Variable pitch fan blade system

A gas turbine engine may have a fan section that includes a variable pitch mechanism for changing the pitch of fan blades. The gas turbine engine may include a fan hub that is rotatable about an engine central longitudinal axis of the gas turbine engine and a blade receiver that is for holding a fan blade and that is rotatably coupled to the fan hub and may be rotatable about a radial axis of the gas turbine engine. The gas turbine engine may include a spring element disposed between the fan hub and the blade receiver to bias the blade receiver radially outward and an annular retaining ring may be disposed around the radial axis and between the fan hub and the blade receiver to limit radially outward movement of the blade receiver to prevent the blade receiver from decoupling and disengaging from the fan hub.

FIELD

The present disclosure relates to gas turbine engines, and more specifically, variable pitch fan blades of gas turbine engines.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. A fan section may drive air along a bypass flowpath while a compressor section may drive air along a core flowpath. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. The compressor section typically includes low pressure and high pressure compressors, and the turbine section includes low pressure and high pressure turbines.

Some gas turbine engines may have a variable pitch fan blade configuration that enables the fan blades to be rotated about their respective radial axes to change the angle of attack of the fan blades.

SUMMARY

In various embodiments, the present disclosure provides a gas turbine engine that includes a fan hub, a blade receiver, a spring element, and an annular retaining ring. The fan hub, according to various embodiments, is rotatable about an engine central longitudinal axis of the gas turbine engine. The blade receiver is for holding a fan blade and the blade receiver is rotatably coupled to the fan hub and rotatable about a radial axis of the gas turbine engine, according to various embodiments. The spring element is disposed between the fan hub and the blade receiver and the spring element is biased to exert a radially outward force against the blade receiver, according to various embodiments. The annular retaining ring, according to various embodiments, is disposed around the radial axis and disposed between the fan hub and the blade receiver. The annular retaining ring may be configured to limit radially outward movement of the blade receiver to prevent the blade receiver from decoupling and disengaging from the fan hub.

In various embodiments, the annular retaining ring is circumferentially segmented. For example, the annular retaining ring may include a plurality of circumferential segments. In various embodiments, the gas turbine engine further includes a bearing disposed between the fan hub and the blade receiver. In various embodiments, the annular retaining ring is in direct engagement with the blade receiver. The direct engagement between the annular retaining ring and the blade receiver may be more than a point contact. In various embodiments, the annular retaining ring is in direct engagement with the fan hub.

In various embodiments, the gas turbine engine further includes a variable pitch mechanism having an actuation arm coupled to the blade receiver. A splined index ring may be coupled between the actuation arm of the variable pitch mechanism and the blade receiver. The annular retaining ring, according to various embodiments, is in direct engagement with the actuation arm of the variable pitch mechanism. In various embodiments, the spring element is positioned radially outward of the annular retaining ring. In various embodiments, the spring element has a wear coating. In various embodiments, the spring element is self-lubricating. In various embodiments, the spring element includes a wave spring that extends around the radial axis.

Also disclosed herein, according to various embodiments, is a method of assembling a gas turbine engine. The method may include positioning a spring element between a fan hub and a blade receiver. The method may further include compressing the spring element to enlarge an annular gap between the fan hub and the blade receiver. Still further, the method may include inserting a circumferential segment of a segmented annular retaining ring through the annular gap.

In various embodiments, inserting the circumferential segment through the annular gap is performed after compressing the spring element and while the annular gap is enlarged. In various embodiments, compressing the spring element includes moving the blade receiver radially inward. In various embodiments, the circumferential segment is one of a plurality of circumferential segments of the segmented annular retaining ring, wherein inserting the circumferential segment comprises individually inserting each circumferential segment of the plurality of circumferential segments through the annular gap to form the segmented annular retaining ring. In various embodiments, the method further includes positioning a splined index ring between the blade receiver and an actuation arm of a variable pitch mechanism after individually inserting each circumferential segment of the plurality of circumferential segments through the annular gap.

The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosure, it should be understood that other embodiments may be realized and that logical changes and adaptations in design and construction may be made in accordance with this disclosure and the teachings herein without departing from the spirit and scope of the disclosure. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation.

As used herein, “aft” refers to the direction associated with the exhaust (e.g., the back end) of a gas turbine engine. As used herein, “forward” refers to the direction associated with the intake (e.g., the front end) of a gas turbine engine.

A first component that is “axially outward” of a second component means that a first component is positioned at a greater distance in the aft or forward direction away from the longitudinal center of the gas turbine along the longitudinal axis of the gas turbine, than the second component. A first component that is “axially inward” of a second component means that the first component is positioned closer to the longitudinal center of the gas turbine along the longitudinal axis of the gas turbine, than the second component.

A first component that is “radially outward” of a second component means that the first component is positioned at a greater distance away from the engine central longitudinal axis than the second component. A first component that is “radially inward” of a second component means that the first component is positioned closer to the engine central longitudinal axis than the second component. In the case of components that rotate circumferentially about the engine central longitudinal axis, a first component that is radially inward of a second component rotates through a circumferentially shorter path than the second component. The terminology “radially outward” and “radially inward” may also be used relative to references other than the engine central longitudinal axis. For example, a first component of a combustor that is radially inward or radially outward of a second component of a combustor is positioned relative to the central longitudinal axis of the combustor.

With reference toFIG. 1, a gas turbine engine20is shown according to various embodiments. Gas turbine engine20may be a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines may include, for example, an augmentor section among other systems or features. In operation, fan section22can drive coolant (e.g., air) along a path of bypass airflow B while compressor section24can drive coolant along a core flowpath C for compression and communication into combustor section26then expansion through turbine section28. Although depicted as a turbofan gas turbine engine20herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

Gas turbine engine20may generally comprise a low speed spool30and a high speed spool32mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure36or engine case via several bearing systems38,38-1, and38-2. Engine central longitudinal axis A-A′ is oriented in the z direction on the provided x-y-z axes. It should be understood that various bearing systems38at various locations may alternatively or additionally be provided, including for example, bearing system38, bearing system38-1, and bearing system38-2.

Low speed spool30may generally comprise an inner shaft40that interconnects a fan42, a low pressure compressor44and a low pressure turbine46. Inner shaft40may be connected to fan42through a geared architecture48that can drive fan42at a lower speed than low speed spool30. Geared architecture48may comprise a gear assembly60enclosed within a gear housing62. Gear assembly60couples inner shaft40to a rotating fan structure. High speed spool32may comprise an outer shaft50that interconnects a high pressure compressor52and high pressure turbine54. A combustor56may be located between high pressure compressor52and high pressure turbine54. A mid-turbine frame57of engine static structure36may be located generally between high pressure turbine54and low pressure turbine46. Mid-turbine frame57may support one or more bearing systems38in turbine section28. Inner shaft40and outer shaft50may be concentric and rotate via bearing systems38about the engine central longitudinal axis A-A′, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.

The airflow of core flowpath C may be compressed by low pressure compressor44then high pressure compressor52, mixed and burned with fuel in combustor56, then expanded over high pressure turbine54and low pressure turbine46. Turbines46,54rotationally drive the respective low speed spool30and high speed spool32in response to the expansion.

Gas turbine engine20may be, for example, a high-bypass ratio geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine20may be greater than about six (6). In various embodiments, the bypass ratio of gas turbine engine20may be greater than ten (10). In various embodiments, geared architecture48may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture48may have a gear reduction ratio of greater than about 2.3 and low pressure turbine46may have a pressure ratio that is greater than about five (5). In various embodiments, the bypass ratio of gas turbine engine20is greater than about ten (10:1). In various embodiments, the diameter of fan42may be significantly larger than that of the low pressure compressor44, and the low pressure turbine46may have a pressure ratio that is greater than about five (5:1). Low pressure turbine46pressure ratio may be measured prior to inlet of low pressure turbine46as related to the pressure at the outlet of low pressure turbine46prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other gas turbine engines including direct drive turbofans. A gas turbine engine may comprise an industrial gas turbine (IGT) or a geared aircraft engine, such as a geared turbofan, or non-geared aircraft engine, such as a turbofan, or may comprise any gas turbine engine as desired.

The fan section22of the gas turbine engine20may have a variable pitch fan blade system100, according to various embodiments and with reference toFIG. 2. The variable pitch fan blade system100depicted inFIG. 2is a cross-sectional view of area85generally indicated inFIG. 1. The variable pitch fan blade system100is generally configured to enable rotation of fan blades about their respective centerline axes (referred to herein as “radial axes”106because the fan blades105generally extend radially outward from the engine central longitudinal axis A-A′ of the gas turbine engine20). Changing the pitch of the fan blades105, also referred to as changing the angle of attack of the fan blades105, affects the flow of air through the gas turbine engine20.

The variable pitch fan blade system100of the present disclosure provides an alternative and improved structure and assembly for radially retaining the fan blades.

In various embodiments, the gas turbine engine20includes a fan hub110, a blade receiver120for holding a fan blade105, a spring element130, and an annular retaining ring140. The fan hub110is rotatable about the engine central longitudinal axis A-A′ of the gas turbine engine20and the blade receiver120is rotatable about the radial axis106(e.g., centerline axis of the fan blade105), according to various embodiments. Said differently, the blade receiver120may be rotatably coupled to the fan hub110or may within a channel defined by the fan hub110. The spring element130is disposed between the fan hub110and the blade receiver120and is biased to exert a radially outward force against the blade receiver120, according to various embodiments. The annular retaining ring140may be disposed around the radial axis106and may be disposed between the fan hub110and the blade receiver120. The annular retaining ring140is generally configured to limit radially outward movement of the blade receiver120, thereby preventing the blade receiver120from decoupling and/or disengaging from the fan hub110, according to various embodiments.

In various embodiments, the gas turbine engine20further includes a variable pitch mechanism170having an actuation arm172that is coupled to the blade receiver120to control and actuate the pitch adjustments of the fan blades105via the fan blade receiver120. In various embodiments, a splined index ring160may be coupled between the actuation arm172of the variable pitch mechanism170and the blade receiver120. The splined index ring160may be secured in place via a retaining ring162which is assembled onto the blade receiver120.

In various embodiments, one or more bearings150may be disposed between the fan hub110and the blade receiver120to enable relative rotation between the fan hub110and the blade receiver120. In various embodiments, the one or more bearings150may be disposed between the fan hub110and the actuation arm172of the variable pitch mechanism170to enable relative rotation between the fan hub110and the actuation arm172.

In various embodiments, and with continued reference toFIG. 2, the spring element130and the annular retaining ring140rotate with the blade receiver120about the radial axis106. Though not having substantial bearing functionality with respect to the rotation of the blade receiver120about the radial axis106within the confines of the fan hub110, the spring element130and the annular retaining ring140may still experience some degree of relative rotation with the adjacent structures (e.g., the blade receiver120, the fan hub110, and/or the actuation arm172of the variable pitch mechanism170). Accordingly, the spring element130and/or the annular retaining ring140may have a wear coating and/or may be made of a self-lubricating material.

In various embodiments, the annular retaining ring140cooperates with the spring element130to radially retain the blade receiver120. Said differently, the biased spring element130may exert a radially outward force on the blade receiver120and the annular retaining ring140may counter the radially outward force. In various embodiments, the radially outward force created by the spring element130prevents the blade receiver120from falling radially inwards, for example when the fan hub110is not rotating around the central longitudinal axis A-A′ of the gas turbine engine20and thus there is no centrifugal load on the blade receiver120. In various embodiments, the annular retaining ring140prevents or at least limits movement of the blade receiver120in the opposite (i.e., radially outward) direction, thereby preventing the blade receiver120from decoupling and disengaging from the fan hub110.

In various embodiments, the spring element130is a wave spring that extends around the blade receiver. In various embodiments, the spring element130may include one or more helical coil springs that are configured to exert the radially outward force on the blade receiver120. The annular retaining ring140may be segmented and thus may have discrete circumferential segments that are detachable from each other. Forming the annular retaining ring140from a plurality of circumferential segments may enhance the ease of assembling the variable pitch fan blade system100. Additional details pertaining to the annular retention ring and how it is installed and/or assembled are included below with reference toFIGS. 3, 4A, and 4B.

In various embodiments, and with reference toFIG. 3, a method390of assembling a gas turbine engine is provided. The method390may include positioning the spring element130between the fan hub110and the blade receiver120at step392. The method390may further include compressing the spring element130to enlarge an annular gap125(seeFIG. 4A) at step394. In various embodiments, the annular gap125is defined between the fan hub110and the blade receiver120. With the annular gap125enlarged, the method390may include inserting individual circumferential segments through the annular gap at step396. The method390may further include positioning and orienting the circumferential segments so as to form the annular retaining ring140in the desired position/location between the fan hub110and the blade receiver120.

In various embodiments, and with reference toFIG. 4A, step394of the method390is provided and shown.FIG. 4Ashows a partial view of the variable pitch fan blade system100. Said differently,FIG. 4A, according to various embodiments, depicts the radially inward portion of the blade receiver120and depicts pertinent portions of the actuation arm172of the variable pitch mechanism170. Step394of the method390includes compressing the spring element130, which moves the blade receiver120in a radially inward direction and enlarges the annular gap125between the blade receiver120and the actuation arm172of the variable pitch mechanism170. In response to the annular gap125being enlarged, circumferential segments of the annular retaining ring140may be individually inserted through the annular gap125and the annular retaining ring140may be reassembled in the desired location between the blade receiver120and the fan hub110and/or the actuation arm172. In response to the annular retaining ring140being properly formed and positioned, the compression force causing the spring element130to compress may be released, thereby causing the blade receiver120to move in a radially outward direction due by the radially outward force exerted by the spring element130.

In various embodiments, and with reference toFIG. 4B, the annular retaining ring140is provided having multiple circumferential segments141,142,143,144,145,146. WhileFIG. 4Bshows the annular retaining ring140having six circumferential segments141,142,143,144,145,146, the annular retaining ring140may be divided into a different number of segments (e.g., 4, 8, 12, etc.). In various embodiments, the number of circumferential segments is dependent on the size of the annular gap125and the features and orientation of the adjacent components.

In various embodiments, the annular retaining ring140is made from a metallic material. For example, the annular retaining ring140may be made from stainless steel. In various embodiments, the blade receiver120and the actuation arm172of the variable pitch mechanism170annular have mating surfaces124,174, respectively, that are configured to match the cross-sectional shape of the annular retaining ring140. In such embodiments, the direct engagement between the annular retaining ring140and the respective mating surfaces124,174is more than a point contact. The term “more than a point contact” refers to direct mechanical contact between workpieces in which the contact interface between the two workpieces is a line or an area (e.g., more than a point). This increased contact area, as compared to an interface that is a single point (e.g., a ball bearing against a raceway), between two workpieces (e.g., between the annular retaining ring140and the blade receiver120) has a higher load capacity. Said differently, direct engagement between the annular retaining ring140and the blade receiver120may comprise a line engagement or an area engagement that provide sufficient contact area for transferring loads during, for example, operating of the fan section22of the gas turbine engine20. Accordingly, in various embodiments, the annular retaining ring140is made from a material that has high shear/transverse strength properties.

Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. Different cross-hatching is used throughout the figures to denote different parts but not necessarily to denote the same or different materials.