Low-ratio fan blade for a gas turbine engine

Abstract

Methods, apparatus, systems, and articles of manufacture are disclosed for implementing low radius ratio fan blades for gas turbine engines. A fan blade for a gas turbine engine includes an airfoil including a leading edge and a trailing edge extending between a root and a tip of the airfoil, a distance between the leading edge and the trailing edge defining a flow path length, a leading edge hub point defined by a radially innermost point of the leading edge; and an axial dovetail including a pair of opposing pressure faces, an axial length of the axial dovetail being less than the flow path length, a top edge of the axial dovetail separating the pressure faces from the airfoil, the top edge of the axial dovetail being radially outward of the leading edge hub point.

Description

[0001] Related Applications

[0002] This patent is a continuation of and claims priority to U.S. Patent Application No. 17 / 535,291, filed November 24, 2021. The entire contents of U.S. Patent Application No. 17 / 535,291 are incorporated herein by reference. Priority to U.S. Patent Application No. 17 / 535,291 is hereby claimed. TECHNICAL FIELD

[0003] The present disclosure relates generally to gas turbines, and more particularly, to fan blades for a gas turbine engine. BACKGROUND

[0004] Gas turbine engines generally include, in serial flow order, an inlet section, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air enters the inlet section and flows to the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section, producing combustion gases. The combustion gases flow from the combustion section through a hot gas path defined within the turbine section, and then exit the turbine section via the exhaust section. BRIEF DESCRIPTION OF DRAWINGS

[0005] Figure 1 An example gas turbine engine is shown.

[0006] Figure 2 An example cross-sectional side view of an inlet section of an example gas turbine engine is shown.

[0007] Figure 3 An example fan blade is shown.

[0008] Figure 4 An example cross-sectional side view of an inlet section of an example gas turbine engine including a portion of an example fan blade is shown.

[0009] Figure 5 A partial perspective side view of an example inlet section including a plurality of example fan blades is shown.

[0010] The drawings are not to scale. Instead, thicknesses of layers or regions can be exaggerated in the drawings for illustration purposes. Although layers and regions are shown to have clear lines and boundaries in the drawings, some or all of these lines and / or boundaries can be idealized. In reality, boundaries and / or lines can be blurred, mixed, and / or irregular. Generally, identical reference numbers have been used to identify identical elements throughout the several figures and the accompanying written description. As used herein, unless otherwise stated, the term "over" describes a relative position of two parts. A first part is over a second part if at least one part is between the Earth and the first part than the second part. Likewise, as used herein, a first part is "under" a second part when the first part is closer to the Earth than the second part. As described above, a first part can be over or under a second part with one or more of: having other parts in between, having no other parts in between, the first part and the second part being in contact, or the first part and the second part not being in direct contact with each other. As used in this patent, stating that any part (e.g., a layer, film, zone, region, or plate) is on (e.g., positioned on, located on, disposed on, or formed on, etc.) another part in any manner indicates that the part is in contact with the other part, or that the part is located above the other part with one or more intervening parts in between. As used herein, unless otherwise stated, a connection reference (e.g., attached, coupled, connected, and joined) can include intervening members between the elements referenced by the connection reference and / or relative movement between those elements. Thus, a connection reference does not necessarily infer that two elements are directly connected and / or in fixed relation to one another. As used herein, stating that any part is "in contact" with another part is defined to mean that there are no intervening parts between the two parts.

[0011] Unless explicitly stated otherwise, descriptors such as "first," "second," "third," etc. are used herein without connoting or otherwise implying any meaning as to priority, physical order, arrangement on a list, and / or any ordering in any manner, but are merely used as labels and / or arbitrary designations to distinguish elements in order to facilitate an understanding of the disclosed examples. In some examples, a descriptor "first" can be used to refer to an element in the detailed description, while a different descriptor (e.g., "second" or "third") can be used in the claims to refer to the same element. In such instances, it should be understood that such descriptors are used merely for clarity in identifying those elements that can otherwise share the same designation, for example. As used herein, "approximately" and "about" refer to dimensions that can not be exact due to manufacturing tolerances and / or other real-world imperfections. As used herein, "substantially the same dimensions" refers to dimensions that can not be exactly the same due to manufacturing tolerances and / or other real-world imperfections. Thus, "substantially the same dimensions" refers to + / - 10% of the dimensions, unless otherwise stated. As used herein, the phrase "communication," including variants thereof, encompasses direct communication and / or indirect communication through one or more intermediary components, and does not require direct physical (e.g., wired) communication and / or continuous communication, but additionally includes selective communication at periodic intervals, predetermined intervals, non-periodic intervals, and / or one-time events. DETAILED DESCRIPTION

[0012] Many known technologies aim to improve the performance of a gas turbine engine. Some known technologies aim to reduce the mass of a gas turbine engine while at least maintaining technical specifications and / or performance. For example, some technologies can aim to improve the performance (e.g., thrust, fuel economy, etc.) of a gas turbine engine while maintaining a given packaging size (e.g., diameter of a fan case). An example technical specification of a gas turbine engine that impacts performance is a radius ratio (e.g., ratio of the radius of the hub of a fan section to the radius of the tip of a fan blade). It is desirable for a fan to have a small radius ratio because a small radius ratio increases the fan inlet area for a given fan diameter, allowing for increased thrust. Examples disclosed herein can provide a reduced radius ratio by utilizing fan blades with shortened blade roots to lower the leading edge hub point.

[0013] In the following detailed description, reference is made to the accompanying drawings that form a part hereof, and in which are shown, by way of illustration, specific examples in which the subject matter can be practiced. These examples are described in sufficient detail to enable those skilled in the art to practice the subject matter, and it is to be understood that other examples can be utilized. The following detailed description is, therefore, not to be taken in a limiting sense. Certain features from the various aspects of the description below can be combined to form yet further aspects of the subject matter discussed below.

[0014] The terms “upstream” and “downstream” refer to the relative direction with respect to the fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. As used herein, “vertical” refers to a direction perpendicular to the ground. As used herein, “horizontal” refers to a direction parallel to the centerline of the turbofan 100. As used herein, “lateral” refers to a direction perpendicular to the axial vertical direction (e.g., into and out of the plane of the drawing, etc.). Figure 1 , 2 ,

[0015] Various terms are used herein to describe the orientation of features. As used herein, the orientation of features, forces, and moments are described with reference to the axial, radial, and circumferential directions of the vehicle with which the features, forces, and moments are associated. Typically, the drawings are annotated with a set of axes including an axial axis A, a radial axis R, and a circumferential axis C. Additionally or alternatively, the drawings are annotated with a set of axes including a roll axis R, a pitch axis P, and a yaw axis Y.

[0016] “Comprise,” “comprising,” “include,” “including,” and “the complement of” (and all forms of these terms) as used herein are used in their open-ended, conventional sense, and are intended to allow for the inclusion of additional elements, terms, etc. As used herein, the phrase “at least” as used, e.g., as the preamble to a claim, is open-ended, in the same manner as the terms “comprising” and “including” are open-ended. The term “and / or” as used in a phrase such as A, B, and / or C, for example, means A alone, B alone, C alone, A with B, A with C, B with C, or A with B and with C. As used herein in the context of describing structures, components, items, objects, and / or things, the phrase “at least one of’ A and B is intended to refer to implementations including any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing structures, components, items, objects, and / or things, the phrase “at least one of’ A or B is intended to refer to implementations including any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. As used herein in the context of describing performance or execution of processes, instructions, actions, activities, and / or steps, the phrase “at least one of’ A and B is intended to refer to implementations including any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing performance or execution of processes, instructions, actions, activities, and / or steps, the phrase “at least one of’ A or B is intended to refer to implementations including any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B.

[0017] As used herein, singular references (e.g., “a,” “an,” “the,” “first,” “second,” etc.) do not exclude the plural. As used herein, the term “a” or “an” entity refers to one or more than one of that entity. The terms “a” (or “an”), “one or more,” and “at least one” are used interchangeably herein. Furthermore, although individually listed, a plurality of means, elements or method actions can be implemented by, e.g., a single unit or processor. Additionally, although individual features can be included in different examples or claims, these can possibly be combined, and the inclusion in different examples or claims does not imply that a combination of features is not feasible and / or advantageous.

[0018] A gas turbine engine includes a fan section proximate an engine air intake. The fan section includes a plurality of circumferentially spaced apart fan blades. Each fan blade extends radially outward from a rotor disk or hub. An example fan blade includes an airfoil and an integral dovetail at an airfoil root. The airfoil extends in a radial direction from the airfoil root to a blade tip. A surface of the airfoil extends in a chordal direction from a leading edge to a trailing edge. A width of the airfoil, defined by a distance between the leading edge and the trailing edge, defines a flow path length of the airfoil. The fan blade dovetail is received in a complementary dovetail slot formed in the rotor disk. Example fan blades can be made of metal (e.g., titanium) or composite material (e.g., carbon-epoxy composite system).

[0019] In known fan blades, the length of the dovetail is approximately the same as the width of the airfoil (e.g., within 10%). The size (e.g., length, cross-sectional area, etc.) of the dovetail is proportional to the force required to retain the fan blade within the dovetail slot. The force required to retain the fan blade within the dovetail slot is proportional to parameters of the fan section and / or fan blade (e.g., tip speed of the fan blade and weight of the fan blade). In some examples, the force required to retain the fan blade is reduced. For example, due to the implementation of a geared fan, the fan section of the gas turbine engine can have a reduced speed relative to the compressor or combustion section of the gas turbine engine. In this example, the tip speed of the fan blade is reduced, thereby reducing the force required to retain the fan blade. In other examples, the fan blade can be lightened by using advanced materials (e.g., composite systems), thereby reducing the force required to retain the fan blade within the dovetail slot. In these examples, the size (e.g., length, cross-sectional area, etc.) of the dovetail can be reduced.

[0020] The inlet area of the fan is defined by the space between the rotor disk and the fan blade tip. As used herein, the radius ratio is the ratio of the radius of the rotor disk to the radius of the blade tip. As noted above, for a given fan diameter, reducing the radius ratio allows for an increase in the fan inlet area. In some examples, for a given fan diameter, if the radius ratio is reduced, the thrust can be increased. In other examples, if the radius ratio is reduced, a lower number of blades can be implemented while maintaining engine performance, thereby reducing the mass, complexity, cost, and fuel usage of the engine. Examples disclosed herein reduce the radius ratio of a gas turbine engine fan by implementing a shortened dovetail length such that the axial dovetail length is less than the flow path length of the fan blade. Examples disclosed herein reduce the radius ratio of a gas turbine engine fan by implementing a fan blade having a leading edge hub point radially inward of the attachment pressure face.

[0021] Reference will now be made in detail to the examples of the present disclosure, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the present disclosure and is not meant as a limitation on the present disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present disclosure without departing from the scope or spirit of the disclosure. For instance, features illustrated or described as part of one example, can be used with another example to yield a still further example. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims and their equivalents.

[0022] Figure 1 is a schematic cross-sectional view of a prior art turbofan gas turbine engine 100 (“turbofan 100”). As shown, the turbofan 100 defines a longitudinal or axial centerline axis 102 extending therethrough for reference. Generally, the turbofan 100 can include a core turbine 104 or gas turbine engine disposed downstream of a fan section 106. Figure 1

[0023] The core turbine 104 generally includes a substantially tubular outer casing 108 (“turbine casing 108”) defining an annular inlet 110. The outer casing 108 can be formed of a single casing or multiple casings. The outer casing 108 encloses, in serial flow relationship, a compressor section having a booster or low pressure compressor 112 (“LP compressor 112”) and a high pressure compressor 114 (“HP compressor 114”), a combustion section 116, a turbine section having a high pressure turbine 118 (“HP turbine 118”) and a low pressure turbine 120 (“LP turbine 120”), and an exhaust section 122. A high pressure shaft or spool 124 (“HP shaft 124”) drivingly couples the HP turbine 118 and the HP compressor 114. A low pressure shaft or spool 126 (“LP shaft 126”) drivingly couples the LP turbine 120 and the LP compressor 112. The LP shaft 126 can also be coupled to a fan shaft or spool or axle 128 (“fan shaft 128”) of the fan section 106. In some examples, the LP shaft 126 can be directly coupled to the fan shaft 128 (i.e., a direct drive configuration). In alternative configurations, the LP shaft 126 can be coupled to the fan shaft 128 via a reduction gearbox 130 (e.g., an indirect drive or geared drive configuration).

[0024] Figure 1 ​​As shown, fan section 106 includes a plurality of fan blades 132 coupled to and extending radially outward from fan shaft 128. An annular fan housing or nacelle 134 circumferentially surrounds at least a portion of fan section 106 and / or core turbine 104. Nacelle 134 is supported relative to core turbine 104 by a plurality of circumferentially spaced outlet guide vanes 136. Furthermore, a downstream section 138 of nacelle 134 may surround an external portion of core turbine 104 to define a bypass airflow passage 140 therebetween.

[0025] like Figure 1 As shown, air 142 enters the inlet portion 144 of the turbofan 100 during operation of the turbofan 100. A first portion 146 of the air 142 flows into the bypass airflow passage 140, while a second portion 148 of the air 142 flows into the annular inlet 110 of the LP compressor 112. One or more successive stages of the LP compressor stator blades 150 and rotor blades 152, coupled to the LP shaft 126, progressively compress the second portion 148 of the air 142 flowing through the LP compressor 112 and directing it to the HP compressor 114. Next, one or more successive stages of the HP compressor stator blades 154 and rotor blades 156, coupled to the HP shaft 124, further compress the second portion 148 of the air 142 flowing through the HP compressor 114. This provides compressed air 158 to the combustion section 116, where it mixes with fuel and burns to provide combustion gases 160.

[0026] Combustion gas 160 flows through HP turbine 118, where one or more successive stages of HP turbine stator blades 162 and HP turbine rotor blades 164, coupled to HP shaft 124, extract a first portion of kinetic and / or thermal energy from the combustion gas 160. This energy extraction supports the operation of HP compressor 114. Combustion gas 160 then flows through LP turbine 120, where one or more successive stages of LP turbine stator blades 166 and LP turbine rotor blades 168, coupled to LP shaft 126, extract a second portion of thermal and / or kinetic energy from it. This energy extraction causes LP shaft 126 to rotate, thereby supporting the operation of LP compressor 112 and / or the rotation of fan shaft 128. Combustion gas 160 then exits core turbine 104 through its exhaust section 122.

[0027] Along with turbofan 100, core turbine 104 serves a similar purpose and a similar environment is seen in land-based gas turbines, turbojet engines where the ratio of the first portion 146 to the second portion 148 of air 142 is smaller than that in turbofans (e.g., turbofan 100), and ductless fan engines where fan section 106 lacks nacelle 134. In each turbofan, turbojet, and ductless engine, a reduction gear (e.g., reduction gearbox 130) may be included between any shaft and spool. For example, reduction gearbox 130 may be disposed between LP shaft 126 and fan shaft 128 of fan section 106.

[0028] Figure 2 It shows that it can be used Figure 1 The example inlet portion 144 is implemented in the example turbine fan 100 shown. The example inlet portion 144 includes a fan section 106 that rotates about a centerline axis 102 via a fan shaft 128 powered by an LP turbine 120 (not shown). The fan section 106 includes a rotor disk 202 and a plurality of circumferentially spaced fan or rotor blades 132. Figure 2 (Only one is shown in the image) Extends radially outward from rotor disk 202. Fan blades 132 may be metallic or non-metallic. For example, fan blades 132 may be made of carbon fiber-epoxy composite material or other similar materials. Rotor disk 202 includes a front end 204 axially spaced from rear end 206 and a radially outer surface 208 extending therebetween.

[0029] Located downstream of fan section 106 is an LP compressor 112 with axially spaced blades and blade rows, its blades engaging with LP shaft 126. In the illustrated example, LP shaft 126 is securely engaged to the rear end 206 of rotor disk 202 by a plurality of bolts 210. Rotor 212 is engaged to the front end 204 of rotor disk 202 to provide an aerodynamic flow path for air 142 entering fan section 106.

[0030] Figure 3 It shows that it can be used Figure 1 and / or Figure 2 An example fan blade 300 is implemented in the inlet portion 144 of the turbine fan 100 shown. The example fan blade 300 has a pressure side 302 and an opposing suction side 304. The example fan blade 300 is an integral part including a root section 306 and an airfoil 308. The root section 306 includes a straight axial dovetail tenon 310 having a pair of opposing pressure surfaces 311. When the fan blade 300 is assembled on a rotor disk (e.g., Figure 2 When the rotor disk 202 depicted in the figure is in the rotor disk, the axial dovetail tenon 310 is provided in the corresponding dovetail groove in the rotor disk (e.g., as shown in the figure). Figure 5 As shown). AlthoughFigure 3 The example fan blade 300 includes an axial dovetail 310, but other example fan blades can include dovetails having different geometries, such as skewed, circular-arc, etc. Figure 3 The axial dovetail 310 of the example fan blade 300 can represent any such geometrically shaped dovetail (e.g., skewed, circular-arc, etc.).

[0031] Still referring to Figure 3 The airfoil 308 of the example fan blade 300 extends in a chordwise direction (e.g., along the axis B) from a leading edge 312 to a trailing edge 314. Additionally, the example airfoil 308 extends in a spanwise or radial direction from a root 316 to a tip 318. Figure 3 The example fan blade 300 can be constructed from a variety of materials, including metals, metal alloys, non-metallic composites, and combinations thereof. In the illustrated example, the fan blade 300 is constructed from a composite material. The term “composite material” generally refers to a material that includes a reinforcing material, such as fibers or particles that are supported in a binder or matrix material. The composite can include a plurality of layers or plies that are embedded in a matrix and oriented substantially parallel to the pressure side 302 and the suction side 304. An example of a suitable material is a carbonaceous (e.g., graphite) fiber that is embedded in a resin material, such as an epoxy resin. These are commercially available as fibers that are arranged unidirectionally into a tape that is impregnated with resin. This “pre-impregnated” tape can be formed into a part shape and cured through an autoclave process or compression molding to form a lightweight, relatively uniform filler article.

[0032] As discussed above, during operation of a gas turbine engine (e.g., the turbofan 100), air (e.g., the air 142) enters the inlet portion (e.g., Figure 1 and / or Figure 2 the inlet portion 144 of the example fan blade 300) and flows substantially axially toward the compressor, combustor, and / or turbine section. Thus, the air flows through the rotors, stators, and / or fan blades in the gas turbine engine. For example, when the fan blade 300 is mounted to the rotor disk (e.g., the rotor disk 202 depicted in Figure 2 , the air can flow through the example fan blade 300. As used herein, the flow path length is the distance that the air flows through the fan blade 300. In Figure 3 the example, the flow path length of the fan blade 300 is defined by the distance between the leading edge 312 and the trailing edge 314.

[0033] In Figure 3 the example, the axial dovetail 310 includes a forward end 320 and an aft end 322. The length of the axial dovetail 310 is defined by the distance between the forward end 320 and the aft end 322. In known fan blades, such as the fan blade 132 of Figure 2 , the forward end of the axial dovetail is aligned with the leading edge of the airfoil. In the example fan blade 300, the forward end 320 of the axial dovetail 310 is offset from the leading edge 312 of the airfoil 308.Figure 3 In the example, the front end 320 of the axial dovetail tenon 310 is positioned in a chordal direction from the leading edge 312 toward the trailing edge 314. Similarly, in examples such as Figure 2 In the known fan blades of fan blade 132, the rear end of the axial dovetail tenon is aligned with the trailing edge of the airfoil. Figure 3 In the example, the rear end 322 of the axial dovetail tenon 310 is positioned in the chordal direction from the rear edge 314 toward the front edge 312. As a result, in Figure 3 In some examples, the length of the axial dovetail 310 is less than the flow path length of the airfoil 308 (e.g., reduced by 30%). In some examples, the front end 320 of the axial dovetail 310 can be positioned from the leading edge 312 of the airfoil 308, while the rear end 322 is aligned with the trailing edge 314. In these examples, the length of the axial dovetail 310 is also less than the length of the airfoil 308 (e.g., reduced by 30%). The hub point 324 of the leading edge 312 is defined by the radially innermost point of the leading edge 312 (e.g., closest to the centerline axis 102 of the turbine fan 100). Figure 3 In the example, hub point 324 is radially inward of the top edge 326 of the pressure surface 311 of the axial dovetail tenon 310. In other words, the top edge 326 of the axial dovetail tenon 310 is radially outward of hub point 324 of the leading edge 312 (e.g., further away from the centerline axis 102 of the turbine fan 100).

[0034] Figure 4 It shows including Figure 3 The example cross-sectional side view of the inlet portion 400 of a turbine fan, showing a portion of the fan blade 300. For example, the example fan blade 300 disclosed herein can be used with... Figure 1 The turbine fan 100 is implemented together. An example inlet portion 400 includes a fan section 402 that rotates about a centerline axis 102 via a fan shaft (not shown) powered by a low-pressure turbine (not shown). The fan section 402 includes a hub or rotor disk 406. The rotor disk 406 includes a body 408 extending axially from a front surface 410 to a rear surface 412. The body 408 of the example rotor disk 406 includes a plurality of dovetail slots configured to receive axial dovetail tenons 310 of the fan blades 300. Each dovetail slot extends axially from the front surface 410 to the rear surface 412. Therefore, the length of each dovetail slot is the distance between the front surface 410 and the rear surface 412. Furthermore, the length of the body 408 is the distance between the front surface 410 and the rear surface 412.

[0035] Extending radially from the rotor disk 406 are multiple circumferentially spaced fan blades 300. Figure 4 (Only one is shown in the image). The rotor disk 406 and multiple fan blades 300 form the rotor assembly 404. (See image for details.) Figure 4As shown, rotor assembly 404 is part of fan section 402. In other examples, rotor assembly 404 including fan blades 300 is part of a compressor section (e.g., LP compressor 112, HP compressor 114) of turbofan 100. In other examples, rotor assembly 404 including fan blades 300 is part of a turbine section (e.g., HP turbine 118, LP turbine 120) of turbofan 100. Axial dovetail 310 of fan blade 300 is disposed within a dovetail slot (e.g., dovetail slot 504) within rotor disk 406. Forward face 410 of rotor disk 406 is aligned with forward end 320 of axial dovetail 310. Similarly, aft face 412 of rotor disk 406 is aligned with aft end 322 of axial dovetail 310. As a result, the length of the dovetail slot is substantially the same (e.g., within 10%) as the length of axial dovetail 310. Similarly, the length of body 408 is substantially the same (e.g., within 10%) as the length of axial dovetail 310. As described above in connection with Figure 5 , the length of axial dovetail 310 is less than the flow path length of airfoil 308. As a result, the length of body 408 of rotor disk 406 is reduced as compared to the length of the body of known rotor disks (e.g., rotor disk 202). Figure 3 As described above in connection with Figure 2 , the length of axial dovetail 310 is less than the flow path length of airfoil 308. As a result, the length of body 408 of rotor disk 406 is reduced as compared to the length of the body of known rotor disks (e.g., rotor disk 202).

[0036] As described above in connection with Figure 3 , hub point 324 of leading edge 312 is radially inward (e.g., closer to centerline axis 102) of a top edge 326 of pressure face 311 of axial dovetail 310. As a result, the radius of rotor disk 406 can be reduced. The reduced radius of rotor disk 406 can result in a reduced weight as compared to known rotor disks. In addition, the reduced radius can result in a reduced radius ratio of a gas turbine engine implementing Figure 4 As described above in connection with Figure 4 , the length of axial dovetail 310 is less than the flow path length of airfoil 308. As a result, the length of body 408 of rotor disk 406 is reduced as compared to the length of the body of known rotor disks (e.g., rotor disk 202). Figure 4 In examples where the length of axial dovetail 310 is less than the flow path length of airfoil 308, the flow path of air through inlet portion 400 extends radially inward as compared to the flow path of air through inlet portion 144. As a result, given that the blade tip diameter of inlet portion 400 is the same as the blade tip diameter of inlet portion 144, the radius ratio of a gas turbine engine implementing Figure 4 As described above in connection with Figure 2 , the length of axial dovetail 310 is less than the flow path length of airfoil 308. As a result, the length of body 408 of rotor disk 406 is reduced as compared to the length of the body of known rotor disks (e.g., rotor disk 202). Figure 4 In examples where the length of axial dovetail 310 is less than the flow path length of airfoil 308, the flow path of air through inlet portion 400 extends radially inward as compared to the flow path of air through inlet portion 144. As a result, given that the blade tip diameter of inlet portion 400 is the same as the blade tip diameter of inlet portion 144, the radius ratio of a gas turbine engine implementing Figure 2The radius ratio of the gas turbine engine (e.g., turbofan 100) of the inlet portion 144 is reduced. As discussed above, in some examples, the reduced radius ratio can result in the technical effect of increased thrust for a given fan diameter. In other examples, with the reduced radius ratio, fewer blade counts can be achieved while maintaining engine performance, thereby reducing the mass, complexity, cost, and fuel usage of the engine.

[0037] Figure 5 An example inlet portion 400 of a turbofan including a plurality of fan blades 300 is shown. For example, the example inlet portion 400 disclosed herein can be implemented in the turbofan 100 of FIG. 1. Figure 3 and / or Figure 4 An example inlet portion 400 of a turbofan including a plurality of fan blades 300 is shown. For example, the example inlet portion 400 disclosed herein can be implemented in the turbofan 100 of FIG. 1. Figure 1 The example inlet portion 400 includes a spinner 502 and a rotor disk 406. The example rotor disk 406 includes a plurality of dovetail slots 504. Each dovetail slot 504 is configured to receive an axial dovetail 310 of one fan blade 300. In an assembled configuration, each dovetail slot 504 receives an axial dovetail 310 of one fan blade 300. However, some fan blades are omitted from the figure in order to visualize the dovetail slots 504. For the same reason, a portion of the spinner 502 is omitted.

[0038] In some examples, the apparatus includes means for propelling air. For example, the means for propelling air can be implemented by the airfoil 308. In some examples, the apparatus includes means for assembling. For example, the means for assembling can be implemented by the axial dovetail 310. In some examples, the apparatus includes means for combining. For example, the means for combining can be implemented by the rotor disk 202. In some examples, the apparatus includes means for receiving. For example, the means for receiving can be implemented by the dovetail slot 504.

[0039] “Comprise,” “comprising,” “include,” “including,” and “the complement of” (and all forms of these terms) as used herein are used in their open-ended, conventional sense, and are intended to allow for the inclusion of additional elements, terms, etc. As used herein, the phrase “at least” as used, e.g., as the preamble to a claim, is open-ended, in the same manner as the terms “comprising” and “including” are open-ended. The term “and / or” as used in a phrase such as A, B, and / or C, for example, means A alone, B alone, C alone, A with B, A with C, B with C, or A with B and with C. As used herein in the context of describing structures, components, items, objects, and / or things, the phrase “at least one of’ A and B is intended to refer to implementations including any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing structures, components, items, objects, and / or things, the phrase “at least one of’ A or B is intended to refer to implementations including any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. As used herein in the context of describing performance or execution of processes, instructions, actions, activities, and / or steps, the phrase “at least one of’ A and B is intended to refer to implementations including any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing performance or execution of processes, instructions, actions, activities, and / or steps, the phrase “at least one of’ A or B is intended to refer to implementations including any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B.

[0040] As used herein, singular references (e.g., “a,” “an,” “the,” “first,” “second,” etc.) do not exclude the plural. As used herein, the term “a” or “an” object refers to one or more of the object. The terms “a” (or “an”), “one or more,” and “at least one” are used interchangeably herein. Furthermore, although individually listed, a plurality of devices, elements or method actions can be implemented by, e.g., a single entity or object. Additionally, although individual features can be included in different examples or claims, these can possibly be combined, and the inclusion of such features is not meant to preclude from the feasibility of the combination and / or claims.

[0041] From the foregoing, it will be appreciated that example systems, methods, apparatus, and articles of manufacture have been disclosed that provide a fan blade for a gas turbine engine that results in a reduced radius ratio of the gas turbine engine. The reduced axial dovetail length of the fan blades disclosed herein provides a reduced rotor disk diameter, and thus a reduced radius ratio of the gas turbine engine. For a given fan diameter, the reduced radius ratio of the gas turbine engine provides a technical effect of increased performance (e.g., thrust). Additionally or alternatively, the reduced radius ratio can provide a technical effect of reduced gas turbine engine weight while maintaining technical performance. In other examples, the reduced radius ratio provides a technical effect of reduced number of components and / or complexity of the gas turbine engine while maintaining technical performance.

[0042] Example methods, apparatus, systems, and articles of manufacture for implementing low radius ratio fan blades of a gas turbine engine are disclosed herein. Further examples and combinations thereof include the following:

[0043] Example 1 includes a fan blade for a gas turbine engine, comprising: an airfoil comprising a leading edge and a trailing edge extending between a root and a tip of the airfoil, a distance between the leading edge and the trailing edge defining a flow path length, a leading edge hub point defined by a radially innermost point of the leading edge; and an axial dovetail comprising a pair of opposing pressure faces, an axial length of the axial dovetail being less than the flow path length, a top edge of the axial dovetail separating the pressure faces from the airfoil, the top edge of the axial dovetail being radially outward of the leading edge hub point.

[0044] Example 2 includes the fan blade of any preceding clause, wherein the fan blade is formed of a composite material.

[0045] Example 3 includes the fan blade of any preceding clause, wherein the fan blade is formed of a metal or metal alloy.

[0046] Example 4 includes the fan blade of any preceding clause, wherein the axial length of the axial dovetail is from a forward end of the axial dovetail to an aft end of the axial dovetail.

[0047] Example 5 includes the fan blade of any preceding clause, wherein the forward end of the axial dovetail is axially offset from the leading edge of the airfoil.

[0048] Example 6 includes the fan blade of any preceding clause, wherein the aft end of the axial dovetail is axially offset from the trailing edge of the airfoil.

[0049] Example 7 includes a rotor assembly comprising: a rotor disk comprising a body comprising an array of dovetail slots; and an array of fan blades, each fan blade having: an airfoil comprising a leading edge and a trailing edge, a distance between the leading edge and the trailing edge defining a flow path length, a leading edge hub point defined by a radially innermost point of the leading edge; and an axial dovetail comprising a pair of opposing pressure faces, an axial length of the axial dovetail being less than the flow path length, a top edge of the axial dovetail separating the pressure faces from the airfoil, the top edge of the axial dovetail being radially outward of the leading edge hub point.

[0050] Example 8 includes the rotor assembly of any preceding item, wherein the rotor assembly is part of a fan section of a gas turbine engine.

[0051] Example 9 includes the rotor assembly of any preceding item, wherein the rotor assembly is part of a compressor section of a gas turbine engine.

[0052] Example 10 includes the rotor assembly of any preceding item, wherein the rotor assembly is part of a turbine section of a gas turbine engine.

[0053] Example 11 includes the rotor assembly of any preceding item, wherein an axial length of each of the dovetail slots is less than the flow path length.

[0054] Example 12 includes the rotor assembly of any preceding item, wherein an axial length of the rotor assembly is less than the flow path length.

[0055] Example 13 includes the rotor assembly of any preceding item, wherein the axial length of the axial dovetail is from a forward end of the axial dovetail to an aft end of the axial dovetail.

[0056] Example 14 includes the rotor assembly of any preceding item, wherein the forward end of the axial dovetail is axially offset from the leading edge of the airfoil.

[0057] Example 15 includes the rotor assembly of any preceding item, wherein the aft end of the axial dovetail is axially offset from the trailing edge of the airfoil.

[0058] Example 16 includes a gas turbine engine comprising: a compressor; a combustion section; a turbine; a shaft rotatably coupling the compressor and the turbine; and a fan section comprising: a rotor disk comprising an array of dovetail slots; and an array of fan blades, each fan blade having: an airfoil comprising a leading edge and a trailing edge, a distance between the leading edge and the trailing edge defining a flowpath length, a leading edge hub point defined by a radially innermost point of the leading edge; and an axial dovetail comprising a pair of opposing pressure faces, an axial length of the axial dovetail being less than the flowpath length, a top edge of the axial dovetail separating the pressure faces from the airfoil, the top edge of the axial dovetail being radially outward of the leading edge hub point.

[0059] Example 17 includes the gas turbine engine of any preceding paragraph, wherein the array of fan blades is to reduce a radius ratio of the gas turbine engine.

[0060] Example 18 includes the gas turbine engine of any preceding paragraph, wherein the axial length of the axial dovetail is from a forward end of the axial dovetail to an aft end of the axial dovetail.

[0061] Example 19 includes the gas turbine engine of any preceding paragraph, wherein the forward end of the axial dovetail is axially offset in a direction toward the compressor compared to an axial position of the leading edge of the airfoil.

[0062] Example 20 includes the gas turbine engine of any preceding paragraph, wherein the aft end of the axial dovetail is axially offset in a direction away from the compressor compared to an axial position of the trailing edge of the airfoil.

[0063] Although certain example systems, methods, devices, and articles of manufacture have been disclosed herein, the scope of coverage of this patent is not limited thereto. On the contrary, this patent covers all systems, methods, devices, and articles of manufacture falling within the scope of the claims.

[0064] The accompanying claims are hereby incorporated into the detailed description, wherein each claim independently stands as a separate embodiment of this disclosure.

Claims

1. A fan blade for a gas turbine engine, characterized in that, include: An airfoil includes a leading edge and a trailing edge extending between the root and the tip of the airfoil, the distance between the leading edge and the trailing edge defining a flow path length, and the leading edge hub point being defined by the radially innermost point of the leading edge; and An axial dovetail tenon includes a pair of opposing pressure surfaces. The axial length of the axial dovetail tenon is less than the flow path length at the root of the airfoil. The top edge of the axial dovetail tenon separates the pressure surfaces from the airfoil. The top edge of the axial dovetail tenon is radially outside the leading edge hub point.

2. The fan blade according to claim 1, characterized in that, The fan blades are made of a composite material.

3. The fan blade according to claim 1, characterized in that, The fan blades are made of metal or metal alloy.

4. The fan blade according to claim 1, characterized in that, The axial length of the axial dovetail joint is from the front end of the axial dovetail joint to the rear end of the axial dovetail joint.

5. The fan blade according to claim 4, characterized in that, The front end of the axial dovetail tenon is axially offset from the leading edge of the airfoil.

6. The fan blade according to claim 4, characterized in that, The rear end of the axial dovetail tenon is axially offset from the trailing edge of the airfoil.

7. A rotor assembly, characterized in that, include: A rotor disk, the rotor disk comprising a body, the body comprising an array of dovetail grooves; and An array of fan blades, each fan blade having: An airfoil includes a leading edge and a trailing edge extending between the root and the tip of the airfoil, the distance between the leading edge and the trailing edge defining a flow path length, and the leading edge hub point being defined by the radially innermost point of the leading edge; and An axial dovetail tenon includes a pair of opposing pressure surfaces. The axial length of the axial dovetail tenon is less than the flow path length at the root of the airfoil. The top edge of the axial dovetail tenon separates the pressure surfaces from the airfoil. The top edge of the axial dovetail tenon is radially outside the leading edge hub point.

8. The rotor assembly according to claim 7, characterized in that, The rotor assembly is part of the fan section of a gas turbine engine.

9. The rotor assembly according to claim 7, characterized in that, The rotor assembly is part of the compressor section of a gas turbine engine.

10. The rotor assembly according to claim 7, characterized in that, The rotor assembly is part of the turbine section of a gas turbine engine.

11. The rotor assembly according to claim 7, characterized in that, The axial length of each dovetail groove in the dovetail groove is less than the length of the flow path.

12. The rotor assembly according to claim 7, characterized in that, The axial length of the main body of the rotor disk is less than the length of the flow path.

13. The rotor assembly according to claim 7, characterized in that, The axial length of the axial dovetail joint is from the front end of the axial dovetail joint to the rear end of the axial dovetail joint.

14. The rotor assembly according to claim 13, characterized in that, The front end of the axial dovetail tenon is axially offset from the leading edge of the airfoil.

15. The rotor assembly according to claim 13, characterized in that, The rear end of the axial dovetail tenon is axially offset from the trailing edge of the airfoil.

16. A gas turbine engine, characterized in that, include: compressor; Combustion zone; turbine; A shaft rotatably connects the compressor and the turbine; and Fan section, the fan section including: Rotor disk, the rotor disk comprising an array of dovetail grooves; and An array of fan blades, each fan blade having: An airfoil includes a leading edge and a trailing edge extending between the root and the tip of the airfoil, the distance between the leading edge and the trailing edge defining a flow path length, and the leading edge hub point being defined by the radially innermost point of the leading edge; and An axial dovetail tenon includes a pair of opposing pressure surfaces. The axial length of the axial dovetail tenon is less than the flow path length at the root of the airfoil. The top edge of the axial dovetail tenon separates the pressure surfaces from the airfoil. The top edge of the axial dovetail tenon is radially outside the leading edge hub point.

17. The gas turbine engine according to claim 16, characterized in that, The array of fan blades is used to reduce the radius ratio of the gas turbine engine, which is the ratio of the radius of the rotor disk to the radius of the blade tip.

18. The gas turbine engine according to claim 16, characterized in that, The axial length of the axial dovetail joint is from the front end of the axial dovetail joint to the rear end of the axial dovetail joint.

19. The gas turbine engine according to claim 18, characterized in that, The front end of the axial dovetail tenon is axially offset in the direction toward the compressor, compared to the axial position of the leading edge of the airfoil.

20. The gas turbine engine according to claim 18, characterized in that, The rear end of the axial dovetail tenon is axially offset away from the compressor, compared to the axial position of the trailing edge of the airfoil.

Images