CN117006090A - Actuation assembly for a fan of a gas turbine engine - Google Patents

Abstract

A fan assembly for a gas turbine engine is provided, comprising: a plurality of fan blades including a first fan blade; and an actuation assembly, the actuation assembly comprising: a first linkage connected to the first fan blade; a first pivot point rotatable with the first fan blade, the first linkage being further connected to the first pivot point; a control point movable relative to the first pivot point and connected to the first linkage for changing the relative position of the first fan blade within the plurality of fan blades.

Description

Actuation assembly for a fan of a gas turbine engine

Technical Field

The present disclosure relates to an actuation assembly for a fan of a gas turbine engine.

Background

Gas turbine engines typically include a turbine and a rotor assembly. Gas turbine engines, such as turbofan engines, may be used for aircraft propulsion. In the case of a turbofan engine, the rotor assembly may be configured as a fan assembly. In at least some constructions, the turbofan engine may include an outer nacelle surrounding a plurality of fan blades of a fan of the fan assembly. The outer nacelle may provide benefits related to noise and blade containment. However, the inclusion of an outer nacelle may limit the fan diameter of the fan assembly because the size and weight of the outer nacelle typically also increases for larger diameter fans.

Thus, some turbofan engines may eliminate the outer nacelle. However, the inventors of the present disclosure have found that such a configuration may present certain problems, and methods of addressing these problems are welcome in the art.

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 cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure.

FIG. 2 is a schematic cross-sectional view of a forward end of the exemplary gas turbine engine of FIG. 1.

FIG. 3 is a view of a portion of a non-uniform blade actuator system, as seen along a longitudinal axis of a gas turbine engine, according to an exemplary aspect of the present disclosure.

FIG. 4 is a view of a portion of a non-uniform vane actuator system, as seen along a longitudinal axis of a gas turbine engine, according to another exemplary aspect of the present disclosure.

FIG. 5 is a schematic view of a fan section and an actuation assembly according to an exemplary aspect of the present disclosure, wherein the actuation assembly is depicted in an intermediate position.

FIG. 6 is a schematic view of a fan section and an actuation assembly according to an exemplary aspect of the present disclosure, wherein the actuation assembly is depicted in an offset position.

Detailed Description

Reference will now be made in detail to embodiments of the disclosure, 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 disclosure.

The word "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 specifically indicated otherwise.

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

The term "at least one", e.g. "at least one of A, B and C", in the context means any combination of a only, B only, C only or A, B and C.

The term "turbine" 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 versions of one or more of these engines.

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

The terms "low" and "high", or their respective degrees of comparison (e.g., more-, where applicable), when used with a compressor, turbine, shaft or spool piece, etc., each refer to relative speeds within the engine, unless otherwise specified. For example, a "low turbine" or "low speed turbine" defines a component configured to operate at a lower rotational speed than a "high turbine" or "high speed turbine" of the engine, such as a maximum allowable rotational speed.

The terms "forward" and "aft" refer to relative positions within the gas turbine engine or carrier, and refer to the normal operating attitude of the gas turbine engine or carrier. For example, for a gas turbine engine, forward refers to a location closer to the engine intake and aft refers to a location closer to the engine nozzle or exhaust.

As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to a centerline of a gas turbine engine. Furthermore, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to a centerline of the gas turbine engine. Furthermore, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations that extend in an arc about a centerline of a gas turbine engine.

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 the part and/or system. For example, approximating language may refer to the remaining amount of 1%, 2%, 4%, 10%, 15%, or 20%. 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.

The terms "coupled," "fixed," "attached," and the like, refer to a direct coupling, fixed or attachment, as well as an indirect coupling, fixed or attachment via one or more intermediate components or features, unless otherwise specified herein.

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

The present disclosure relates generally to an actuation assembly for a fan assembly of a gas turbine engine and a gas turbine engine including the actuation assembly. In at least some example embodiments, the fan assembly may be a ductless fan assembly, i.e., may not include an outer nacelle surrounding the fan assembly. During certain operations, the fan assembly may receive airflow that is not aligned with the fan axis of the fan assembly. For example, during operations in which the gas turbine engine defines a large angle of attack, such as take-off or climb operations, the airflow received by the fan may be misaligned with the fan axis. Similarly, during low speed operation with strong crosswinds, the airflow received by the fan may be misaligned with the fan axis. In such a configuration, the misaligned airflow may cause the fan blades on one side of the engine to have a higher load than the fan blades on the other side of the engine, thereby causing undesirable forces (also referred to as "1P" loads) to be generated on the fan assembly and the gas turbine engine at least once per revolution of the fan assembly.

To address this issue, the inventors propose an actuation assembly for a fan assembly that is capable of reconfiguring the fan blades to more evenly distribute the force under operating conditions that receive airflow that is not aligned with the fan axis. In particular, the inventors propose an actuation assembly having a first linkage connected to a first fan blade; a first pivot point rotatable with the first fan blade, the first linkage being connected to the first pivot point; a control point movable relative to the first pivot point and connected to the first linkage for changing the relative position of the first fan blade within the plurality of fan blades.

In certain exemplary aspects, the control point is movable between an intermediate position in which the fan blades each define an equal circumferential spacing and an offset position in which the fan blades define a varying circumferential spacing that varies based on the circumferential position of the respective fan blade. In this way, the actuation assembly may circumferentially distribute the fan blades to equalize forces on the fan assembly and the gas turbine engine despite misalignment of the incoming airflow with the fan axis.

Referring now to the drawings, in which like numerals indicate like elements throughout the several views, FIG. 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, for the embodiment of FIG. 1, the gas turbine engine is a high bypass turbofan jet engine, sometimes referred to as a "turbofan engine". As shown in FIG. 1, gas turbine engine 10 defines an axial direction A (extending parallel to longitudinal axis 12 for reference), a radial direction R, and a circumferential direction C extending about longitudinal axis 12. In general, the gas turbine engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14.

The exemplary turbine 16 shown generally includes a generally tubular outer housing 18 defining an annular inlet 20. The outer casing 18 encloses, in serial flow relationship, a compressor section including a booster or Low Pressure (LP) compressor 22 and a High Pressure (HP) compressor 24; a combustion section 26; a turbine section including a High Pressure (HP) turbine 28 and a Low Pressure (LP) turbine 30; and an injection exhaust nozzle section 32. A High Pressure (HP) shaft 34 (which may additionally or alternatively be a spool) drivingly connects HP turbine 28 to HP compressor 24. A Low Pressure (LP) shaft 36 (which may additionally or alternatively be a spool) drivingly connects LP turbine 30 to LP compressor 22. The compressor section, combustion section 26, turbine section, and injection exhaust nozzle section 32 together define a working gas flow path 37.

For the depicted embodiment, the fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As shown, the fan blades 40 extend outwardly from the disk 42 generally in the radial direction R. Each fan blade 40 is rotatable about a pitch axis P relative to the disk 42 in that the fan blades 40 are operatively coupled to a suitable pitch mechanism 44, which pitch mechanism 44 is configured to collectively (e.g., consistently) change the pitch of the fan blades 40. The gas turbine engine 10 also includes a power gearbox 46, and the fan blades 40, disk 42, and pitch mechanism 44 are rotatable about the longitudinal axis 12 through the LP shaft 36 together across the power gearbox 46. The power gearbox 46 includes a plurality of gears for adjusting the rotational speed of the fan 38 relative to the rotational speed of the LP shaft 36 so that the fan 38 may rotate at a more efficient fan speed.

Still referring to the exemplary embodiment of FIG. 1, the disk 42 is covered by a rotatable front hub 48 (sometimes also referred to as a "spinner") of the fan section 14, the front hub 48 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 40.

Further, the exemplary fan section 14 includes an annular fan housing or outer nacelle 50 that circumferentially surrounds at least a portion of the fan 38 and/or turbine 16. It should be appreciated that in the depicted embodiment, the outer nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Further, a downstream section 54 of the outer nacelle 50 extends over the exterior of the turbine 16 to define a bypass airflow passage 56 therebetween.

During operation of the gas turbine engine 10, a volume of air 58 enters the gas turbine engine 10 through the outer nacelle 50 and an associated inlet 60 of the fan section 14. As a volume of air 58 passes through the fan blades 40, a first portion 62 of the air is directed or routed into the bypass airflow passage 56, and a second portion 64 of the air is directed or routed into the working gas flow path 37, or more specifically, into the LP compressor 22. The ratio of the first portion of air 62 to the second portion of air 64 is commonly referred to as the bypass ratio. The pressure of the second portion 64 of air then increases as it is passed through the HP compressor 24 and into the combustion section 26, where it mixes with fuel and combusts to provide combustion gases 66.

The combustion gases 66 are channeled through HP turbine 28 wherein heat energy and/or a portion of the kinetic energy from combustion gases 66 are extracted via sequential stages of HP turbine stator vanes 68 coupled to outer casing 18 and HP turbine rotor blades 70 coupled to HP shaft 34, thereby causing HP shaft 34 to rotate, thereby supporting the operation of HP compressor 24. The combustion gases 66 are then channeled through LP turbine 30 wherein heat energy and a second portion of the kinetic energy from combustion gases 66 are extracted via sequential stages of LP turbine stator vanes 72 coupled to outer housing 18 and LP turbine rotor blades 74 coupled to LP shaft 36, thereby causing LP shaft 36 to rotate, thereby supporting operation of LP compressor 22 and/or rotation of fan 38.

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

However, it should be appreciated that the exemplary gas turbine engine 10 depicted in FIG. 1 is by way of example only, and that in other exemplary embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, although the illustrated gas turbine engine 10 is configured as a ducted gas turbine engine (i.e., including the outer nacelle 50), in other embodiments, the gas turbine engine 10 may be a ductless gas turbine engine (such that the fan 38 is a ductless fan with the outlet guide vanes 52 cantilevered from the outer casing 18). Additionally or alternatively, although gas turbine engine 10 is shown configured as a geared gas turbine engine (i.e., including power gearbox 46) and a variable pitch gas turbine engine (i.e., including fan 38 configured as a variable pitch fan), in other embodiments gas turbine engine 10 may additionally or alternatively be configured to directly drive the gas turbine engine (such that LP shaft 36 rotates at the same speed as fan 38), as a fixed pitch gas turbine engine (e.g., fan 38 includes fan blades 40 that are not rotatable about pitch axis P), or both. It should also be appreciated that in other exemplary embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of the present disclosure may be incorporated (as appropriate) into, for example, a turboprop gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine. Moreover, in other exemplary embodiments, aspects of the present disclosure may be incorporated into, for example, aeroderivative gas turbine engines, marine gas turbine engines, wind turbines, and the like.

Referring now to FIG. 2, a schematic cross-sectional view of a forward end of a gas turbine engine 10 is provided in accordance with an exemplary embodiment of the present disclosure. Specifically, FIG. 2 provides a schematic cross-sectional view of fan section 14 of gas turbine engine 10. In certain exemplary embodiments, the exemplary gas turbine engine 10 of FIG. 2 may be configured in substantially the same manner as the exemplary gas turbine engine 10 of FIG. 1. Thus, the same or similar numbers may refer to the same or similar parts.

As shown in FIG. 2, the fan section 14 (also referred to herein as a "fan assembly") generally includes a fan 38 configured as a variable pitch fan having a plurality of fan blades 40 coupled to a disk 42. Briefly, it should be appreciated that the fan 38 is configured as a forward thrust fan configured to generate thrust in a forward direction for the gas turbine engine 10 (and, for example, an aircraft incorporating the gas turbine engine 10). The "forward direction" may correspond to the forward direction of an aircraft containing the gas turbine engine 10, and in the illustrated embodiment refers to a direction pointing to the left.

Still referring to fig. 2, each fan blade 40 includes a base 80 at an inner end in the radial direction R. Each fan blade 40 is coupled to the disk 42 at a base 80 via a respective trunnion mechanism 82. Trunnion mechanism 82 facilitates rotation of a respective fan blade 40 about a pitch axis P of the respective fan blade 40. For the illustrated embodiment, the base 80 is configured as a dovetail that is received within a correspondingly shaped dovetail slot of the trunnion mechanism 82.

However, in other exemplary embodiments, the base 80 may be attached to the trunnion mechanism 82 in any other suitable manner. For example, the base 80 may be attached to the trunnion mechanism 82 using a pinned connection or any other suitable connection. In other exemplary embodiments, the base 80 may be integrally formed with the trunnion mechanism 82.

Further, as with the exemplary gas turbine engine 10 of FIG. 1, the fan 38 of the exemplary gas turbine engine 10 depicted in FIG. 2 is mechanically coupled to the turbine 16 (not depicted, see FIG. 1). More specifically, the exemplary variable pitch fan 38 of the gas turbine engine 10 of FIG. 2 is rotatable about the longitudinal axis 12 of the gas turbine engine 10 through the LP shaft 36 (not depicted, see FIG. 1) across the power gearbox 46. Specifically, the disk 42 is attached to the power gearbox 46 by a fan rotor 84, which fan rotor 84 includes one or more separate structural members 86 for the illustrated embodiment. The power gearbox 46 is in turn attached to the LP shaft 36 (not depicted, see fig. 1) such that rotation of the LP shaft rotates the fan rotor 84 and the plurality of fan blades 40, respectively. Notably, as also shown, the fan section 14 additionally includes a front hub 48 (which is capable of rotating with, for example, the disk 42 and the plurality of fan blades 40).

In addition, the fan 38 additionally includes a stationary fan frame 88 and one or more fan bearings 96 for supporting rotation of various rotating components of the fan 38 (e.g., the plurality of fan blades 40). More specifically, fan frame 88 supports the various rotating components of fan 38 via one or more fan bearings 96. For the illustrated embodiment, the one or more fan bearings 96 include a forward roller bearing 98 and an aft ball bearing 100. However, in other exemplary embodiments, any other suitable number and/or type of bearings may be provided to support rotation of the plurality of fan blades 40. For example, in other exemplary embodiments, the one or more fan bearings 96 may include a pair (two) of tapered roller bearings or any other suitable bearing.

Additionally, the exemplary fan 38 of the gas turbine engine 10 includes a pitch mechanism 44 for rotating each of the plurality of fan blades 40 about its respective pitch axis P.

Furthermore, the exemplary fan 38 of the gas turbine engine 10 illustrated in FIG. 2 further includes an actuation assembly 100. The actuation assembly 100 may generally be configured to move the plurality of fan blades 40 of the fan 38 between a uniform spacing in the circumferential direction C and a non-uniform spacing in the circumferential direction C, as will be explained in more detail below.

The actuation assembly 100 includes a plurality of linkages 102 connected to a corresponding plurality of fan blades 40. In particular, the actuation assembly 100 includes a first linkage 102A of the plurality of linkages 102 connected to a first fan blade 40A of the plurality of fan blades 40. More specifically, for the illustrated embodiment, first linkage 102A is rigidly coupled to trunnion mechanism 82, which trunnion mechanism 82 is in turn coupled to fan blade 40. The actuation assembly 100 further includes a first pivot point 104A rotatable with the first fan blade 40A, wherein the first linkage 102A is also connected to the first pivot point 104A. In particular, for the illustrated embodiment, the first pivot point 104A is capable of rotating with the fan rotor 84, and more particularly, is rigidly coupled to the fan rotor 84 by an extension arm 106. In this manner, the first linkage 102A and the first pivot point 104A are configured to rotate with the plurality of fan blades 40 during operation of the fan section 14 (also referred to herein as a fan assembly).

In addition, the illustrated example actuation assembly 100 also includes a control point 108. The control point 108 is depicted in fig. 2 as being in an intermediate position, but is configured to move in a radial direction R of the gas turbine engine 10 relative to a fan axis (not separately labeled; aligned with the longitudinal axis 12) of the fan 38 and the longitudinal axis 12 of the gas turbine engine 10. More specifically, control point 108 is configured to move relative to first pivot point 104A and is also coupled to first linkage 102A to change the configuration of plurality of fan blades 40, and more specifically, to change the relative position of first fan blade 40A within plurality of fan blades 40, as described below.

For the illustrated embodiment, a control point actuator 110 is provided to move the control point 108 relative to the longitudinal axis 12. The control point actuator 110 may be grounded to the static structure of the gas turbine engine 10 through the power gearbox 46.

Referring now to FIG. 3, a separate view of the first fan blade 40A and a subset of the actuation assemblies 100 of FIG. 2 is provided, as seen in the axial direction A of the gas turbine engine 10. It should be appreciated that the first fan blade 40A and the actuation assembly 100 are depicted in a reference plane 124 (see fig. 2) defined perpendicular to the longitudinal axis 12 of the gas turbine engine 10. As indicated by arrow 112, control point 108 is configured to move relative to longitudinal axis 12 within reference plane 124. Movement of the control point 108 is configured to move the first linkage 102A to control the angle 114 of the first fan blade 40A relative to the radial direction R about the first pivot point 104A. In particular, when in the neutral position, as shown in FIG. 3, the angle 114 of the first fan blade 40A relative to the radial direction R about the first pivot point 104A may be equal to zero. An angle 114 is defined between the radial direction R of the first fan blade 40A and the pivot axis P.

Fig. 3 depicts in phantom the control point 108 moving to a non-uniform position away from the longitudinal axis 12. When moved to the non-uniform position, the angle 114 defined by the fan blade 40 and the radial direction R about the first pivot point 104A may be greater than zero, such as greater than 5 ° and less than 45 °, such as at least 7.5 °, such as at least 10 °, such as at least 15 °, such as less than 30 °, and so forth. The angle 114 may be defined by the pitch axis P of the first fan blade 40A and the radial direction R.

Notably, to facilitate such movement of the control point 108 relative to the longitudinal axis 12 of the gas turbine engine 10 within the reference plane 124, the actuation assembly 100 includes a first linkage 102A. More specifically, for the illustrated embodiment, the first linkage 102A includes a first member 116 and a second member 118. The first member 116 is slidable relative to the second member 118. More specifically, for the illustrated embodiment, the first member 116 is retractable into the second member 118.

In this manner, the first linkage 102A may be a variable length linkage to facilitate movement of the control point 108 relative to the first pivot point 104A within the reference plane 124. For example, when the control point 108 is in the neutral position, the first linkage 102A is longer (shown in phantom in fig. 3) than the first linkage 102A when the control point 108 is in the non-uniform position. In this manner, it should be appreciated that the length of the linkage 102 may vary as the rotor rotates relative to the control point 108. More specifically, the length of the linkage 102 changes in a sinusoidal pattern once per revolution.

However, it should be appreciated that in other exemplary embodiments of the present disclosure, one or more linkages 102 of the plurality of linkages 102 may be configured in any other suitable manner. For example, referring now to fig. 4, a first fan blade 40A and an actuation assembly 100 according to another exemplary aspect of the present disclosure are provided. The exemplary first fan blade 40A and actuation assembly 100 of fig. 4 may be configured in substantially the same manner as the exemplary first fan blade 40A and actuation member of fig. 3. For example, the actuation assembly 100 generally includes a first linkage 102A that couples the first fan blade 40A; a first pivot point 104A rotatable with the first fan blade 40A (and the first linkage 102A is further connected to the first pivot point 104A); and a control point 108, the control point 108 being movable relative to the first pivot point 104A and relative to the longitudinal axis 12. The first linkage 102A is further connected to a control point 108.

However, for the illustrated embodiment, the first linkage 102A is not a slidable linkage, but rather includes a pivot joint 120. More specifically, the first linkage 102A includes a first member 116 and a second member 118, the first member 116 being pivotably connected to the second member 118 at a pivot joint 120. The first member 116 and the second member 118 define an angle 122, for example, between 15 ° and 165 °. Fig. 4 depicts the actuation assembly 100 in an intermediate position, and further depicts the first fan blade 40A and the actuation assembly 100 in a non-uniform position in phantom, wherein the control point 108 has moved within the reference plane 124 relative to the longitudinal axis 12. Control point 108 is generally movable in any suitable direction within reference plane 124, as indicated by arrow 112.

Referring now to fig. 5 and 6, a fan section 14 in accordance with exemplary aspects of the present disclosure is provided. The exemplary fan section 14 of fig. 5 and 6 may be configured in a similar manner as the exemplary fan section 14 described above with reference to fig. 2-4. In particular, the fan section 14 of fig. 5 and 6 includes a plurality of fan blades 40 and an actuation assembly 100. The actuation assembly 100 includes a plurality of linkages 102, each linkage 102 coupled to a respective one of the plurality of fan blades 40. Each linkage 102 of the plurality of linkages 102 may be configured in a similar manner as the first linkage 102A described above with respect to, for example, fig. 3 or 4.

Further, the actuation assembly 100 includes a plurality of pivot points 104 and a control point 108. Each pivot point 104 of the plurality of pivot points 104 is rotatable with a corresponding one of the plurality of fan blades 40. Each linkage 102 of the plurality of linkages 102 is connected to a respective pivot point 104 of the plurality of pivot points 104 and is further connected to a control point 108.

The control point 108 is movable relative to the plurality of pivot points 104 and is configured to change the relative position of at least one fan blade 40 within the plurality of fan blades 40. Specifically, the control point 108 is movable within a reference plane 124 (defined perpendicular to the longitudinal axis 12 of the gas turbine engine 10 incorporating the fan section 14; see FIG. 2; the plane depicted in FIGS. 5 and 6) relative to a fan axis (not labeled) and the longitudinal axis 12 of the gas turbine engine 10 incorporating the fan section 14 to vary the configuration of the plurality of fan blades 40 of the fan 38.

With particular reference first to FIG. 5 in particular, the actuation assembly 100 is depicted in an intermediate position in which the control point 108 is aligned with the longitudinal axis 12 of the gas turbine engine 10 incorporating the fan section 14. When the actuation assembly 100 is in the neutral position, each of the fan blades 40 defines a uniform spacing in the circumferential direction C. More specifically, when control point 108 is in an intermediate position aligned with longitudinal axis 12, the plurality of fan blades 40 define a first blade pitch 126 at a first circumferential position 128 and a second blade pitch 130 at a second circumferential position 132. The first blade pitch 126 is equal to the second blade pitch 130. The first blade pitch 126 and the second blade pitch 130 are each a linear distance between tips of two adjacent fan blades 40 at respective circumferential positions 128, 132.

In contrast, referring now specifically to fig. 6, the actuation assembly 100 is depicted in an offset position (also referred to herein as a non-uniform position), and more specifically, the control point 108 is in an offset position, separated from the longitudinal axis 12 within the reference plane 124. When the actuation assembly 100 is in the offset position, the plurality of fan blades 40 define a non-uniform spacing in the circumferential direction C. More specifically, when control point 108 is in an offset position within reference plane 124 that is separated from longitudinal axis 12, the plurality of fan blades 40 redefine a first blade pitch 126 at a first circumferential position 128 and a second blade pitch 130 at a second circumferential position 132. However, when the actuation assembly 100 is in the offset position, the first blade pitch 126 is different than the second blade pitch 130.

It will be appreciated that by moving the control point 108 to the offset position, at least some of the plurality of fan blades define an angle with the radial direction R (see, e.g., angle 114 in fig. 3 and 4) as compared to when the control point 108 is in the neutral position. In other words, by moving the control point 108 to the offset position, blade yaw is induced to at least some of the plurality of fan blades 40. The angle or value of blade yaw may be determined by the amount the control point 108 moves away from the longitudinal axis 12 within the reference plane 124, the direction the control point 108 moves away from the longitudinal axis 12 within the reference plane 124, the geometry of the plurality of linkages 102, or a combination thereof.

When the control point 108 is moved to the offset position, the circumferential spacing of the plurality of fan blades 40 and the angle of the plurality of fan blades 40 with respect to the radial direction R are set for each individual circumferential position. More specifically, as fan blades 40 rotate in circumferential direction C, they move into a spacing and blade angle configuration at a particular circumferential position determined by the offset position of control point 108 and the geometry of linkage 102. In this manner, it can be appreciated that as the first fan blade 40A of the plurality of fan blades 40 rotates in the circumferential direction C, the angle defined by the first fan blade 40A and the radial direction R changes from positive to negative and back again, e.g., in a sinusoidal pattern. Similarly, as the first fan blade 40A of the plurality of fan blades 40 rotates in the circumferential direction C, the spacing of the first fan blade 40A from circumferentially adjacent fan blades 40 also varies in a sinusoidal pattern.

As will be appreciated, changing the angle of the fan blade 40 relative to the radial direction R increases or decreases the angular velocity of the fan blade 40. For example, in the view of fig. 6, the first fan blade 40A located on the right defines an angle with the radial direction R that increases as the first fan blade 40 rotates in the circumferential direction C, as indicated by arrow 134. Accordingly, the rotational speed of the first fan blade 40A increases with increasing angle to the radial direction R. This results in an additional angular velocity of the first fan blade 40A and an additional linear velocity of the first fan blade 40A (see arrow 136), thereby effectively reducing the inflow angle with the oncoming airflow of the fan 38.

In contrast, still referring to fig. 6, a second fan blade 40B of the plurality of fan blades 40 that is located to the left of the view of fig. 6 defines an angle with the radial direction R that decreases as the second fan blade 40B rotates in the circumferential direction C, as indicated by arrow 138. In this way, the rotational speed of the second fan blade 40B decreases as the angle to the radial direction R also decreases. This may result in a decrease in the angular velocity of the second fan blade 40B and a decrease in the linear velocity of the second fan blade 40B (see arrow 140). This effectively increases the inflow angle to the oncoming airflow of the fan 38.

These local variations of the fan 38 in the circumferential direction C may affect the loading of the fan blades 40 based on the circumferential position of the respective fan blades 40. In this manner, the fan blades 40 may be configured to reduce 1P loading due to head-on airflow using the actuation assembly 100, wherein the fan 38 defines an oblique angle with the longitudinal axis 12, whether from a steep angle of attack, a negative angle of attack, starboard or port side wind, or the like.

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

a fan assembly for a gas turbine engine, comprising: a plurality of fan blades including a first fan blade; an actuation assembly, the actuation assembly comprising: a first linkage connected to the first fan blade; a first pivot point rotatable with the first fan blade, the first linkage being further connected to the first pivot point; and a control point movable relative to the first pivot point and connected to the first linkage for changing the relative position of the first fan blade among the plurality of fan blades.

The fan assembly of one or more of the preceding clauses, wherein the first linkage and the first pivot point are configured to rotate with the plurality of fan blades.

The fan assembly of one or more of the preceding clauses, wherein the first linkage comprises a first member and a second member, wherein the first member is slidable relative to the second member.

The fan assembly of one or more of the preceding clauses, wherein the first linkage comprises a first member and a second member, wherein the first member and the second member define an angle between 15 degrees and 165 degrees.

The fan assembly of one or more of the preceding clauses, wherein the fan assembly defines a fan axis, wherein the control point is movable relative to the fan axis.

The fan assembly of one or more of the preceding clauses, wherein the fan assembly defines a reference plane perpendicular to the fan axis, wherein the control point is movable relative to the fan axis within the reference plane.

The fan assembly of one or more of the preceding clauses, wherein the actuation assembly further comprises a plurality of linkages and a plurality of pivot points, wherein each pivot point is rotatable with a respective fan blade of the plurality of fan blades, wherein each linkage of the plurality of linkages is connected to a respective fan blade of the plurality of fan blades, to a respective pivot point of the plurality of pivot points, and to the control point.

The fan assembly of one or more of the preceding clauses, wherein the fan assembly defines a fan axis, wherein the control point is movable relative to the fan axis to change the configuration of the plurality of fan blades.

The fan assembly of one or more of the preceding clauses, wherein the control point is movable to an offset position separate from the fan axis, wherein the plurality of fan blades define a first blade pitch at a first circumferential position and a second blade pitch at a second circumferential position when the control point is in the offset position, wherein the first blade pitch is different from the second blade pitch.

The fan assembly of one or more of the preceding clauses, wherein the control point is movable to an intermediate position aligned with the fan axis, wherein the plurality of fan blades define a first blade pitch at a first circumferential position and a second blade pitch at a second circumferential position when the control point is in the intermediate position, wherein the first blade pitch is equal to the second blade pitch.

A gas turbine engine, comprising: a fan assembly of one or more of the preceding clauses.

A gas turbine engine, comprising: a turbine; and a fan assembly rotatable by the turbine, the fan assembly including a plurality of fan blades including a first fan blade and an actuation assembly including: a first linkage connected to the first fan blade; a first pivot point rotatable with the first fan blade, the first linkage being connected to the first pivot point; a control point movable relative to the first pivot point and connected to the first linkage for changing the relative position of the first fan blade among the plurality of fan blades.

The gas turbine engine of one or more of the preceding clauses, wherein the first linkage and the first pivot point are rotatable with the plurality of fan blades.

The gas turbine engine of one or more of the preceding clauses, wherein the fan assembly defines a fan axis, wherein the control point is movable relative to the fan axis.

The gas turbine engine of one or more of the preceding clauses, wherein the fan assembly defines a reference plane perpendicular to the fan axis, wherein the control point is movable relative to the fan axis within the reference plane.

The gas turbine engine of one or more of the preceding clauses, wherein the actuation assembly further comprises a plurality of linkages and a plurality of pivot points, wherein each pivot point is rotatable with a respective fan blade of the plurality of fan blades, wherein each linkage of the plurality of linkages is connected to a respective fan blade of the plurality of fan blades, to a respective pivot point of the plurality of pivot points, and to the control point.

The gas turbine engine of one or more of the preceding clauses, wherein the fan assembly defines a fan axis, wherein the control point is movable to an offset position separate from the fan axis, wherein the plurality of fan blades define a first blade pitch at a first circumferential position and a second blade pitch at a second circumferential position when the control point is in the offset position, wherein the first blade pitch is different from the second blade pitch.

The gas turbine engine of one or more of the preceding clauses, wherein the fan assembly defines a fan axis, wherein the control point is movable to an intermediate position separate from the fan axis, wherein the plurality of fan blades define a first blade pitch at a first circumferential position and a second blade pitch at a second circumferential position when the control point is in the intermediate position, wherein the first blade pitch is equal to the second blade pitch.

An actuation assembly for a fan assembly of a gas turbine engine, the fan assembly comprising a plurality of fan blades including a first fan blade, the actuation assembly comprising: a first linkage configured to be connected to the first fan blade; a first pivot point rotatable with the first fan blade when the actuation assembly is installed in the gas turbine engine, the first linkage device being connected to the first pivot point; and a control point movable relative to the first pivot point and connected to the first linkage for changing a relative position of the first fan blade within the plurality of fan blades when the actuation assembly is installed in the gas turbine engine.

The actuation assembly of one or more of the preceding clauses, wherein the first linkage comprises a first member and a second member, wherein the first member is slidable relative to the second member.

The actuation assembly of one or more of the preceding clauses, wherein the first linkage comprises a first member and a second member, wherein the first member and the second member define an angle between 15 degrees and 165 degrees.

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. These 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 language of the claims.

Claims (10)

1. A fan assembly for a gas turbine engine, comprising:

a plurality of fan blades, the plurality of fan blades comprising a first fan blade; and

an actuation assembly, the actuation assembly comprising:

a first linkage connected to the first fan blade;

a first pivot point rotatable with the first fan blade, the first linkage further connected to the first pivot point; and

a control point movable relative to the first pivot point and connected to the first linkage for changing a relative position of the first fan blade within the plurality of fan blades.

2. The fan assembly of claim 1, wherein the first linkage and the first pivot point are configured to rotate with the plurality of fan blades.

3. The fan assembly of claim 1, wherein the first linkage comprises a first member and a second member, wherein the first member is slidable relative to the second member.

4. The fan assembly of claim 1 wherein the first linkage comprises a first member and a second member, wherein the first member and the second member define an angle between 15 degrees and 165 degrees.

5. The fan assembly of claim 1, wherein the fan assembly defines a fan axis, wherein the control point is movable relative to the fan axis.

6. The fan assembly of claim 5 wherein the fan assembly defines a reference plane perpendicular to the fan axis, wherein the control point is movable within the reference plane relative to the fan axis.

7. The fan assembly of claim 1, wherein the actuation assembly further comprises a plurality of linkages and a plurality of pivot points, wherein each pivot point is rotatable with a respective fan blade of the plurality of fan blades, wherein each linkage of the plurality of linkages is connected to a respective fan blade of the plurality of fan blades, to a respective pivot point of the plurality of pivot points, and to the control point.

8. The fan assembly of claim 7, wherein the fan assembly defines a fan axis, wherein the control point is movable relative to the fan axis to change the configuration of the plurality of fan blades.

9. The fan assembly of claim 8, wherein the control point is movable to an offset position separate from the fan axis, wherein the plurality of fan blades define a first blade pitch at a first circumferential position and a second blade pitch at a second circumferential position when the control point is in the offset position, wherein the first blade pitch is different from the second blade pitch.

10. The fan assembly of claim 8, wherein the control point is movable to an intermediate position aligned with the fan axis, wherein when the control point is in the intermediate position, the plurality of fan blades define a first blade pitch at a first circumferential position and a second blade pitch at a second circumferential position, wherein the first blade pitch is equal to the second blade pitch.