CN116557346A - Airfoil assemblies with different orientation stages - Google Patents

Abstract

An airfoil assembly for a gas turbine engine, comprising: an inner support structure configured to enclose a longitudinal axis of the gas turbine engine; an outer support structure configured to enclose a longitudinal axis of the gas turbine engine, the outer support structure enclosing an inner support structure; and a stage comprising a plurality of airfoils extending from the inner support structure toward the outer support structure, the plurality of airfoils comprising: a first airfoil defining a first sweep angle, a first axial position, and a first lean angle; and a second airfoil defining a second sweep angle, a second axial position, and a second pitch angle, wherein the second sweep angle is different from the first sweep angle, the second axial position is different from the first axial position, or a combination thereof.

Description

Airfoil assemblies with different orientation stages

PRIORITY INFORMATION

The present application claims priority from polish patent application number p.440316 filed on 7, 2, 2022.

Technical Field

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

Background

Gas turbine engines typically include a fan and a turbine arranged in flow communication with each other. Further, the turbines of gas turbine engines typically include a compressor section, a combustion section, a turbine section, and an exhaust section in series flow order. In operation, air is provided from the fan to the inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. The fuel is mixed with compressed air and combusted within the combustion section to provide combustion gases. The combustion gases are channeled from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then directed through the exhaust section, e.g., to the atmosphere.

The fan is driven by the turbine. The fan includes a plurality of circumferentially spaced apart fan blades extending radially outwardly from a rotor disk. The rotation of the fan blades creates an airflow through the inlet to the compressor section of the turbine, as well as an airflow over the turbine. For some engines, a plurality of outlet guide vanes are provided downstream of the fan for straightening the airflow from the fan to increase the amount of thrust generated by the fan, for example.

Improvements to the outlet guide vanes of gas turbine engines, as well as other airfoil assemblies within gas turbine engines, would be welcomed in the art.

Drawings

A full and enabling disclosure of the present invention, 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 ducted gas turbine engine in accordance with various embodiments of the present subject matter.

FIG. 2 is a schematic cross-sectional view of an exemplary ductless gas turbine engine in accordance with various embodiments of the present subject matter.

FIG. 3 is a perspective view of an exemplary airfoil assembly according to various embodiments of the present subject matter.

FIG. 4 is a schematic cross-sectional view of a stage of an airfoil assembly according to an exemplary aspect of the disclosure.

FIG. 5 is a schematic cross-sectional view of a stage of an airfoil assembly according to another exemplary aspect of the disclosure.

FIG. 6 is a schematic cross-sectional view of a stage of an airfoil assembly according to another exemplary aspect of the disclosure.

FIG. 7 is a schematic cross-sectional view of a stage of an airfoil assembly according to another exemplary aspect of the disclosure.

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

Detailed Description

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

The term "exemplary" is used herein to mean "serving as an example, instance, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Moreover, all embodiments described herein are to be considered as exemplary unless expressly stated otherwise.

As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another, and are not intended to represent the location or importance of the respective components.

The terms "forward" and "aft" refer to relative positions within a turbine, gas turbine engine, or carrier, and refer to their normal operational attitude. For example, for a gas turbine engine, reference is made to a location closer to the engine inlet and then to a location closer to the engine nozzle or exhaust.

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

Unless otherwise indicated herein, the terms "coupled," "fixed," "attached," and the like are intended to mean both direct coupling, fixed, or attached and indirect coupling, fixed, or attached via one or more intermediate components or features.

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

The term "at least one" in the context of, for example, "at least one of A, B and C" refers to any combination of a only, B only, C only, or A, B and C.

Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by one or more terms, such as "about," "approximately," and "substantially," are not to be limited to the precise value specified. In at least some cases, the approximating language may correspond to the precision of an instrument for measuring the value or the precision of a method or machine for constructing or manufacturing a part and/or system. For example, approximating language may refer to within 1%, 2%, 4%, 10%, 15%, or 20% of the balance. These approximate margins may apply to individual values, to any one or both of the endpoints of a defined numerical range, and/or to margins of a range between the endpoints.

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.

As used herein, "third stream" refers to a non-primary air stream that is capable of increasing fluid energy to produce a small portion of the total propulsion system thrust. The third stream may typically receive inlet air (air from a duct channel downstream of the main fan) rather than free-stream air (as with the main fan). The pressure ratio of the third stream may be higher than the pressure ratio of the main propulsion stream (e.g., bypass or propeller driven propulsion stream). Thrust may be generated by a dedicated nozzle or by mixing the air flow through the third stream with the main thrust stream or core air flow, for example into a common nozzle.

In certain exemplary embodiments, the operating temperature of the airflow through the third stream may be below the maximum compressor discharge temperature of the engine, and more particularly, may be below 350 degrees Fahrenheit (e.g., below 300 degrees Fahrenheit, such as below 250 degrees Fahrenheit, such as below 200 degrees Fahrenheit, and at least as high as ambient temperature). In certain exemplary embodiments, these operating temperatures may facilitate heat transfer through or from the third stream and the separate fluid stream to the gas stream. Moreover, in certain exemplary embodiments, the airflow through the third flow may contribute less than 50% of the total engine thrust (and at least, for example, 2% of the total engine thrust) under takeoff conditions, or more particularly, operating at rated takeoff power at sea level, static flight speeds, 86 degrees Fahrenheit ambient temperature operating conditions.

Moreover, in certain exemplary embodiments, aspects of the airflow through the third flow (e.g., airflow, mixing, or exhaust characteristics) and thus the above-described exemplary percentage contribution to the total thrust force may be passively adjusted during engine operation or purposefully modified through the use of engine control features (e.g., fuel flow, motor power, variable stators, variable intake guide vanes, valves, variable exhaust geometry, or fluid features) to adjust or optimize overall system performance over a wide range of potential operating conditions.

The term "turbine" or "turbomachine" refers to a machine that includes one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together produce a torque output.

The term "gas turbine engine" refers to an engine having a turbine as all or part of its power source. Exemplary gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and the like, as well as hybrid electric versions of one or more of these engines.

The term "combustion section" refers to any heat addition system for a turbine. For example, the term combustion section may refer to a section that includes one or more of a deflagration-type combustion assembly, a rotary detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assembly. In certain exemplary embodiments, the combustion section may include an annular combustor, a can-type combustor, a tubular combustor, a Trapped Vortex Combustor (TVC), or other suitable combustion system, or a combination thereof.

Unless otherwise indicated, the terms "low", "high" or their respective degrees of comparison (e.g., lower, higher, if applicable) each refer to relative speeds within the engine when used with a compressor, turbine, shaft or spool piece, etc. For example, a "low turbine" or "low speed turbine" defines a component configured to operate at a lower rotational speed (e.g., a maximum allowable rotational speed) than a "high turbine" or "high speed turbine" at the engine.

In certain aspects of the present disclosure, an airfoil assembly for a turbine is provided. The airfoil assembly generally includes an inner support structure and a plurality of airfoils extending outwardly relative to the inner support structure along a radial direction (R). Typically, at least one airfoil of a stage is oriented in a different manner relative to one or more other airfoils of the stage, such as a different sweep angle, a different tilt angle, and/or a different axial position.

In certain exemplary embodiments, two or more airfoils of a stage may be differently oriented such that dynamic pressure distributions associated with the airfoils and/or rotor blades of the turbine interact in order to reduce sound distributions generated by or attenuated by the airfoil assembly (generated from an upstream source such as an upstream rotor assembly).

In certain additional or alternative exemplary embodiments, the airfoil assembly may further include an outer support structure. For example, the airfoil assembly may be used in a duct type gas turbine engine. In such embodiments, the airfoil may generally extend between the inner and outer support structures. With this configuration, at least one airfoil of a stage may be configured differently such that the airfoil array stiffness of the airfoil assembly falls within a predetermined range.

Referring now to the drawings, FIG. 1 illustrates a front cross-sectional view of an exemplary embodiment of a gas turbine engine that may incorporate one or more inventive aspects of the present disclosure. In particular, the exemplary gas turbine engine of FIG. 1 is configured as a single rotor ducted engine 10 that defines an axial direction A, a radial direction R, and a circumferential direction C (extending about the axial direction A; see, e.g., FIG. 3). As can be seen in fig. 1, the engine 10 takes the form of a closed rotor propulsion system and has a rotor assembly 12 (e.g., a fan assembly), the rotor assembly 12 including an array of airfoils disposed about a central longitudinal axis 14 of the engine 10, and more specifically, an array of rotor blades 16 disposed about the central longitudinal axis 14 of the engine 10. Furthermore, as will be explained in greater detail below, the engine 10 additionally includes a non-rotating airfoil assembly 18 (i.e., non-rotating relative to the longitudinal axis 14) positioned aft of the rotor assembly 12 that includes an array of airfoils 20 (e.g., outlet guide vanes) also disposed about the longitudinal axis 14.

The rotor blades 16 are disposed in a typically equally spaced relationship about the central longitudinal axis 14, and each has a root 22 and a tip 24 with a span defined therebetween. Similarly, the airfoils 20 are also arranged in a typically equally spaced relationship about the central longitudinal axis 14, and each has a root and tip and a span defined therebetween. The rotor assembly 12 further includes a hub 43 positioned forward of the plurality of rotor blades 16.

Further, engine 10 includes a turbine 30, and in the exemplary embodiment, turbine 30 includes a low speed system and a high speed system. The high speed system of the turbine 30 generally includes a high speed compressor 34, a high speed turbine 36, and a high speed shaft 38 extending therebetween and connecting the high speed compressor 34 and the high speed turbine 36. The high speed compressor 34 (or at least rotating components thereof), the high speed turbine 36 (or at least rotating components thereof), and the high speed shaft 38 may be collectively referred to as a high speed spool 35 of the engine. Further, the combustion section 40 is located between the high speed compressor 34 and the high speed turbine 36. The combustion section 40 may include one or more arrangements for receiving a mixture of fuel and air and providing a flow of combustion gases through the high speed turbine 36 to drive the high speed spool 35.

The low speed system similarly includes a low speed turbine 42, a low speed compressor or booster 44, and a low speed shaft 46 extending between and connecting the low speed compressor 44 and the low speed turbine 42, 44. The low speed compressor 44 (or at least rotating components thereof), the low speed turbine 42 (or at least rotating components thereof), and the low speed shaft 46 may be collectively referred to as a low speed spool 45 of the engine.

Still referring to FIG. 1, the turbine 30 is typically enclosed in a shroud 48. Further, it should be appreciated that the shroud 48 at least partially defines the inlet 50 and the injection exhaust nozzle section 52 and includes an inlet flow path 54 extending between the inlet 50 and the injection exhaust nozzle section 52. For the illustrated embodiment, the inlet 50 is an annular or axisymmetric 360 degree inlet located between the rotor blade assembly 12 and the stationary or stationary outlet guide vane assembly 18, and provides a path for incoming atmospheric air to enter the core flow path 54 (and the compressors 44, 34, the combustion section 40, and the turbines 36, 42) inside the airfoil 20 along the radial direction R. Such a location may be advantageous for a number of reasons, including management of icing performance and protection of the inlet 50 from various objects and materials that may be encountered during operation. However, in other embodiments, the inlet 50 may be positioned at any other suitable location, such as, for example, aft of the airfoil assembly 18, arranged in a non-axisymmetric manner, etc.

As briefly described above, the engine 10 includes an airfoil assembly 18. The airfoil assembly 18 extends from the shroud 48 and is positioned aft of the rotor assembly 12. The airfoils 20 of the airfoil assembly 18 may be mounted to a fixed frame or other mounting structure and not rotate relative to the longitudinal axis 14. For reference purposes, fig. 1 also depicts a forward direction with arrow F, which in turn defines the front and rear portions of the system. As shown in FIG. 1, the rotor assembly 12 is positioned forward of the turbine 30 in a "puller" configuration, and the exhaust ports 52 are positioned aft of the guide airfoils 20. As will be appreciated, the airfoils 20 of the airfoil assembly 18 may be configured to straighten the airflow from the rotor assembly 12 (e.g., reduce turbulence in the airflow) to increase the efficiency of the engine 10. For example, the size, shape, and configuration of the airfoils 20 may impart a counter-swirling vortex to the airflow from the rotor blade 16 such that the airflow has a greatly reduced degree of swirl in a downstream direction behind the two rows of airfoils (e.g., blades 16, airfoils 20), which may translate into an increased level of induction efficiency. Further, the airfoils 20 (e.g., outlet guide vanes) may support one or more of the turbine 30, the shroud 48, or the nacelle 80 relative to each other, a frame or other stationary structure supporting the engine 10, or an associated infrastructure, carrier, or the like.

Still referring to FIG. 1, it may be desirable for the rotor blade 16, the airfoil 20, or both to incorporate a pitch change mechanism such that the airfoils (e.g., blade 16, airfoil 20, etc.) may be rotated relative to the pitch rotation axis, either independently or in combination with each other. Such pitch variation may be used to vary thrust and/or eddy current effects under various operating conditions, including adjusting the magnitude or direction of thrust generated at the rotor blade 16, or providing thrust reversal features, which may be useful in certain operations, such as when an aircraft lands, or desirably modulating acoustic noise generated at least in part by the rotor blade 16, the airfoil 20, or aerodynamic interactions from the rotor blade 16 relative to the airfoil 20. More specifically, for the embodiment of FIG. 1, rotor assembly 12 is depicted as having pitch change mechanisms 58 for rotating rotor blades 16 about their respective pitch axes 60, and airfoil assembly 18 is depicted as having pitch change mechanisms 26 for rotating airfoils 20 about their respective pitch axes 64.

As depicted, the rotor assembly 12 is driven by the turbine 30, and more specifically, by the low speed spool 45. More specifically, engine 10 in the embodiment shown in FIG. 1 includes a power gearbox 56, and rotor assembly 12 is driven by low speed spool 45 of turbine 30 through power gearbox 56. The power gearbox 56 may include a gear set for reducing the rotational speed of the low speed spool 45 relative to the low speed turbine 42 such that the rotor assembly 12 may rotate at a slower rotational speed than the low speed spool 45. In this manner, the rotating rotor blades 16 of the rotor assembly 12 may rotate about the longitudinal axis 14 and generate thrust to propel the engine 10, and thus the aircraft associated therewith, in the forward direction F. As further shown in FIG. 1, the exemplary engine 10 includes a nacelle 80, with the nacelle 80 at least partially surrounding the rotor assembly 12 and the turbine 30, defining a bypass airflow passage 82 therebetween.

During operation of the ducted engine 10, a quantity of air 59 enters the engine 10 through the nacelle 80 and/or the associated inlet 51 of the rotor assembly 12. As a quantity of air 59 passes through rotor blade 16, a first portion of the quantity of air 59, as indicated by arrow 62, is directed or directed into bypass airflow passage 82, and a second portion of the quantity of air 59, as indicated by arrow 65, is directed or directed into core flow path 54, or more specifically, into low speed compressor 44. The ratio between the first portion of air 62 and the second portion of air 65 is commonly referred to as the bypass ratio. The pressure of the second portion of air 65 then increases as it passes through the high speed compressor 34 and into the combustion section 40 where it mixes with fuel and combusts to provide combustion gases 66.

The combustion gases 66 are channeled through high-speed turbine 36 wherein a portion of the thermal and/or kinetic energy from combustion gases 66 is extracted via the high-speed turbine stator airfoils and sequential stages of high-speed turbine rotor blades coupled to high-speed spool 35, causing high-speed spool 35 to rotate, thereby supporting operation of high-speed compressor 34. The combustion gases 66 are then channeled through low speed turbine 42 wherein a second portion of the thermal and kinetic energy is extracted from combustion gases 66 via the low pressure turbine stator airfoils and the sequential stages of low speed turbine rotor blades coupled to low speed spool 45, rotating low speed spool 45, thereby supporting operation of low speed compressor 44 and/or rotation of rotor assembly 12.

The combustion gases 66 are then channeled through injection exhaust nozzle section 52 of turbine 30 to provide propulsion thrust.

At the same time, as the first portion of air 62 is channeled through bypass airflow passage 82 prior to being discharged from nozzle exhaust section 76 of ducted engine 10, the pressure of first portion of air 62 increases substantially, also providing propulsion thrust. The high speed turbine 36, the low speed turbine 42, and the injection exhaust nozzle section 52 at least partially define a hot gas path 55 for directing the combustion gases 66 through the turbine 30.

However, it should be appreciated that the exemplary single rotor ducted engine depicted in FIG. 1 is by way of example only, and that in other exemplary embodiments, engine 10 may have any other suitable configuration, including, for example, any other suitable number of shafts or spools, turbines, compressors, etc. Additionally or alternatively, in other exemplary embodiments, any other suitable gas turbine engine may be provided. For example, in other exemplary embodiments, the gas turbine engine may be a ductless engine, a turbofan engine, a turboshaft engine, a turboprop engine, a turbojet engine, or the like.

For example, referring now to FIG. 2, a front cross-sectional view of another exemplary embodiment of a gas turbine engine 10 is disclosed that may incorporate one or more inventive aspects of the present disclosure. In particular, the exemplary gas turbine engine 10 of FIG. 2 is configured as a single rotor ductless engine. The exemplary embodiment of FIG. 2 may be configured in substantially the same manner as the exemplary engine 10 described above with respect to FIG. 1, and the same or similar reference numbers may refer to the same or similar components. For example, in the embodiment depicted in FIG. 2, engine 10 includes a rotor assembly 12, rotor assembly 12 including an array of airfoils disposed about a central longitudinal axis 14 of engine 10, and more specifically an array of rotor blades 16 disposed about central longitudinal axis 14 of engine 10. The exemplary engine 10 additionally includes a non-rotating airfoil assembly 18 (i.e., non-rotating relative to the longitudinal axis 14) positioned aft of the rotor assembly 12 that includes an array of airfoils also disposed about the longitudinal axis 14, and more particularly includes an array of airfoils 20 disposed about the longitudinal axis 14.

However, as will be appreciated, for the open rotor propulsion system embodiment shown in FIG. 2, the engine 10 includes an array of rotor blades 16 and an array of airfoils 20, which are ductless or open. Further to the exemplary embodiment of FIG. 2, engine 10 includes a third flow 31 extending from core flow path 54 to a location external to shroud 48. The third stream 31 may receive air from a core flow path 54 within the turbine 30 downstream of at least one compression stage (e.g., downstream of the first stage of the low speed compressor 44) and upstream of the high speed compressor 34. In additional or alternative embodiments, at least one array of rotor blades 16 or one array of airfoils 20 may be ducted, with at least one airfoil 20, array of rotor blades 16 in an open arrangement.

As will be appreciated, the gas turbine engine 10 of the present disclosure includes a variety of airfoil assemblies including, for example, a fan (e.g., rotor assembly 12 with fan blades 16) and an outlet guide vane assembly (e.g., airfoil assembly 18 with airfoil 20/outlet guide vane). It will be further appreciated that the gas turbine engine 10 of the present disclosure includes airfoil assemblies within the turbine 30 of the gas turbine engine 10, such as stages of inlet guide vanes (e.g., downstream of the inlet 50 and upstream of the low speed compressor), stages of stator vanes (e.g., within the low speed compressor 44), stages of rotor blades (e.g., within the low speed compressor 44), and the like.

For example, referring now to FIG. 3, a perspective view of an exemplary stage 100 of the airfoil assembly of the present disclosure is provided. More specifically, the airfoil assembly of FIG. 3 is configured in a similar manner as the exemplary airfoil assembly 18 of FIG. 1. The exemplary stage 100 will be generally described with respect to a ducted gas turbine engine, but the following disclosure may be equally applicable to any other gas turbine engine, such as ductless gas turbine engines, turbofan engines, open propeller engines, turbojet engines, power generating gas turbine engines, and the like.

The stage 100 of the airfoil assembly 18 is generally provided with an inner support structure 102 (e.g., a first circular frame member or fan hub frame), an outer support structure 104 (e.g., a circular frame member or fan casing), and a plurality of airfoils 20 (e.g., outlet guide vanes or counter-rotating airfoils), the plurality of airfoils 20 being disposed in a circumferential array about the inner support structure 102 or radially between the inner support structure 102 and the outer support structure 104. According to an exemplary embodiment, the airfoil assembly 18 shares a longitudinal axis with the longitudinal axis 14 of the engine 10 (FIG. 1). Accordingly, the inner support structure 102 may be configured to enclose the longitudinal axis 14 of the engine 10. It should be appreciated that the inner support structure 102 may be configured to be fixed relative to or by a fixed support, frame, etc. of the pipeline engine 10. The outer support structure 104 may be configured to enclose the longitudinal axis 14 of the pipeline engine 10 or the inner support structure 102. In one embodiment, the outer support structure 104 may be configured to be fixed relative to or by a fixed support, frame, etc. of the pipeline engine 10. For example, the outer support structure 104 may be supported and/or fixed at least in part via the plurality of airfoils 20 of the stage 100 of the airfoil assembly 18.

In this manner, it should be appreciated that the airfoil assembly 18 may be used to mount the outlet guide vanes into a ducted gas turbine engine. In this regard, the airfoil 20 may provide a load path within the stage 100 from an inner support structure 102 (e.g., a fan hub frame (and thus a turbine)) and an outer support structure 104 (e.g., the nacelle 80 (see FIG. 1) or a component coupled or formed with the nacelle 80).

In alternative embodiments, the stages of the airfoil assembly may be configured for a ductless gas turbine engine, such as an open rotor gas turbine engine (see FIG. 2). Such a stage for a ductless gas turbine engine may be configured similar to stage 100 of airfoil assembly 18 described above with respect to FIG. 3. For example, an exemplary stage of an airfoil assembly configured for use with a ductless engine may generally include an inner support structure, such as a first circular frame member; a bonnet of the engine; a component supported, coupled to, or formed with the hood; and/or other support structures supported relative to a frame or other fixed structure of the ductless gas turbine engine. The exemplary stage of an airfoil assembly configured for use with a ductless engine also typically includes a plurality of airfoils 20, such as counter-rotating airfoils, disposed in a circumferential array about the inner support structure 102.

However, for various embodiments of such exemplary airfoil assemblies configured for use with ductless engines, there is no external support structure. In additional or alternative embodiments, a stage of an airfoil assembly configured for use with a ductless engine may include an outer support structure, such as a shroud, segmented shroud, or the like, extending between two or more airfoils at respective distal ends of the airfoils. In such embodiments, there may be no outer support structure extending between any two or more consecutive airfoils within the stage. Thus, it should be appreciated that one or more outer support structures may only partially enclose the plurality of airfoils and/or inner support structures. In this manner, it should be appreciated that the airfoil may be used to support an outer support structure relative to a ductless gas turbine engine.

In additional or alternative embodiments, the stages of the airfoil assembly may be configured for use in any suitable turbine or any suitable stator airfoil assembly of a gas turbine engine. It should be appreciated that the present disclosure is equally applicable to suitably configured stator stages of compressors, turbines, etc. of gas turbine engines.

Referring now to fig. 4-7, schematic cross-sectional views of a stage 100 of an exemplary airfoil assembly 18 are shown in accordance with aspects of the present disclosure. The exemplary stage 100 of fig. 4-7 may be configured in a similar manner as the exemplary stage 100 of fig. 3. In general, each of FIGS. 4-7 illustrates an embodiment in which one or more airfoils 20 of a stage 100 define at least one respective orientation that is different from the remaining airfoils 20 of the stage 100.

The embodiments of fig. 4-7 will be described and generally are configured as stages of an airfoil assembly suitable for incorporation within a ducted gas turbine engine (e.g., engine 10 described above with reference to fig. 1). Further, the embodiments of fig. 4-7 are described and illustrated as surrounding a longitudinal axis 14 of an associated gas turbine engine, such as the longitudinal axis 14 of the engine 10 of fig. 1 or 2. However, in other embodiments, it should be appreciated that the longitudinal axis 14 of the stage 100 of the airfoil assembly 18 may be configured differently than the longitudinal axis of the associated gas turbine engine. For example, the longitudinal axis of the stage 100 may be radially displaced and/or define an angle relative to the longitudinal axis of the associated turbine.

Further, for the depicted embodiment, the rotor blades 16 of the rotor assembly 12 are positioned axially forward and/or upstream of the stage 100 of the airfoil assembly 18, e.g., directly upstream of the airfoil assembly 18. In at least one embodiment, the rotor assembly 12 and the airfoil assembly 18 may define a single rotor-stator stage of the turbine. Typically, the rotor assembly 12 is powered by a rotating spool of a turbine (e.g., a low speed spool 45 or similarly configured rotating spool as described with reference to fig. 1 and 2). In the exemplary embodiment of fig. 4-7, rotor assembly 12 and rotor blades 16 may generally be configured as a fan assembly and fan blades, respectively, of a gas turbine engine, such as fan assembly 12 of ducted engine 10 of fig. 1 or a similarly configured ducted gas turbine engine. Additionally or alternatively, the illustrated rotor assembly 12 and rotor blades 16 may generally be configured as a rotor assembly and rotor blades, respectively, of a gas turbine engine, such as the ductless engine 10 of FIG. 2 or a similarly configured ductless gas turbine engine rotor assembly 12.

However, it should be understood that the present disclosure is equally applicable to any rotor-stator stage of a turbine. For example, the rotor blades 16 may include, but are not limited to, compressor rotor blades, turbine rotor blades, and the like. In such embodiments, the stages 100 of the airfoil assembly 18 may correspond to stator stages of a compressor (including compressor stator airfoils), a turbine (including turbine stator airfoils), and the like, respectively.

Thus, each of the exemplary embodiments of fig. 4-7 includes an inner support structure 102, the inner support structure 102 surrounding the respective longitudinal axis 14 of the stage 100. Each of the exemplary embodiments of fig. 4-7 also includes a plurality of airfoils 20 extending from the inner support structure 102. The airfoil 20 may be coupled, fixed, supported, etc. with respect to the inner support structure 102. The inner support structure 102 itself may be supported, coupled, and/or fixed relative to a frame, bracket, other fixed structure, etc. of the gas turbine engine. In one or more embodiments, it should be appreciated that the plurality of airfoils 20 are configured to remain stationary during operation of the associated gas turbine engine 10.

In particular, as shown in the illustrated embodiment, the respective stage 100 may be configured for use in a ducted gas turbine engine. Thus, each of the embodiments of fig. 4-7 is shown with an outer support structure 104 (depicted in phantom). In one embodiment, the outer support structure 104 may be configured to be fixed relative to or by a fixed support, structure, frame, nacelle, fan housing, hood, etc. of the gas turbine engine 10. Additionally or alternatively, each airfoil 20 within the exemplary stage 100 may extend from the inner support structure 102 to the outer support structure 104. Additionally or alternatively, one or more airfoils 20, such as all airfoils 20, may be coupled, fixed, or supported relative to the inner support structure 102, the outer support structure 104, or both. For example, in various embodiments, the exemplary stage 100 of fig. 4-7 may be configured for a gas turbine engine, such as a ducted gas turbine engine, configured the same or similar to the ducted engine 10 of fig. 1, such as a turbofan engine. In various embodiments, such as those configured for a ducted gas turbine engine and/or a turbofan engine, the plurality of airfoils 20 may be configured as a plurality of outlet guide vanes.

However, it should be understood that the following description applies equally to other suitably configured stages and associated airfoil assemblies. For example, in additional or alternative embodiments, each of the exemplary stages 100 depicted in fig. 4-7 may be configured for a ductless gas turbine engine configured the same or similar to the ductless engine 10 of fig. 2, such as an open rotor gas turbine engine. Thus, each of FIGS. 4-7 shows the outer support structure 104 in phantom, as the following description applies equally to a turbomachine that is ductless and/or does not include an outer support structure. In various such embodiments, the plurality of airfoils 20 of the stage 100 may be configured as a plurality of counter-rotating airfoils.

In additional or alternative embodiments, any of the example stages 100 depicted in fig. 4-7 may be configured for a ductless engine including an outer support structure 106 (depicted in phantom in the example embodiments of fig. 4-7). The outer support structure 106 may extend between respective distal ends 124 of two or more airfoils 20 of the stage 100 of the airfoil assembly 18. For example, in one embodiment, the outer support structure 106 may extend in the circumferential direction C between two airfoils of the stage 100. Additionally or alternatively, one or more such airfoils 20 associated with the outer support structure 106, such as all airfoils 20 surrounded or partially surrounded by the outer support structure 106, may be coupled to the outer support structure 106, fixed to the outer support structure 106, supporting the outer support structure 106, and so forth. It should also be appreciated that the outer support structure 106 may extend between the non-continuous airfoils 20 of the stage 100. Thus, the outer support structure 106 may at least partially enclose three or more airfoils 20. Such an outer support structure 106 may be coupled to one or more airfoils 20, fixed to one or more airfoils 20, supported by one or more airfoils 20, etc., one or more airfoils 20 being, for example, two airfoils, such as two airfoils defining a circumferential end of the outer support structure 106, or all airfoils 20 surrounded by the outer support structure 106.

In this manner, it will be appreciated that one or more airfoils 20 surrounded by the outer support structure 106 may be used to support the outer support structure 106 relative to an associated ductless gas turbine engine and/or an associated ductless gas turbine engine core (e.g., the turbine 30 of FIG. 2 or a similarly configured turbine). In additional or alternative embodiments, the outer support structure 106 extending between any two or more consecutive airfoils 20 within the stage 100 may be absent. Accordingly, it should be appreciated that the one or more outer support structures 106 may only partially surround the airfoil 20 and/or the inner support structure 102 within the stage 100. In one embodiment, the one or more outer support structures 106 may be configured as one or more shields, such as segmented shields.

Referring still generally to the exemplary embodiment of fig. 4-7, the plurality of airfoils 20 of the stage 100 may include at least a first airfoil 120, a second airfoil 220, and a third airfoil 320. However, it should be appreciated that the stage 100 may include any suitable number of airfoils 20, such as four or more, 10 or more, 20 or more, etc., and that only a portion of the airfoils 20 of the respective exemplary stage 100 are depicted in fig. 4-7. In several embodiments, the first airfoil 120, the second airfoil 220, and the third airfoil 320 may be continuous airfoils within the respective stages 100 of the airfoil assembly 18. However, in additional or alternative embodiments, one or more airfoils 20 may be positioned generally in the circumferential direction (C) between the first airfoil 120 and the second airfoil 220 and/or between the second airfoil 220 and the third airfoil 320 in the stage 100.

In general, each airfoil 20 of the depicted embodiment includes a root end 122 and a distal end 124 (only described with respect to the root end 122 and distal end 124 of the first airfoil 120 for clarity). Each airfoil 20 may include a body 126 (depicted relative to the body 126 of the first airfoil 120 only for clarity) extending between the root end 122 and the distal end 124 along the span of the respective airfoil 20. The body 126 of each airfoil 20 may generally define an airfoil profile. For example, as shown with respect to the first airfoil 120 of fig. 4-7, the body 126 of each airfoil 20 may include a leading edge 128, a trailing edge 130, a suction side 132 extending between the leading edge 128 and the trailing edge 130 and between the root end 122 and the distal end 124, and a pressure side 134 extending between the leading edge 128 and the trailing edge 130 and between the root end 122 and the distal end 124. Each airfoil 20 generally defines a chord extending between a leading edge 128 and a trailing edge 130. In some embodiments, the chords of one or more airfoils 20 of a stage 100 may be the same along the respective spans of the body 126. Additionally or alternatively, the direction of the chord may vary along the respective spans of one or more airfoils 20 (e.g., airfoils configured to have varying dihedral angles, sweeps, or twists along the respective spans of the airfoils).

According to various aspects of the present disclosure, the exemplary stage 100 of the airfoil assembly 18 may include two or more airfoils 120 defining at least one differential and/or different orientation relative to each other. For example, two or more airfoils 120 may define different sweep angles (fig. 4), two or more airfoils 120 may define different tilt angles (fig. 5), two or more airfoils 120 may define different axial positions (fig. 6), two or more airfoils 120 may define different ring spacing (fig. 7), or any combination of the foregoing. Each of these embodiments is described in more detail below.

Rotor assembly 12 generally provides and/or generates a downstream flow of a working fluid (e.g., into the atmosphere). Each rotor blade 16 produces a sound profile when rotated. Furthermore, such flow of the working fluid creates one or more different acoustic contours due to collisions and/or interactions of the working fluid with the corresponding airfoil 20 of the stage 100 of the airfoil assembly 18. The accumulation of these sound profiles often results in an undesirable, unacceptable or unnecessarily high level sound output from the associated turbine. However, various aspects of the present disclosure may mitigate undesirable noise levels from gas turbine engines by generally utilizing destructive interference phenomena. For example, two or more airfoils 20 of a stage 100 may be differently oriented such that the sound profiles associated with any two or more airfoils 20 or rotor blades 16 interact to reduce or eliminate the noise level generated by the associated stage 100 of the airfoil assembly 18, the associated rotor-stator stage, or the total sound output from the associated turbine.

In various embodiments, the first airfoil 120 and the second airfoil 220 of the stage 100 have at least one different orientation such that a phase shift is created between the sound contours produced by the first airfoil 120 and the second airfoil 220 during operation of the associated gas turbine engine 10. In additional or alternative embodiments, the first airfoil 120 and the second airfoil 220 of the stage 100 have at least one different orientation such that a phase shift is created between one or more of the first airfoil 120 and the second airfoil 220 and the additional airfoil 20 (e.g., the third airfoil 320) during operation of the associated gas turbine engine 10. In additional or alternative embodiments, the first airfoil 120 and the second airfoil 220 of the stage 100 have at least one different orientation such that a phase shift is created between the sound profile produced by one or both of the first airfoil 120 and the second airfoil 220 and the one or more rotating rotor blades 16.

It should be appreciated that such a phase shift may be determined with reference to one or more predetermined locations relative to stage 100 and/or an associated gas turbine engine. For example, the first airfoil 120 and the second airfoil 220 may have at least one different orientation such that destructive interference occurs between the sound profiles of any of the above-described airfoils and/or blades at the stage 100 of the airfoil assembly 18, at the inlet of the relevant gas turbine engine (e.g., the inlet of a ducted gas turbine engine, such as inlet 51 shown in fig. 1), at the exhaust outlet of the relevant gas turbine engine (e.g., the exhaust nozzle of a ducted gas turbine engine, such as jet exhaust nozzle section 52 or nozzle exhaust section 76 shown in fig. 1), or at a known location or distance relative to the first airfoil 120 and the second airfoil 220 (e.g., the passenger compartment of the relevant vehicle or portion thereof or the operator position of the relevant gas turbine engine).

In additional or alternative embodiments, the first airfoil 120 and the second airfoil 220 of the stage 100 have at least one different orientation such that the airfoil array stiffness of the airfoil assembly 18 falls within a predetermined range. For example, two or more differently oriented airfoils 20 (e.g., first airfoil 120, second airfoil 220, and/or third airfoil 320) may change the overall airfoil array stiffness or airfoil array stiffness at one or more discrete locations of the airfoil assembly 18. For example, the first airfoil 120 may define a first stiffness and the second airfoil 220 may define a second stiffness. The first stiffness may be at least about 5% greater than the second stiffness, such as at least about 10% greater, such as at least about 15% greater, such as up to 200% greater (calculated by (first stiffness-second stiffness)/first stiffness; wherein stiffness is measured in newtons per meter). The first stiffness and the second stiffness are measured in the same direction.

It should be appreciated that such differently oriented airfoils 20 may be used to adjust the overall stiffness relative to the axial direction a, the circumferential direction C, or a combination of both, and/or such local stiffness at one or more locations of the airfoil assembly 18.

Referring now in particular to FIG. 4, a close-up cross-sectional schematic view of a stage 100 of the airfoil assembly 18 is shown in accordance with an exemplary embodiment of the present disclosure. More specifically, fig. 4 is a schematic view of stage 100 taken at a point along circumferential direction C along a plane defined by radial direction R and axial direction a. As shown, the stage 100 may include two or more airfoils 20 defining discrete sweep angles (e.g., a first sweep angle 136 of the first airfoil 120, a second sweep angle 236 of the second airfoil 220, a third sweep angle 336 of the third airfoil 320, etc.). The sweep angle of the airfoils 20 of the stage 100 generally corresponds to the angle defined between the direction of the span of the respective airfoil 20 and the direction of the longitudinal axis 14 in the plane defined by the axial direction a and the radial direction R. For example, as shown with respect to the first airfoil 120, in a plane defined by the axial direction a and the radial direction R, a first sweep angle 136 is defined between the spanwise direction of the first airfoil 120 and the longitudinal axis 14.

It should be appreciated that each airfoil 20 of the stage 100 may define a discrete sweep angle. In additional or alternative embodiments, two or more airfoils 20 may define the same sweep angle (e.g., first sweep angle 136) while two or more airfoils 20 define the same sweep angle (e.g., second sweep angle 236). Additionally or alternatively, the stage 100 may include one or more mating, groups, or individual further airfoils 20 defining any number of discrete sweep angles as desired or required.

In the depicted embodiment, the first and second grazing angles 136, 236 differ by at least 5%, such as by at least 10%, such as by at least 15%, such as by at least 20%, such as by as much as 90% (calculated by (first grazing angle-second grazing angle)/first grazing angle).

Further, for the depicted embodiment, the first sweep angle 136 is less than the second sweep angle 236, and the second sweep angle 236 is less than the third sweep angle 336. The difference between the second sweep angle 236 and the third sweep angle 336 may be within a similar range as defined above for the first sweep angle 136 and the second sweep angle 236.

It should be appreciated that each airfoil 20 of the stage 100 may define a discrete sweep angle. In additional or alternative embodiments, two or more airfoils 20 may define a first common sweep angle (e.g., first sweep angle 136), while two or more other airfoils 20 define a different common sweep angle (e.g., second sweep angle 236). Additionally or alternatively, the stage 100 may include one or more pairs, groups, or individual further airfoils 20 defining any number of discrete sweep angles desired or required.

Referring now in particular to FIG. 5, a close-up cross-sectional schematic view of a stage 100 of the airfoil assembly 18 is shown in accordance with an alternative or additional exemplary embodiment of the present disclosure. More specifically, fig. 5 is a schematic view of a stage 100 taken at a point along the axial direction a along a plane defined by the radial direction R and the circumferential direction C. As shown, the stage 100 may include two or more airfoils 20 defining discrete pitch angles (e.g., a first pitch angle 138 of the first airfoil 120, a second pitch angle 238 of the second airfoil 220, a third pitch angle 338 of the third airfoil 320, etc.). The inclination angle of the airfoils 20 of the stage 100 generally corresponds to the angle defined between the spanwise and circumferential directions C of the respective airfoils 20 in a plane defined by the radial direction R and circumferential direction C. For example, as shown with respect to the first airfoil 120, in a plane defined by the radial direction R and the circumferential direction C, a first angle of inclination 138 is defined between the spanwise direction and the circumferential direction C of the first airfoil 120.

In the depicted embodiment, the first tilt angle 138 differs from the second tilt angle 238 by at least 5%, such as by at least 10%, such as by at least 15%, such as by at least 20%, such as by as much as 90% (calculated by (first tilt angle-second tilt angle)/first tilt angle).

Further, for the depicted embodiment, the first tilt angle 138 is less than the second tilt angle 238, and the second tilt angle 238 is less than the third tilt angle 338. The difference between the second tilt angle 238 and the third tilt angle 338 may be within a similar range as defined above for the first tilt angle 138 and the second tilt angle 238.

It should be appreciated that each airfoil 20 of a stage 100 may define a discrete angle of inclination. In additional or alternative embodiments, two or more airfoils 20 may define a first common tilt angle (e.g., first tilt angle 138), while two or more other airfoils 20 define a different common tilt angle (e.g., second tilt angle 238). Additionally or alternatively, the stage 100 may include one or more mating, group, or individual further airfoils 20 defining any number of discrete lean angles as desired or required.

Referring now in particular to FIG. 6, a close-up cross-sectional schematic view of a stage 100 of the airfoil assembly 18 is shown in accordance with an additional or alternative exemplary embodiment of the present disclosure. More specifically, fig. 6 is a schematic view of stage 100 taken at a point along circumferential direction C along a plane defined by radial direction R and axial direction a. As shown, the stage 100 may include two or more airfoils 20 positioned at discrete axial locations (e.g., a first axial location 140 of the first airfoil 120, a second axial location 240 of the second airfoil 220, a third axial location 340 of the third airfoil 320, etc.). The axial position of the airfoils 20 of the stage 100 generally corresponds to the position of the respective airfoils 20 along the longitudinal axis 14, for example in a depicted plane defined by the axial direction a and the radial direction R. More specifically, as used herein, axial position refers to a forward-most point of the airfoil, e.g., a forward-most point of the leading edge of the airfoil. For example, the first axial position 140 is the forward-most point of the first leading edge 128 of the first airfoil, the second axial position 240 is the forward-most point of the second leading edge 228 of the second airfoil 220, and the third axial position 340 is the forward-most point of the third leading edge 328 of the third airfoil 320.

In the depicted embodiment, the first axial position 140 is offset from the second axial position 240 by at least about 0.25 inches, such as at least about 0.75 inches, such as at least about 1.25 inches, such as at least about 2 inches, such as at least about 2.5 inches, such as at least about 4 inches, such as at most about 20 inches.

Further to the depicted embodiment, the first axial position 140 is forward of the second axial position 240, and the second axial position 240 is forward of the third axial position 340. The difference between the second axial position 240 and the third axial position 340 may be within a similar range as defined above for the first axial position 140 and the second axial position 240.

It should be appreciated that each airfoil 20 of a stage 100 may be positioned at a discrete axial location. In additional or alternative embodiments, two or more airfoils 20 may be positioned at a first common axial position (e.g., first axial position 140), while two or more airfoils 20 may be positioned at a second common axial position (e.g., second axial position 240). Additionally or alternatively, the stage 100 may include one or more mating, group, or individual additional airfoils 20 positioned at any number of discrete axial positions as desired or required.

Referring now in particular to FIG. 7, a close-up cross-sectional schematic view of a stage 100 of the airfoil assembly 18 is shown in accordance with an alternative or additional exemplary embodiment of the present disclosure. More specifically, fig. 7 is a schematic view of a stage 100 taken at a point along the axial direction a along a plane defined by the radial direction R and the circumferential direction C. As shown, the stage 100 may include two or more airfoils 20 defining discrete ring pitches (e.g., a first ring pitch between a first ring position 142 of a first airfoil 120 and a ring position of a next successive airfoil 20, a second ring pitch between a second ring position 242 of a second airfoil 220 and a ring position of a next successive airfoil 20, a third ring pitch between a third ring position 342 of a third airfoil 320 and a ring position of a next successive airfoil 20, etc.).

The ring spacing of the airfoils 20 of a stage 100 generally corresponds to the difference between the circumferential position of the respective airfoil 20 and the circumferential position of the next successive airfoil 20 in the circumferential direction C of the next successive airfoil 20. Such ring spacing may generally be defined as the difference between the angular position of the respective airfoil 20 and the angular position of the next successive airfoil 20. Additionally or alternatively, such ring spacing may be defined as the circumferential distance between the circumferential position of the respective airfoil 20 and the circumferential position of the next successive airfoil 20. It should be appreciated that such ring spacing may be determined at the same radial distance of the respective airfoil 20 and the next successive airfoil 20 relative to the longitudinal axis 14. Further, the ring spacing of each airfoil 20 of the stage 100 may be determined at the same radial distance relative to the longitudinal axis 14.

For example, as shown with respect to the first airfoil 120, the first ring spacing may be defined as the difference between the angular position of the first ring location 142 (e.g., at the root 122) and the angular position of the second ring location 142 (e.g., at the root 222 of the second airfoil 220). Additionally or alternatively, the first ring spacing may be defined as a circumferential distance between the first ring location 142 and the second ring location 242 (e.g., a circumferential distance between the root 122 and the root 222).

It should be appreciated that each airfoil 20 of a stage 100 may define a discrete ring spacing relative to the next successive airfoil 20 of the stage 100. In additional or alternative embodiments, two or more airfoils 20 may define the same ring spacing (e.g., a first ring spacing) while two or more airfoils 20 define the same ring spacing (e.g., a second ring spacing). Additionally or alternatively, the stage 100 may include one or more mating, groups, or individual further airfoils 20 defining any number of discrete ring pitches as desired or required.

From the above disclosure, it should be appreciated that including an airfoil assembly having a first airfoil (defining a first sweep angle, a first axial position, and a first pitch angle) and a second airfoil (defining a second sweep angle, a second axial position, a second ring spacing, and a second pitch angle), wherein the second sweep angle is different than the first sweep angle, the second axial position is different than the first axial position, the second pitch angle is different than the first pitch angle, or a combination thereof, may produce a more advantageous sound profile downstream of the airfoil assembly. Notably, the inventors of the present disclosure have found that varying the first axial position and the second axial position may be particularly relevant in a duct arrangement, taking into account the airflow characteristics in the duct environment.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. If these other examples 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, they are intended to be within the scope of the claims.

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

an airfoil assembly for a gas turbine engine defining an axial direction (a), a longitudinal axis extending along the axial direction (a), a radial direction (R), and a circumferential direction (C) relative to the longitudinal axis, the airfoil assembly comprising: an inner support structure configured to enclose the longitudinal axis of the gas turbine engine; an outer support structure configured to surround the longitudinal axis of the gas turbine engine, the outer support structure surrounding the inner support structure; and a stage comprising a plurality of airfoils extending from the inner support structure toward the outer support structure, the plurality of airfoils comprising: a first airfoil defining a first sweep angle, a first axial position, and a first lean angle; and a second airfoil defining a second sweep angle, a second axial position, and a second lean angle, wherein the second sweep angle is different from the first sweep angle, the second axial position is different from the first axial position, the second lean angle is different from the first lean angle, or a combination thereof.

The airfoil assembly according to one or more of these strips, wherein the inner support structure is configured to be fixed relative to a stationary support of the gas turbine engine, and wherein the plurality of airfoils are configured to remain stationary during operation of the gas turbine engine.

The airfoil assembly according to one or more of these strips, wherein the first airfoil and the second airfoil are each configured to reduce noise generated by the airfoil assembly via destructive interference during operation of the gas turbine engine.

The airfoil assembly according to one or more of these strips, wherein the second axial position is different from the first axial position.

The airfoil assembly according to one or more of these strips, wherein the second sweep angle is different from the first sweep angle.

The airfoil assembly according to one or more of these strips, wherein the second angle of inclination is different from the first angle of inclination.

The airfoil assembly according to one or more of these strips, wherein the first airfoil and the second airfoil are consecutive airfoils in the circumferential direction (C) of the stage of the airfoil assembly.

The airfoil assembly according to one or more of these strips, wherein at least one additional airfoil of the plurality of airfoils is positioned circumferentially between the first airfoil and the second airfoil.

The airfoil assembly according to one or more of these strips, wherein the outer support structure of the airfoil assembly is an outer support structure configured to be fixed to a support of the gas turbine engine, and wherein each airfoil of the plurality of airfoils extends from the inner support structure to the outer support structure.

The airfoil assembly according to one or more of these strips, wherein the outer support structure of the airfoil assembly is a shroud surrounding the plurality of airfoils, and wherein the plurality of airfoils are rotatable fan blades.

A gas turbine engine defining an axial direction (a), a longitudinal axis extending along the axial direction (a), a radial direction (R), and a circumferential direction (C) relative to the longitudinal axis, the gas turbine engine comprising: a turbine; a fan assembly rotatable by the turbine; and an airfoil assembly configured with the turbine, the fan assembly, or both, the airfoil assembly comprising: an inner support structure configured to enclose the longitudinal axis of the gas turbine engine; an outer support structure configured to enclose the longitudinal axis of the gas turbine engine and the inner support structure; and a stage comprising a plurality of airfoils extending from the inner support structure toward the outer support structure, the plurality of airfoils comprising: a first airfoil defining a first sweep angle, a first axial position, and a first lean angle; and a second airfoil defining a second sweep angle, a second axial position, and a second lean angle, wherein the second sweep angle is different from the first sweep angle, the second axial position is different from the first axial position, the second lean angle is different from the first lean angle, or a combination thereof.

The gas turbine engine of one or more of these strips, further comprising: an outer nacelle, and wherein the outer support structure is an outer support structure coupled to or integrated into the outer nacelle, wherein each airfoil of the plurality of airfoils extends from the inner support structure to the outer support structure.

The gas turbine engine of one or more of these strips, wherein the fan assembly comprises a fan, wherein the fan comprises a plurality of fan blades, and wherein the outer nacelle encloses the plurality of fan blades.

The gas turbine engine according to one or more of these strips, wherein the plurality of airfoils are a plurality of outlet guide vanes.

The gas turbine engine of one or more of these clauses, wherein the second tilt angle is different from the first tilt angle, wherein the second sweep angle is different from the first sweep angle, or both.

The gas turbine engine as recited in one or more of these strips, wherein the airfoil assembly is positioned within the turbine.

An airfoil assembly for a gas turbine engine defining an axial direction (a), a longitudinal axis extending along the axial direction (a), a radial direction (R), and a circumferential direction (C) relative to the longitudinal axis, the airfoil assembly comprising: an inner support structure configured to enclose the longitudinal axis of the gas turbine engine; and a stage including a plurality of airfoils extending outwardly from the inner support structure along the radial direction, the plurality of airfoils being a plurality of ductless airfoils and including: a first airfoil defining a first sweep angle and a first lean angle; and a second airfoil defining a second sweep angle and a second lean angle, wherein the second sweep angle is different from the first sweep angle, the second lean angle is different from the first lean angle, or a combination thereof.

The airfoil assembly according to one or more of these strips, wherein said gas turbine engine comprises a ductless fan, and wherein said plurality of airfoils are a plurality of outlet guide vanes for said ductless fan.

The airfoil assembly according to one or more of these strips, wherein the second angle of inclination is different from the first angle of inclination.

The airfoil assembly according to one or more of these strips, wherein the second sweep angle is different from the first sweep angle.

Claims (10)

1. An airfoil assembly for a gas turbine engine defining an axial direction, a longitudinal axis extending along the axial direction, a radial direction, and a circumferential direction relative to the longitudinal axis, the airfoil assembly comprising:

an inner support structure configured to enclose the longitudinal axis of the gas turbine engine;

an outer support structure configured to surround the longitudinal axis of the gas turbine engine, the outer support structure surrounding the inner support structure; and

a stage comprising a plurality of airfoils extending from the inner support structure toward the outer support structure, the plurality of airfoils comprising:

a first airfoil defining a first sweep angle, a first axial position, and a first lean angle; and

a second airfoil defining a second sweep angle, a second axial position and a second pitch angle,

Wherein the second sweep angle is different from the first sweep angle, the second axial position is different from the first axial position, the second tilt angle is different from the first tilt angle, or a combination thereof.

2. The airfoil assembly of claim 1, wherein the inner support structure is configured to be fixed relative to a stationary support of the gas turbine engine, and wherein the plurality of airfoils are configured to remain stationary during operation of the gas turbine engine.

3. The airfoil assembly of claim 1, wherein the first airfoil and the second airfoil are each configured to reduce noise generated by the airfoil assembly via destructive interference during operation of the gas turbine engine.

4. The airfoil assembly of claim 1, wherein the second axial position is different from the first axial position.

5. The airfoil assembly of claim 1, wherein the second sweep angle is different from the first sweep angle.

6. The airfoil assembly of claim 1, wherein the second angle of inclination is different than the first angle of inclination.

7. The airfoil assembly of claim 1, wherein the first airfoil and the second airfoil are continuous airfoils in the circumferential direction of the stage of the airfoil assembly.

8. The airfoil assembly of claim 1, wherein at least one additional airfoil of the plurality of airfoils is positioned circumferentially between the first airfoil and the second airfoil.

9. The airfoil assembly of claim 1, wherein the outer support structure of the airfoil assembly is an outer support structure configured to be secured to a support of the gas turbine engine, and wherein each airfoil of the plurality of airfoils extends from the inner support structure to the outer support structure.

10. The airfoil assembly of claim 1, wherein the outer support structure of the airfoil assembly is a shroud surrounding the plurality of airfoils, and wherein the plurality of airfoils are rotatable fan blades.