Patent Publication Number: US-2023160312-A1

Title: Low radius ratio fan blade for a gas turbine engine

Description:
FIELD OF THE DISCLOSURE 
     This disclosure relates generally to gas turbines and, more particularly, to fan blades for a gas turbine engine. 
     BACKGROUND 
     A gas turbine engine generally includes, 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, thereby creating 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 THE DRAWINGS 
         FIG.  1    illustrates an example gas turbine engine. 
         FIG.  2    illustrates an example cross-sectional side view of an inlet section of the example gas turbine engine. 
         FIG.  3    illustrates an example fan blade. 
         FIG.  4    illustrates an example cross-sectional side view of an inlet section of an example gas turbine engine including a portion of the example fan blade. 
         FIG.  5    illustrates a partial perspective side view of the example inlet section including a plurality of the example fan blade. 
     
    
    
     The figures are not to scale. Instead, the thickness of the layers or regions may be enlarged in the drawings. Although the figures show layers and regions with clean lines and boundaries, some or all of these lines and/or boundaries may be idealized. In reality, the boundaries and/or lines may be unobservable, blended, and/or irregular. In general, the same reference numbers will be used throughout the drawing(s) and accompanying written description to refer to the same or like parts. As used herein, unless otherwise stated, the term “above” describes the relationship of two parts relative to Earth. A first part is above a second part, if the second part has at least one part between Earth and the first part. Likewise, as used herein, a first part is “below” a second part when the first part is closer to the Earth than the second part. As noted above, a first part can be above or below a second part with one or more of: other parts therebetween, without other parts therebetween, with the first and second parts touching, or without the first and second parts being in direct contact with one another. As used in this patent, stating that any part (e.g., a layer, film, area, region, or plate) is in any way on (e.g., positioned on, located on, disposed on, or formed on, etc.) another part, indicates that the referenced part is either in contact with the other part, or that the referenced part is above the other part with one or more intermediate part(s) located therebetween. As used herein, connection references (e.g., attached, coupled, connected, and joined) may include intermediate members between the elements referenced by the connection reference and/or relative movement between those elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and/or in fixed relation to each other. As used herein, stating that any part is in “contact” with another part is defined to mean that there is no intermediate part between the two parts. 
     Unless specifically stated otherwise, descriptors such as “first,” “second,” “third,” etc., are used herein without imputing or otherwise indicating any meaning of priority, physical order, arrangement in a list, and/or ordering in any way, but are merely used as labels and/or arbitrary names to distinguish elements for ease of understanding the disclosed examples. In some examples, the descriptor “first” may be used to refer to an element in the detailed description, while the same element may be referred to in a claim with a different descriptor such as “second” or “third.” In such instances, it should be understood that such descriptors are used merely for identifying those elements distinctly that might, for example, otherwise share a same name. As used herein, “approximately” and “about” refer to dimensions that may not be exact due to manufacturing tolerances and/or other real world imperfections. As used herein “substantially the same size” refers to dimensions that may not be exactly identical due to manufacturing tolerances and/or other real world imperfections. Thus, unless otherwise specified, “substantially the same size” refers to +/−10 percent of a dimension. As used herein, the phrase “in communication,” including variations 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 constant communication, but rather additionally includes selective communication at periodic intervals, scheduled intervals, aperiodic intervals, and/or one-time events. 
     DETAILED DESCRIPTION 
     Many known technologies are directed to increasing gas turbine engine performance. Some known technologies are directed to decreasing the mass of a gas turbine engine while at least maintaining technical specifications and/or performance. For example, some technologies may be directed to increasing performance (e.g., thrust, fuel economy, etc.) of a gas turbine engine while maintaining a given package size (e.g., a diameter of a fan casing). An example technical specification of a gas turbine engine that effects performance is a radius ratio (e.g., the ratio of a radius of a hub of a fan section to the radius of a fan blade tip). It is desirable for the fan to have a smaller radius ratio because a smaller radius ratio increase fan inlet area for a given fan diameter, thus allowing for increased thrust. Examples disclosed herein can provide for a decreased radius ratio by utilizing a fan blade with a shortened blade root in order to lower a leading edge hub point. 
     In the following detailed description, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration specific examples that may be practiced. These examples are described in sufficient detail to enable one skilled in the art to practice the subject matter, and it is to be understood that other examples may be utilized. The following detailed description is, therefore, provided to describe an example implementation and not to be taken limiting on the scope of the subject matter described in this disclosure. Certain features from different aspects of the following description may be combined to form yet new aspects of the subject matter discussed below. 
     The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. As used herein, “vertical” refers to the direction perpendicular to the ground. As used herein, “horizontal” refers to the direction parallel to the centerline of the turbofan  100 . As used herein, “lateral” refers to the direction perpendicular to the axial vertical directions (e.g., into and out of the plane of  FIGS.  1 ,  2   , etc.). 
     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 direction, radial direction, and circumferential direction of the vehicle associated with the features, forces and moments. In general, the attached figures are annotated with a set of axes including the axial axis A, the radial axis R, and the circumferential axis C. Additionally or alternatively, the attached figures are annotated with a set of axes including the roll axis R, the pitch axis P, and the yaw axis Y. 
     “Including” and “comprising” (and all forms and tenses thereof) are used herein to be open ended terms. Thus, whenever a claim employs any form of “include” or “comprise” (e.g., comprises, includes, comprising, including, having, etc.) as a preamble or within a claim recitation of any kind, it is to be understood that additional elements, terms, etc. may be present without falling outside the scope of the corresponding claim or recitation. As used herein, when the phrase “at least” is used as the transition term in, for example, a preamble of a claim, it is open-ended in the same manner as the term “comprising” and “including” are open ended. The term “and/or” when used, for example, in a form such as A, B, and/or C refers to any combination or subset of A, B, C such as (1) A alone, (2) B alone, (3) C alone, (4) A with B, (5) A with C, (6) B with C, and (7) 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 the 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 the 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. 
     As used herein, singular references (e.g., “a”, “an”, “first”, “second”, etc.) do not exclude a plurality. The term “a” or “an” entity, as used herein, refers to one or more of that entity. The terms “a” (or “an”), “one or more”, and “at least one” can be used interchangeably herein. Furthermore, although individually listed, a plurality of means, elements or method actions may be implemented by, e.g., a single unit or processor. Additionally, although individual features may be included in different examples or claims, these may 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. 
     Gas turbine engines include a fan section proximate an intake of the engine. The fan section includes a plurality of circumferentially spaced apart fan blades. Each of the fan blades extends radially outward from a rotor disk or hub. Example fan blades include an airfoil and an integral dovetail at the 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 chordwise 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, such as titanium, or a composite material, such as a carbon-epoxy composite system. 
     In known fan blades, a length of the dovetail is approximately the same (e.g., within 10%) the width of the airfoil. A size (e.g., length, cross-sectional area, etc.) of the dovetail is proportionate to the force needed to retain the fan blade within the dovetail slot. The force needed to retain the fan blade within the dovetail slot is proportionate to parameters of the fan section and/or fan blade such as tip speed of the fan blade and weight of the fan blade. In some examples, the force needed to retain the fan blade is decreased. For example, the fan section of the gas turbine engine may have a reduced speed relative to a compressor or combustion section of the gas turbine engine due to implementation of a geared fan. In this example, the tip speed of the fan blade is reduced, thus reducing the force needed to retain the fan blade. In other examples, the fan blade may be reduced in weight by use of advanced materials (e.g., composite systems), thus reducing the force needed 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. 
     An inlet area of the fan is defined by the space between the rotor disk and the fan blade tips. As used herein, a radius ratio is a ratio of a radius of the rotor disk to a radius of the blade tips. As described above, decreasing the radius ratio allows for increased fan inlet area for a given fan diameter. In some examples, if the radius ratio is decreased, thrust can be increased for a given fan diameter. In other examples, if the radius ratio is decreased, a lower blade count can be implemented while maintaining engine performance, thus decreasing mass, complexity, cost, and fuel utilization of the engine. Examples disclosed herein reduce a 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 a radius ratio of a gas turbine engine fan by implementing a fan blade with a leading edge hub point radially inward of an attachment pressure face. 
     Reference now will be made in detail to 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, not limitation of 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 present 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. 
       FIG.  1    is a schematic cross-sectional view of a prior art turbofan-type gas turbine engine  100  (“turbofan  100 ”). As shown in  FIG.  1   , the turbofan  100  defines a longitudinal or axial centerline axis  102  extending therethrough for reference. In general, the turbofan  100  may include a core turbine  104  or gas turbine engine disposed downstream from a fan section  106 . 
     The core turbine  104  generally includes a substantially tubular outer casing  108  (“turbine casing  108 ”) that defines an annular inlet  110 . The outer casing  108  can be formed from 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  may also couple to a fan spool or shaft  128  of the fan section  106  (“fan shaft  128 ”). In some examples, the LP shaft  126  may couple directly to the fan shaft  128  (i.e., a direct-drive configuration). In alternative configurations, the LP shaft  126  may couple to the fan shaft  128  via a reduction gearbox  130  (e.g., an indirect-drive or geared-drive configuration). 
     As shown in  FIG.  1   , the fan section  106  includes a plurality of fan blades  132  coupled to and extending radially outwardly from the fan shaft  128 . An annular fan casing or nacelle  134  circumferentially encloses the fan section  106  and/or at least a portion of the core turbine  104 . The nacelle  134  is supported relative to the core turbine  104  by a plurality of circumferentially-spaced apart outlet guide vanes  136 . Furthermore, a downstream section  138  of the nacelle  134  can enclose an outer portion of the core turbine  104  to define a bypass airflow passage  140  therebetween. 
     As illustrated in  FIG.  1   , air  142  enters an inlet portion  144  of the turbofan  100  during operation thereof. 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 sequential stages of LP compressor stator vanes  150  and LP compressor 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  en route to the HP compressor  114 . Next, one or more sequential stages of HP compressor stator vanes  154  and HP compressor 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 . 
     The combustion gases  160  flow through the HP turbine  118  in which one or more sequential stages of HP turbine stator vanes  162  and HP turbine rotor blades  164  coupled to the HP shaft  124  extract a first portion of kinetic and/or thermal energy from the combustion gases  160 . This energy extraction supports operation of the HP compressor  114 . The combustion gases  160  then flow through the LP turbine  120  where one or more sequential stages of LP turbine stator vanes  166  and LP turbine rotor blades  168  coupled to the LP shaft  126  extract a second portion of thermal and/or kinetic energy therefrom. This energy extraction causes the LP shaft  126  to rotate, thereby supporting operation of the LP compressor  112  and/or rotation of the fan shaft  128 . The combustion gases  160  then exit the core turbine  104  through the exhaust section  122  thereof. 
     Along with the turbofan  100 , the core turbine  104  serves a similar purpose and sees a similar environment in land-based gas turbines, turbojet engines in which the ratio of the first portion  146  of the air  142  to the second portion  148  of the air  142  is less than that of a turbofan (e.g., turbofan  100 ), and unducted fan engines in which the fan section  106  is devoid of the nacelle  134 . In each of the turbofan, turbojet, and unducted engines, a speed reduction device (e.g., the reduction gearbox  130 ) may be included between any shafts and spools. For example, the reduction gearbox  130  may be disposed between the LP shaft  126  and the fan shaft  128  of the fan section  106 . 
       FIG.  2    illustrates an example inlet portion  144  that can be implemented in the example turbofan  100  shown in  FIG.  1   . The example inlet portion  144  includes the fan section  106  which is rotated about the centerline axis  102  by the fan shaft  128  powered by the LP turbine  120  (not shown). The fan section  106  includes a rotor disk  202  from which extends radially outwardly a plurality of circumferentially spaced apart fan or rotor blades  132  (only one shown in  FIG.  2   ). The fan blades  132  may be metallic or nonmetallic. For example, the fan blades  132  may be made from a carbon fiber-epoxy composite or other similar material. The rotor disk  202  includes a forward end  204  axially spaced from an aft end  206 , and a radially outer surface  208  extending therebetween. 
     Disposed downstream of the fan section  106  is the LP compressor  112  having axially spaced apart vane and blade rows, with the blades thereof being joined to the LP shaft  126 . In the illustrated example, the LP shaft  126  is fixedly joined to the aft end  206  of the rotor disk  202  by a plurality of bolts  210 . A spinner  212  is joined to the forward end  204  of the rotor disk  202  to provide an aerodynamic flow path for the air  142  entering the fan section  106 . 
       FIG.  3    illustrates an example fan blade  300  that can be implemented in the inlet portion  144  of the turbofan  100  shown in  FIGS.  1  and/or  2   . The example fan blade  300  has a pressure side  302  and opposed suction side  304 . The example fan blade  300  is an integral component including a root section  306  and an airfoil  308 . The root section  306  includes a straight axial dovetail  310  with a pair of opposed pressure faces  311 . When the fan blade  300  is assembled in a rotor disk (e.g., the rotor disk  202  depicted in  FIG.  2   ), the axial dovetail  310  is disposed in a corresponding dovetail slot in the rotor disk (as shown, for example, in  FIG.  5   ). While the example fan blade  300  of  FIG.  3    includes the axial dovetail  310 , other example fan blades can include a dovetail with a different geometry such as skewed axial, circular arc, etc. The axial dovetail  310  of  FIG.  3    can represent a dovetail of any such geometry (e.g., skewed axial, circular arc, etc.). 
     Still referring to  FIG.  3   , the airfoil  308  of example fan blade  300  extends in a chordwise direction (e.g., along 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 . The example fan blade  300  of  FIG.  3    may be constructed from a variety of materials including metal, metal alloys, nonmetallic composites, and combinations thereof. In the illustrated example, the fan blade  300  is constructed from a composite material. The term “composite” refers generally to a material containing a reinforcement such as fibers or particles supported in a binder or matrix material. The composite may include a number of layers or plies embedded in a matrix and oriented substantially parallel to the pressure and suction sides  302 ,  304 . An example of a suitable material is a carbonaceous (e.g., graphite) fiber embedded in a resin material such as epoxy. These are commercially available as fibers unidirectionally aligned into a tape that is impregnated with a resin. Such “prepreg” tape can be formed into a part shape, and cured via an autoclaving process or press molding to form a light weight, stuff, relatively homogeneous article. 
     As discussed above, during operation of a gas turbine engine (e.g., the turbofan  100 ), air (e.g., the air  142 ) enters an inlet portion (e.g., the inlet portion  144  of  FIGS.  1  and/or  2   ) and flows substantially axially towards compressor, combustor, and/or turbine sections. As such, the air flows across rotor, stator, and/or fan blades in the gas turbine engine. For example, the air can flow across the example fan blade  300  when the fan blade  300  is mounted to a rotor disk (e.g., rotor disk  202  depicted in  FIG.  2   ). As used herein, a flow path length is a distance the air flows across the fan blade  300 . In the example of  FIG.  3   , the flow path length of the fan blade  300  is defined by a distance between the leading edge  312  and the trailing edge  314 . 
     In the example of  FIG.  3   , the axial dovetail  310  includes a fore end  320  and an aft end  322 . A length of the axial dovetail  310  is defined by a distance between the fore end  320  and the aft end  322 . In known fan blades such as the fan blade  132  of  FIG.  2   , a fore end of an axial dovetail is aligned with a leading edge of an airfoil. In the example of  FIG.  3   , the fore end  320  of the axial dovetail  310  is disposed in a chordwise direction from the leading edge  312  toward the trailing edge  314 . Similarly, in known fan blades such as the fan blade  132  of  FIG.  2   , an aft end of an axial dovetail is aligned with a trailing edge of an airfoil. In the example of  FIG.  3   , the aft end  322  of the axial dovetail  310  is disposed in a chordwise direction from the trailing edge  314  toward the leading edge  312 . As a result, in the example of  FIG.  3   , the length of the axial dovetail  310  is less than (e.g., reduced by 30 percent) the flow path length of the airfoil  308 . In some examples, the fore end  320  of the axial dovetail  310  can be disposed from the leading edge  312  of the airfoil  308  while the aft end  322  is aligned with the trailing edge  314 . In these examples, the length of the axial dovetail  310  is also less than (e.g., reduced by 30 percent) the length of the airfoil  308 . A hub point  324  of the leading edge  312  is defined by a radially innermost (e.g., closest to the centerline axis  102  of the turbofan  100 ) point of the leading edge  312 . In the example of  FIG.  3   , the hub point  324  is radially inward of a top edge  326  of the pressure face  311  of the axial dovetail  310 . In other words, the top edge  326  of the axial dovetail  310  is radially outward (e.g., further from the centerline axis  102  of the turbofan  100 ) compared to the hub point  324  of the leading edge  312 . 
       FIG.  4    illustrates an example cross-sectional side view of an inlet portion  400  of a turbofan including a portion of the fan blade  300  shown in  FIG.  3   . The example fan blade  300  disclosed herein can be implemented with the turbofan  100  of  FIG.  1   , for example. The example inlet portion  400  includes a fan section  402  which is rotated about the centerline axis  102  by 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 forward surface  410  to an aft surface  412 . The body  408  of the example rotor disk  406  includes a plurality of dovetail slots configured to receive the axial dovetails  310  of the fan blades  300 . Each of the dovetail slots extend axially from the forward surface  410  to the aft surface  412 . Thus, a length of each of the dovetail slots is a distance between the forward surface  410  and the aft surface  412 . Additionally, a length of the body  408  is the distance between the forward surface  410  and the aft surface  412 . 
     Extending radially from the rotor disk  406  is a plurality of the fan blades  300  (only one shown in  FIG.  4   ) circumferentially spaced apart. The rotor disk  406  and the plurality of fan blades  300  form a rotor assembly  404 . As illustrated in  FIG.  4   , the rotor assembly  404  is a portion of the fan section  402 . In other examples, the rotor assembly  404  including the fan blades  300  is a portion of a compressor section (e.g., the LP compressor  112 , the HP compressor  114 ) of the turbofan  100 . In other examples, the rotor assembly  404  including the fan blades  300  is a portion of a turbine section (e.g., the HP turbine  118 , the LP turbine  120 ) of the turbofan  100 . The axial dovetail  310  of the fan blade  300  is disposed within the dovetail slot (e.g., dovetail slot  504  of  FIG.  5   ) within the rotor disk  406 . The forward surface  410  of the rotor disk  406  is aligned with the fore end  320  of the axial dovetail  310 . Similarly, the aft surface  412  of the rotor disk  406  is aligned with the aft end  322  of the axial dovetail  310 . As a result, the length of the dovetail slot is substantially (e.g., within 10 percent) the same as the length of the axial dovetail  310 . Similarly, the length of the body  408  is substantially (e.g., within 10 percent) the same as the length of the axial dovetail  310 . As described above in connection with  FIG.  3   , the length of the axial dovetail  310  is less than the flowpath length of the airfoil  308 . As a result, the length of the body  408  of the rotor disk  406  is reduced (e.g., 30 percent less) compared to a length of a body of a known rotor disk (e.g., the rotor disk  202  of  FIG.  2   ). As a result of the reduced length, the rotor disk  406  can have a reduced weight (e.g., 30 percent less) compared to a known rotor disk. 
     As described in connection with  FIG.  3   , the hub point  324  of the leading edge  312  is radially inward (e.g., closer to the centerline axis  102 ) of the top edge  326  of the pressure face  311  of the axial dovetail  310 . As a result, a radius of the rotor disk  406  can be reduced. The reduced radius of the rotor disk  406  can result in a reduced weight compared to a known rotor disk. Additionally, the reduced radius can result in a reduced radius ratio for a gas turbine engine implementing the inlet portion  400  of  FIG.  4   . For example, a radially inner boundary of an air flowpath for the inlet portion  400  is illustrated by line  416 . A radially inner boundary of an air flowpath for an inlet portion (e.g., the inlet portion  144 ) with known fan blades (e.g., the fan blades  132 ) is illustrated in  FIG.  4    by line  418 . In the example of  FIG.  4   , the flowpath for the air through the inlet portion  400  extends radially inner compared to the flowpath for the air though the inlet portion  144 . As a result, given the same blade tip diameter for the inlet portion  400  of  FIG.  4    and the inlet portion  144  of  FIG.  2   , a radius ratio of a gas turbine engine implementing the inlet portion  400  of  FIG.  4    is reduced from a radius ratio of a gas turbine engine (e.g., the turbofan  100 ) implementing the inlet portion  144  of  FIG.  2   . As discussed above, in some examples, a reduced radius ratio can result in a technical effect of increased thrust for a given fan diameter. In other examples, with a reduced radius ratio, a lower blade count can be implemented while maintaining engine performance, thus decreasing mass, complexity, cost, and fuel utilization of the engine. 
       FIG.  5    illustrates the example inlet portion  400  of a turbofan including a plurality of the fan blades  300  shown in  FIGS.  3  and/or  4   . The example inlet portion  400  disclosed herein can be implemented with the turbofan  100  of  FIG.  1   , for example. The example inlet portion  400  includes a spinner  502  and the rotor disk  406 . The example rotor disk  406  includes a plurality of dovetail slots  504 . Each of the dovetail slots  504  is configured to receive the axial dovetail  310  of one of the fan blades  300 . In an assembled configuration, each of the dovetail slots  504  receives an axial dovetail  310  of one of the fan blades  300 . However, some of the fan blades are omitted from the illustration for the purposes of visualizing the dovetail slots  504 . For the same reason, a portion of the spinner  502  is omitted. 
     In some examples, the apparatus includes means for pushing air. For example, the means for pushing air may be implemented by airfoil  308 . In some examples, the apparatus includes means for assembling. For example, the means for assembling may be implemented by axial dovetail  310 . In some examples, the apparatus includes means for combining. For example, the means for combining may be implemented by rotor disk  202 . In some examples, the apparatus includes means for receiving. For example, the means for receiving may be implemented by dovetail slots  504 . 
     “Including” and “comprising” (and all forms and tenses thereof) are used herein to be open ended terms. Thus, whenever a claim employs any form of “include” or “comprise” (e.g., comprises, includes, comprising, including, having, etc.) as a preamble or within a claim recitation of any kind, it is to be understood that additional elements, terms, etc., may be present without falling outside the scope of the corresponding claim or recitation. As used herein, when the phrase “at least” is used as the transition term in, for example, a preamble of a claim, it is open-ended in the same manner as the term “comprising” and “including” are open ended. The term “and/or” when used, for example, in a form such as A, B, and/or C refers to any combination or subset of A, B, C such as (1) A alone, (2) B alone, (3) C alone, (4) A with B, (5) A with C, (6) B with C, or (7) 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, or (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, or (3) at least one A and at least one B. As used herein in the context of describing the 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, or (3) at least one A and at least one B. Similarly, as used herein in the context of describing the 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, or (3) at least one A and at least one B. 
     As used herein, singular references (e.g., “a”, “an”, “first”, “second”, etc.) do not exclude a plurality. The term “a” or “an” object, as used herein, refers to one or more of that object. 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 may be implemented by, e.g., the same entity or object. Additionally, although individual features may be included in different examples or claims, these may 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. 
     From the foregoing, it will be appreciated that example systems, methods, apparatus, and articles of manufacture have been disclosed that provide for fan blades for gas turbine engines which result in a reduced radius ratio of the gas turbine engine. The reduced axial dovetail length of the fan blades disclosed herein provide for a reduced rotor disk diameter and, thus, a reduced radius ratio of the gas turbine engine. The reduced radius ratio of the gas turbine engine provides a technical effect of increased performance (e.g., thrust) for a given fan diameter. Additionally or alternatively, the reduced radius ratio can provide a technical effect of reduced weight of the gas turbine engine while maintaining technical performance. In other examples, the reduced radius ratio provides a technical effect of reduced part count and/or complexity of the gas turbine engine while maintaining technical performance. 
     Example methods, apparatus, systems, and articles of manufacture to implement a low radius ratio fan blade for a gas turbine engine are disclosed herein. Further examples and combinations thereof include the following: 
     Example 1 includes a fan blade for a gas turbine engine, comprising 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 opposed pressure faces, an axial length of the axial dovetail 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 radially outward of the leading edge hub point. 
     Example 2 includes the fan blade of any preceding clause, wherein the fan blade is formed from a composite material. 
     Example 3 includes the fan blade of any preceding clause, wherein the fan blade is formed from a metal or metal alloy. 
     Example 4 includes the fan blade of any preceding clause, wherein the axial length of the axial dovetail is from a fore end of the axial dovetail to an aft end of the axial dovetail. 
     Example 5 includes the fan blade of any preceding clause, wherein the fore end of the axial dovetail is offset axially from the leading edge of the airfoil. 
     Example 6 includes the fan blade of any preceding clause, wherein the aft end of the axial dovetail is offset axially from the trailing edge of the airfoil. 
     Example 7 includes a rotor assembly, comprising a rotor disk including a body including an array of dovetail slots; and an array of fan blades, each having an airfoil including 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 including a pair of opposed pressure faces, an axial length of the axial dovetail 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 radially outward of the leading edge hub point. 
     Example 8 includes the rotor assembly of any preceding clause, wherein the rotor assembly is a portion of a fan section of a gas turbine engine. 
     Example 9 includes the rotor assembly of any preceding clause, wherein the rotor assembly is a portion of a compressor section of a gas turbine engine. 
     Example 10 includes the rotor assembly of any preceding clause, wherein the rotor assembly is a portion of a turbine section of a gas turbine engine. 
     Example 11 includes the rotor assembly of any preceding clause, wherein an axial length of each of the dovetail slots is less than the flow path length. 
     Example 12 includes the rotor assembly of any preceding clause, wherein an axial length of the rotor assembly is less than the flow path length. 
     Example 13 includes the rotor assembly of any preceding clause, wherein the axial length of the axial dovetail is from a fore end of the axial dovetail to an aft end of the axial dovetail. 
     Example 14 includes the rotor assembly of any preceding clause, wherein the fore end of the axial dovetail is offset axially from the leading edge of the airfoil. 
     Example 15 includes the rotor assembly of any preceding clause, wherein the aft end of the axial dovetail is offset axially from the trailing edge of the airfoil. 
     Example 16 includes a gas turbine engine, comprising a compressor; a combustion section; a turbine; a shaft to rotatably couple the compressor and the turbine; and a fan section, the fan section including a rotor disk including an array of dovetail slots; and an array of fan blades, each having an airfoil including 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 including a pair of opposed pressure faces, an axial length of the axial dovetail 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 radially outward of the leading edge hub point. 
     Example 17 includes the gas turbine engine of any preceding clause, wherein the array of fan blades are to reduce a radius ratio of the gas turbine engine. 
     Example 18 includes the gas turbine engine of any preceding clause, wherein the axial length of the axial dovetail is from a fore end of the axial dovetail to an aft end of the axial dovetail. 
     Example 19 includes the gas turbine engine of any preceding clause, wherein the fore end of the axial dovetail is offset axially in a direction toward the compressor compared to an axial location of the leading edge of the airfoil. 
     Example 20 includes the gas turbine engine of any preceding clause, wherein the aft end of the axial dovetail is offset axially in a direction away from the compressor compared to an axial location of the trailing edge of the airfoil. 
     Although certain example systems, methods, apparatus, 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, apparatus, and articles of manufacture fairly falling within the scope of the claims of this patent. 
     The following claims are hereby incorporated into this Detailed Description by this reference, with each claim standing on its own as a separate embodiment of the present disclosure.