embodiments, the engine includes a length (L) from a longitudinally (or axial) forward end to a longitudinally aft end. In various embodiments, the engine defines a ratio of L/Dcore that provides for reduced installed drag. In one embodiment, L/Dcore is at least 2. In another embodiment, L/Dcore is at least 2.5. In some embodiments, the L/Dcore is less than 5, less than 4, or less than 3. In various embodiments, it should be appreciated that the L/Dcore is for a single unducted rotor engine.

The reduced installed drag may further provide for improved efficiency, such as improved specific fuel consumption. Additionally, or alternatively, the reduced drag may provide for cruise altitude engine and aircraft operation at the above describe Mach numbers at cruise altitude. Still particular embodiments may provide such benefits with reduced interaction noise between the blade assembly and the vane assembly and/or decreased overall noise generated by the engine by virtue of structures located in an annular duct of the engine.

Additionally, it should be appreciated that ranges of power loading and/or rotor blade tip speed may correspond to certain structures, core sizes, thrust outputs, etc., or other structures at the core engine. However, as previously stated, to the extent one or more structures provided herein may be known in the art, it should be appreciated that the present disclosure may include combinations of structures not previously known to combine, at least for reasons based in part on conflicting benefits versus losses, desired modes of operation, or other forms of teaching away in the art.

This written description uses examples to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be 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 with insubstantial differences from the literal languages of the claims.

While this disclosure has been described as having exemplary designs, the present disclosure can be further modified within the scope of this disclosure. This application is therefore intended to cover any variations, uses, or adaptations of the disclosure using its general principles. Further, this application is intended to cover such departures from the present disclosure as come within known or customary practice in the art to which this disclosure pertains and which fall within the limits of the appended claims.

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

An unducted airfoil assembly, comprising: an airfoil having spaced-apart pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced apart leading and trailing edges, and wherein the airfoil comprises a forward-most axial point; and wherein the airfoil is arranged around 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 the rotational direction relative to a circumferential location of the forward-most axial point.

The unducted airfoil assembly of the preceding clause, wherein the forward-most axial point is defined when the airfoil is oriented at a design orientation for subsonic cruise operation.

The unducted airfoil assembly of 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.

The unducted airfoil assembly of 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.

The unducted airfoil assembly of 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.

The unducted airfoil assembly of 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.

The unducted airfoil assembly of any preceding clause, wherein the forward-most axial point is radially located at or greater than forty percent of the span of the airfoil.

The unducted airfoil assembly of 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.

The unducted airfoil assembly of 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.

The unducted airfoil assembly of any preceding clause, wherein the forward-most axial point is radially located at or greater than fifty percent of the span of the airfoil.

The unducted airfoil assembly of 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 chordwise fractional distance is defined from the leading edge, and wherein the chordwise fractional distance to the maximum thickness is minimum in a tip portion of the airfoil.

The unducted airfoil assembly of any preceding clause, wherein the airfoil comprises a tip portion extending radially from the forward-most axial point to the tip, and wherein a maximum thickness of the airfoil in at least a portion of the tip portion for a chord extending from the leading edge to the trailing edge is located between five percent to thirty percent of the chord as measured from the leading edge.

The unducted airfoil assembly of any preceding clause wherein the trailing edge comprises a sculpted trailing edge feature.

The unducted airfoil assembly of any preceding clause, wherein the airfoil comprises a tip portion extending radially from the forward-most 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 of the airfoil section, and wherein the thickness ratio of any airfoil section in the tip portion is greater than 0.8 at a chord fraction between 0.05 and 0.16.

The unducted airfoil assembly of any preceding clause, wherein the airfoil comprises a tip portion extending radially from the forward-most axial point to the tip, and wherein a maximum thickness for at least a portion of the airfoil in the tip portion is located between 0.05 and 0.40 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least a portion of the airfoil in the tip portion is located between 0.05 and 0.30 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least a portion of the airfoil in the tip portion is located between 0.05 and 0.25 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least a portion of the airfoil in the tip portion is located between 0.05 and 0.20 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least a portion of the airfoil in the tip portion is located between the leading edge and 0.40 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least a portion of the airfoil in the tip portion is located between the leading edge and 0.30 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least a portion of the airfoil in the tip portion is located between the leading edge and 0.25 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least a portion of the airfoil in the tip portion is located between the leading edge and 0.20 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least twenty-five percent (25%) of a radial extent of the tip portion of the airfoil is located forward of 0.40 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least fifty percent (50%) of a radial extent of the tip portion of the airfoil is located forward of 0.40 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least sixty-seven percent (67%) of a radial extent of the tip portion of the airfoil is located forward of 0.40 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least seventy-five percent (75%) of a radial extent of the tip portion of the airfoil is located forward of 0.40 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least twenty-five percent (25%) of a radial extent of the tip portion of the airfoil is located forward of 0.30 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least fifty percent (50%) of a radial extent of the tip portion of the airfoil is located forward of 0.30 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein, a maximum thickness for at least sixty-seven percent (67%) of a radial extent of the tip portion of the airfoil is located forward of 0.30 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein, a maximum thickness for at least seventy-five percent (75%) of a radial extent of the tip portion of the airfoil is located forward of 0.30 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least twenty-five percent (25%) of a radial extent of the tip portion of the airfoil is located forward of 0.25 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least fifty percent (50%) of a radial extent of the tip portion of the airfoil is located forward of 0.25 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least sixty-seven percent (67%) of a radial extent of the tip portion of the airfoil is located forward of 0.25 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a maximum thickness for at least seventy-five percent (75%) of a radial extent of the tip portion of the airfoil is located forward of 0.25 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein the airfoil is a guide vane.

The unducted airfoil assembly of any preceding clause, wherein the airfoil is arranged around a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein the leading edge in the tip portion leans in a direction opposite the rotational direction.

The unducted airfoil assembly of any preceding clause, wherein a chordwise fractional distance from the leading edge of a maximum thickness of an airfoil section of the airfoil is minimum in a tip portion of the airfoil.

The unducted airfoil assembly of any preceding clause, wherein a chordwise fractional location of a maximum thickness of the airfoil is located furthest forward as measured from the leading edge in the tip portion.

The unducted airfoil assembly of any preceding clause, wherein the airfoil is arranged around a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein the airfoil includes a tip portion, and wherein the leading edge in the tip portion leans in a direction opposite the rotational direction.

The unducted airfoil assembly of any preceding clause, wherein the airfoil is arranged around a longitudinal axis and rotates about the longitudinal axis in a rotational direction, and wherein the airfoil includes a tip portion, and wherein a circumferential coordinate of the leading edge in the tip portion leans in a direction opposite the rotational direction.

The unducted airfoil assembly of any preceding clause, wherein circumferential coordinates of the leading edge monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point of the leading edge in a direction away from a rotational direction of the blade beyond a R/Rtip value of 0.6.

The unducted airfoil assembly of any preceding clause, wherein circumferential coordinates of the leading edge monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point of the leading edge in a direction away from a rotational direction of the blade beyond a R/Rtip value of 0.65.

The unducted airfoil assembly of any preceding clause, wherein circumferential coordinates of the leading edge monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point of the leading edge in a direction away from a rotational direction of the blade beyond a R/Rtip value of 0.68.

The unducted airfoil assembly of any preceding clause, wherein circumferential coordinates of the leading edge monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point of the leading edge in a direction away from a rotational direction of the blade beyond a R/Rtip value of 0.72.

The unducted airfoil assembly of any preceding clause, wherein circumferential coordinates of the leading edge monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point of the leading edge in a direction away from a rotational direction of the airfoil for at least a sub-portion of the tip portion beyond a tip radius value of the forward-most axial point.

The unducted airfoil assembly of any preceding clause, wherein circumferential coordinates of a first sub-portion of the leading edge in a tip portion of the airfoil immediately outboard of the forward-most axial point lean in the rotational direction of the airfoil, and for a second sub-portion of the leading edge in the tip portion immediately outboard of the first sub-portion extending to the tip, circumferential coordinates of the leading edge monotonically increase, in a radial direction, in a direction away from the rotational direction.

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

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

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

The unducted airfoil assembly of any preceding clause, wherein the airfoil comprises a tip portion, and wherein a thickness ratio of an airfoil section is defined as a thickness of the airfoil section divided by a maximum thickness of the airfoil section, and wherein the thickness ratio of the tip portion is greater than 0.8 at 0.16 chord fraction.

The unducted airfoil assembly of any preceding clause, wherein the maximum thickness of an airfoil section of the airfoil is between 0.05 and 0.2 chord fraction as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein the thickness ratio is equal to or greater than 0.8 at a chord fraction of 0.05 as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein the thickness ratio is equal to or greater than 0.8 at a chord fraction between 0.05 and 0.16 as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein the thickness ratio is equal to or greater than 0.8 at a chord fraction between 0.05 and 0.15 as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein the thickness ratio is equal to or greater than 0.8 at a chord fraction between 0.05 and 0.10 as measured from the leading edge.

An unducted airfoil assembly, comprising: an airfoil having a root, a medial portion, and a tip portion, the airfoil having spaced-apart pressure and suction sides extending radially in span from the root to a tip defined in the tip portion, and extending axially in chord between spaced apart leading and trailing edges; and wherein a thickness of the airfoil is defined as a distance measured between the pressure side and the suction side, and wherein a thickness ratio is defined as the thickness of an airfoil section divided by a maximum thickness at the airfoil 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 location midway between the leading edge and a chord fraction location of the maximum thickness to a chord fraction location midway between the trailing edge and the chord fraction location of the maximum thickness.

The unducted airfoil assembly of 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 chord fraction range extending from the chord fraction location of the maximum thickness to a chord fraction located midway between the leading edge or the trailing edge.

The unducted airfoil assembly of 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 as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein the thickness ratio is equal to or greater than 0.9 between chord fractions 0.08 and 0.58 as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein the thickness ratio in the tip portion is equal to or greater than 0.90 over a chord fraction range extending from a chord fraction location midway between the leading edge and a chord fraction location of the maximum thickness to a chord fraction location midway between the trailing edge and the chord fraction location of the maximum thickness.

The unducted airfoil assembly of any preceding clause, wherein the thickness ratio in the tip portion is equal to or greater than 0.85 between chord fractions 0.08 and 0.58 as measured from the leading edge.

The unducted airfoil assembly of 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 as measured from the leading edge.

The unducted airfoil assembly of any preceding clause, wherein a position of a point on an airfoil section at the tip at a twenty-five percent (25%) chord fraction on a meanline of the airfoil section is such that a magnitude of a circumferential offset of the point from a pitch change axis of the airfoil is greater than a magnitude of an axial offset of the point from the pitch change axis, and wherein the point is located axially aft of the pitch change axis.

The unducted airfoil assembly of any preceding clause, wherein at least a portion of the airfoil in the tip portion lies axially forward and circumferentially away from the direction of rotation of the airfoil relative to a pitch change axis of the airfoil.

Claims

1-7. (canceled)

8. An unducted airfoil assembly, comprising: an airfoil having spaced-apart pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced apart leading and trailing edges, and wherein the airfoil comprises a forward-most axial point located between the root and the tip, and wherein the airfoil defines a tip portion extending radially from the forward-most axial point to the tip; andwherein a thickness of the airfoil is defined as a distance measured between the pressure side and the suction side, and wherein a maximum thickness for at least 50% of a radial extent of the tip portion for a chord extending from the leading edge to the trailing edge is located forward of 0.40 chord fraction.

9. The unducted airfoil assembly of claim 8, 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 leans in a direction opposite the rotational direction.

10. The unducted airfoil assembly of claim 8, 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 the thickness ratio of any airfoil section in the tip portion is 0.8 or greater at 0.05 chord fraction.

11. The unducted airfoil assembly of claim 8, wherein circumferential coordinates of the leading edge monotonically increase, in a radial direction, relative to a circumferential coordinate of the forward-most axial point of the leading edge in a direction away from a rotational direction of the airfoil for at least a sub-portion of the tip portion beyond a tip radius value of the forward-most axial point.

12. The unducted airfoil assembly of claim 8, wherein the trailing edge comprises a sculpted trailing edge feature.

13. The unducted airfoil assembly of claim 8, wherein the maximum thickness for at least a portion of the airfoil in the tip portion is located forward of a 0.25 chord fraction as measured from the leading edge.

14. The unducted airfoil assembly of claim 8, wherein the airfoil comprises a guide vane.

15-20. (canceled)