Patent Publication Number: US-2020300175-A1

Title: Gearbox for a turbomachine with alternatingly spaced rotor blades

Description:
PRIORITY INFORMATION 
     The present application claims priority to Italian Patent Application Number 102019000003991 filed on 19 Mar. 2019. 
     FIELD 
     The present subject matter relates generally to a turbomachine, and more particularly, to a gearbox for a turbomachine having alternatingly spaced rotor blades. 
     BACKGROUND 
     Gas turbine engines generally include a turbine section downstream of a combustion section that is rotatable with a compressor section to rotate and operate the gas turbine engine to generate power, such as propulsive thrust. General gas turbine engine design criteria often include conflicting criteria that must be balanced or compromised, including increasing fuel efficiency, operational efficiency, and/or power output while maintaining or reducing weight, part count, and/or packaging (i.e. axial and/or radial dimensions of the engine). 
     Within at least certain gas turbine engines, the turbine section may include interdigitated rotors (i.e., successive rows or stages of rotating airfoils or blades). For example, a turbine section may include a turbine having a first plurality of low speed turbine rotor blades and a second plurality of high speed turbine rotor blades rotatable with one another through a reversing gearbox. The first plurality of low speed turbine rotor blades may be interdigitated with the second plurality of high speed turbine rotor blades. Such a configuration may result in a more efficient turbine. 
     However, several problems may arise with such a configuration relating to clearance issues between the first and second pluralities of rotor blades and packaging of the gearbox inward of the turbine. Accordingly, an improved turbine with interdigitated turbine rotor blades would be useful. 
     BRIEF DESCRIPTION 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     In one exemplary embodiment of the present disclosure a gas turbine engine defining a radial direction and an axial direction is provided. The gas turbine engine includes a stationary frame, a compressor and a turbine, the compressor or the turbine including a first plurality of rotor blades and a second plurality of rotor blades, the first plurality of rotor blades and second plurality of rotor blades alternatingly spaced along the axial direction and rotatable with one another; and a gearbox including a first gear coupled to the first plurality of rotor blades, a second gear coupled to the second plurality of rotor blades, and an intermediate gear positioned between the first gear and the second gear and coupled to the stationary frame, the intermediate gear defining an axis of rotation, the axis of rotation defining an angle with the radial direction less than about 75 degrees. 
     In certain exemplary embodiments the angle the axis of rotation defines with the radial direction is less than about 30 degrees. 
     In certain exemplary embodiments the angle the axis of rotation defines with the radial direction is less than about 15 degrees. 
     In certain exemplary embodiments the angle the axis of rotation defines with the radial direction is about zero degrees. 
     In certain exemplary embodiments the intermediate gear is a first intermediate gear of a plurality of intermediate gears spaced along a circumferential direction of the gas turbine engine within the gearbox. 
     In certain exemplary embodiments the intermediate gear defines a forward end and an aft end, wherein the first gear meshes with the intermediate gear at the forward end and wherein the second gear meshes with the intermediate gear at the aft end. 
     For example, in certain exemplary embodiments the first gear and the intermediate gear together define a first intersection line, wherein the second gear and the intermediate gear together define a second intersection line, wherein the first intersection line defines a first intersection angle less than about 75 degrees with the radial direction, and wherein the second intersection line defines a second intersection angle less than about 75 degrees with the radial direction. 
     For example, in certain exemplary embodiments the first intersection angle is less than about 30 degrees with the radial direction, and wherein the second intersection angle is less than about 30 degrees with the radial direction. 
     For example, in certain exemplary embodiments the first intersection angle, the second intersection angle, and the angle the axis of rotation defines with the radial direction are each equal to about zero degrees. 
     For example, in certain exemplary embodiments the first intersection angle and the second intersection angle are each greater than the angle the axis of rotation defines with the radial direction. 
     In certain exemplary embodiments the turbine includes the first plurality of rotor blades and the second plurality of rotor blades, and wherein the stationary frame is a turbine frame. 
     For example, in certain exemplary embodiments the turbine is a low pressure turbine. 
     In certain exemplary embodiments the gas turbine engine in an aeronautical gas turbine engine. 
     In certain exemplary embodiments the intermediate gear is a compound gear including an outer gear and an inner gear rotatable with one another, wherein one of the first gear or second gear meshes with the outer gear of the intermediate gear, and wherein the other of the first gear or second gear meshes with the inner gear of the intermediate gear. 
     For example, in certain exemplary embodiments the outer gear defines a first diameter, wherein the inner gear defines an second diameter, wherein the first diameter is not equal to the second diameter. 
     A gearbox for a gas turbine engine defining a radial direction and an axial direction, the gas turbine engine including a stationary frame, a compressor, and a turbine, the compressor or the turbine comprising a first plurality of rotor blades and a second plurality of rotor blades alternatingly spaced, the gearbox including a first gear configured to be coupled to the first plurality of rotor blades of the gas turbine engine; a second gear configured to be coupled to the second plurality of rotor blades of the gas turbine engine; and an intermediate gear configured to be coupled to the stationary frame of the gas turbine engine, the intermediate gear defining a forward end and an aft end, the first gear meshing with the intermediate gear at the forward end, the second gear meshing with the intermediate gear at the aft end. 
     In certain exemplary embodiments the intermediate gear further defines an axis of rotation defining an angle with the radial direction less than about 75 degrees. 
     For example, in certain exemplary embodiments the angle the axis of rotation defines with the radial direction is less than about 15 degrees. 
     In certain exemplary embodiments the first gear and the intermediate gear together define a first intersection line, wherein the second gear and the intermediate gear together define a second intersection line, wherein the first intersection line defines a first intersection angle less than about 75 degrees with the radial direction, and wherein the second intersection line defines a second intersection angle less than about 75 degrees with the radial direction 
     In certain exemplary embodiments the intermediate gear is a compound gear including an outer gear and an inner gear rotatable with one another, wherein one of the first gear or second gear meshes with the outer gear of the intermediate gear, and wherein the other of the first gear or second gear meshes with the inner gear of the intermediate gear. 
     These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: 
         FIG. 1  is a schematic cross sectional view of an exemplary gas turbine engine incorporating an exemplary embodiment of a turbine section according to an aspect of the present disclosure; 
         FIG. 2  is a schematic, cross sectional view of a turbine section in accordance with an exemplary aspect of the present disclosure; 
         FIG. 3  is a cross sectional view depicting exemplary blade pitch angles of a turbine of a turbine section in accordance with an exemplary embodiment of the present disclosure; 
         FIG. 4  is a perspective, cross-sectional view of a gearbox in accordance with an exemplary embodiment of the present disclosure. 
         FIG. 5  is a schematic, cross sectional view of a turbine section in accordance with another exemplary aspect of the present disclosure; 
         FIG. 6  is a schematic, cross sectional view of a turbine section in accordance with yet another exemplary aspect of the present disclosure; 
         FIG. 7  is a schematic, cross sectional view of a turbine section in accordance with still another exemplary aspect of the present disclosure; and 
         FIG. 8  is a schematic, cross sectional view of a turbine section in accordance with yet another exemplary aspect of the present disclosure. 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention. 
     DETAILED DESCRIPTION 
     Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. 
     As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. 
     The terms “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle, and refer to the normal operational attitude of the gas turbine engine or vehicle. For example, with regard to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust. 
     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. 
     The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein. 
     The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise. 
     The terms “low speed” and “high-speed” refer to relative speeds, such as relative rotational speeds, of two components during operations of the turbomachine, and do not imply or require any minimum or maximum absolute speeds. 
     Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 10 percent margin. 
     Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. 
     Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,  FIG. 1  is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of  FIG. 1 , the gas turbine engine is a high-bypass turbofan jet engine  10 , referred to herein as “turbofan engine  10 .” As shown in  FIG. 1 , the turbofan engine  10  defines an axial direction A (extending parallel to a longitudinal centerline  12  provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not depicted). In general, the turbofan  10  includes a fan section  14  and a core turbine engine  16  disposed downstream from the fan section  14 . 
     The exemplary core turbine engine  16  depicted generally includes a substantially tubular outer casing  18  that defines an annular inlet  20 . The outer casing  18  encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor  22  and a high pressure (HP) compressor  24 ; a combustion section  26 ; a turbine section including a high pressure (HP) turbine  28  and a low pressure (LP) turbine  30 ; and a jet exhaust nozzle section  32 . The compressor section, combustion section  26 , and turbine section together define a core air flowpath  37  extending from the annular inlet  20  through the LP compressor  22 , HP compressor  24 , combustion section  26 , HP turbine section  28 , LP turbine section  30  and jet nozzle exhaust section  32 . A high pressure (HP) shaft or spool  34  drivingly connects the HP turbine  28  to the HP compressor  24 . A low pressure (LP) shaft or spool  36  drivingly connects the LP turbine  30  to the LP compressor  22 . 
     For the embodiment depicted, the fan section  14  includes a fan  38  having a plurality of fan blades  40  coupled to a disk  42  in a spaced apart manner. As depicted, the fan blades  40  extend outwardly from disk  42  generally along the radial direction R. The fan blades  40  and disk  42  are together rotatable about the longitudinal axis  12  by LP shaft  36 . 
     Referring still to the exemplary embodiment of  FIG. 1 , the disk  42  is covered by rotatable spinner cone  48  aerodynamically contoured to promote an airflow through the plurality of fan blades  40 . Additionally, the exemplary fan section  14  includes an annular fan casing or outer nacelle  50  that circumferentially surrounds the fan  38  and/or at least a portion of the core turbine engine  16 . It should be appreciated that for the embodiment depicted, the nacelle  50  is supported relative to the core turbine engine  16  by a plurality of circumferentially-spaced outlet guide vanes  52 . Moreover, a downstream section  54  of the nacelle  50  extends over an outer portion of the core turbine engine  16  so as to define a bypass airflow passage  56  therebetween. 
     During operation of the turbofan engine  10 , a volume of air  58  enters the turbofan  10  through an associated inlet  60  of the nacelle  50  and/or fan section  14 . As the volume of air  58  passes across the fan blades  40 , a first portion of the air  58  as indicated by arrows  62  is directed or routed into the bypass airflow passage  56  and a second portion of the air  58  as indicated by arrow  64  is directed or routed into the LP compressor  22 . The ratio between the first portion of air  62  and the second portion of air  64  is commonly known as a bypass ratio. The pressure of the second portion of air  64  is then increased as it is routed through the high pressure (HP) compressor  24  and into the combustion section  26 , where it is mixed with fuel and burned to provide combustion gases  66 . 
     The combustion gases  66  are routed through the HP turbine  28  where a portion of thermal and/or kinetic energy from the combustion gases  66  is extracted via sequential stages of HP turbine stator vanes  68  that are coupled to an inner casing (not shown) and HP turbine rotor blades  70  that are coupled to the HP shaft or spool  34 , thus causing the HP shaft or spool  34  to rotate, thereby supporting operation of the HP compressor  24 . The combustion gases  66  are then routed through the LP turbine  30  where a second portion of thermal and kinetic energy is extracted from the combustion gases  66  via sequential stages of a first plurality of LP turbine rotor blades  72  that are coupled to an outer drum  73 , and a second plurality of LP turbine rotor blades  74  that are coupled to an inner drum  75 . The first plurality of LP turbine rotor blades  72  and second plurality of LP turbine rotor blades  74  are alternatingly spaced and rotatable with one another through a gearbox  76  to together drive the LP shaft or spool  36 , thus causing the LP shaft or spool  36  to rotate. Such thereby supports operation of the LP compressor  22  and/or rotation of the fan  38 . 
     As will explained in greater detail, below, the gearbox  76  is configured to occupy a smaller radial footprint to, e.g., allow for a smaller diameter turbine, such as a smaller diameter LP turbine  30 . 
     The combustion gases  66  are subsequently routed through the jet exhaust nozzle section  32  of the core turbine engine  16  to provide propulsive thrust. Simultaneously, the pressure of the first portion of air  62  is substantially increased as the first portion of air  62  is routed through the bypass airflow passage  56  before it is exhausted from a fan nozzle exhaust section  78  of the turbofan  10 , also providing propulsive thrust. The HP turbine  28 , the LP turbine  30 , and the jet exhaust nozzle section  32  at least partially define a hot gas path  80  for routing the combustion gases  66  through the core turbine engine  16 . 
     It should be appreciated, however, that the exemplary turbofan engine  10  depicted in  FIG. 1  is by way of example only, and that in other exemplary embodiments, the turbofan engine  10  may have any other suitable configuration. For example, in other exemplary embodiments, the turbofan engine  10  may instead be configured as any other suitable turbomachine including, e.g., any other suitable number of shafts or spools, and excluding, e.g., the fan  38  and/or including, e.g., a gearbox between the fan  38  and the LP shaft or spool  36 , a variable pitch fan  38 , etc. Accordingly, it will be appreciated that in other exemplary embodiments, the turbofan engine  10  may instead be configured as, e.g., a turbojet engine, a turboshaft engine, a turboprop engine, etc., and further may be configured as an aeroderivative gas turbine engine or industrial gas turbine engine. 
     Referring now to  FIG. 2 , a schematic, side, cross-sectional view is provided of a turbine section  100  of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. The exemplary turbine section  100  depicted in  FIG. 2  may be incorporated into, e.g., the exemplary turbofan engine  10  described above with reference to  FIG. 1 . However, in other exemplary embodiments, the turbine section  100  may be integrated into any other suitable machine utilizing a turbine. 
     Accordingly, it will be appreciated that the gas turbine engine within which the turbine section  100  is included generally defines a radial direction R, an axial direction A, a circumferential direction C extending about the axial direction A (see  FIG. 3 ), and a longitudinal centerline  102 . Further, the turbine section  100  includes a turbine  104 . For example, in certain embodiments, the turbine  104  may be a low pressure turbine (such as the exemplary low pressure turbine  30  of  FIG. 1 ), or alternatively may be any other turbine (such as, a high pressure turbine, an intermediate turbine, a dual use turbine functioning as part of a high pressure turbine and/or a low pressure turbine, etc.). 
     Moreover, for the exemplary embodiment depicted, the turbine  104  includes a first plurality of rotor blades, or rather a first plurality of turbine rotor blades  106 , and a second plurality of rotor blades, or rather a second plurality of turbine rotor blades  108 . As will be discussed in greater detail below, the first plurality of turbine rotor blades  106  and second plurality of turbine rotor blades  108  are alternatingly spaced along the axial direction A. 
     Referring first to the first plurality of turbine rotor blades  106 , each of the first plurality of turbine rotor blades  106  extends generally along the radial direction R between a radially inner end  110  and a radially outer end  112 . Additionally, the first plurality of turbine rotor blades  106  includes a first turbine rotor blade  106 A, a second turbine rotor blade  106 B, and a third turbine rotor blade  106 C, each spaced apart from one another generally along the axial direction A. At least two of the first plurality of turbine rotor blades  106  are spaced from one another along the axial direction A and coupled to one another at the respective radially inner ends  110 . More specifically, for the embodiment depicted, each of the first turbine rotor blade  106 A, the second turbine rotor blade  106 B, and the third turbine rotor blade  106 C are coupled to one another through their respective radially inner ends  110 . More specifically, still, each of the first turbine rotor blade  106 A, the second turbine rotor blade  106 B, and the third turbine rotor blade  106 C of the first plurality of turbine rotor blades  106  are coupled at their respective radially inner ends  110  through an inner drum  114 . 
     Further, the second plurality of turbine rotor blades  108 , each also extend generally along the radial direction R between a radially inner end  118  and a radially outer end  120 . Additionally, for the embodiment depicted, the second plurality of turbine rotor blades  108  includes a first turbine rotor blade  108 A, a second turbine rotor blade  108 B, and a third turbine rotor blade  108 C, each spaced apart from another generally along the axial direction A. For the embodiment depicted, at least two of the second plurality of turbine rotor blades  108  are spaced from one another along the axial direction A and coupled to one another at the respective radially outer ends  120 . More specifically, for the embodiment depicted, each of the first turbine rotor blade  108 A, the second turbine rotor blade  108 B, and the third turbine rotor blade  108 C of the second plurality of turbine rotor blades  108  are mechanically coupled to one another through their respective radially outer ends  120 . More specifically, still, each of the first turbine rotor blade  108 A, the second turbine rotor blade  108 B, and the third turbine rotor blade  108 C of the second plurality of turbine rotor blades  108  are coupled at their respective radially outer ends  120  through an outer drum  116 . 
     It should be appreciated, however, that in other exemplary embodiments, the first plurality of turbine rotor blades  106  and/or the second plurality of turbine rotor blades  108  may be coupled together in any other suitable manner, and that as used herein, “coupled at the radially inner ends” and “coupled at the radially outer ends” refers generally to any direct or indirect coupling means or mechanism to connect the respective components. For example, in certain exemplary embodiments, the second plurality of turbine rotor blades  108  may include multiple stages of rotors (not shown) spaced along the axial direction A, with the first turbine rotor blade  108 A, the second turbine rotor blade  108 B, and the third turbine rotor blade  108 C coupled to the respective stages of rotors at the respectively radially inner ends  118  through, e.g. dovetail base portions. The respective stages of rotors may, in turn, be coupled together to therefore “couple the second plurality of turbine rotor blades  108  at their respective radially inner ends  118 .” 
     Referring still to the embodiment depicted in  FIG. 2 , as stated, the first plurality of turbine rotor blades  106  and the second plurality of turbine rotor blades  108  are alternatingly spaced along the axial direction A. As used herein, the term “alternatingly spaced along the axial direction A” refers to the second plurality of turbine rotor blades  108  including at least one turbine rotor blade positioned along the axial direction A between two axially spaced turbine rotor blades of the first plurality of turbine rotor blades  106 . For example, for the embodiment depicted, alternatingly spaced along the axial direction A refers to the second plurality of turbine rotor blades  108  including at least one turbine rotor blade positioned between the first and second turbine rotor blades  106 A,  106 B of the first plurality of turbine rotor blades  106  along the axial direction A, or between the second and third turbine rotor blades  106 B,  106 C of the first plurality of turbine rotor blades  106  along the axial direction A. More specifically, for the embodiment depicted, the first turbine rotor blade  106 A of the first plurality of turbine rotor blades  106  is positioned forward of the first turbine rotor blade  108 A of the second plurality of turbine rotor blades  108 ; the second turbine rotor blade  106 B of the first plurality of turbine rotor blades  106  is positioned between the first and second turbine rotor blades  108 A,  108 B of the second plurality of turbine rotor blades  108 ; and the third turbine rotor blade  106 C of the first plurality of turbine rotor blades  106  is positioned between the second and third turbine rotor blades  108 B,  108 C of the second plurality of turbine rotor blades  108 . 
     Notably, however, in other exemplary embodiments, the first plurality of turbine rotor blades  106  may have any other suitable configuration and/or the second plurality of turbine rotor blades  108  may have any other suitable configuration. For example, it will be appreciated that for the embodiments described herein, the first turbine rotor blade  106 A, second turbine rotor blade  106 B, and third turbine rotor blade  106 C of the first plurality of turbine rotor blades  106  generally represent a first stage of turbine rotor blades, a second stage of turbine rotor blades, and a third stage of turbine rotor blades, respectively. It will similarly be appreciated that the first turbine rotor blade  108 A, second turbine rotor blade  108 B, and third turbine rotor blade  108 C of the second plurality of turbine rotor blades  108  each also generally represent a first stage of turbine rotor blades, a second stage of turbine rotor blades, and a third stage of turbine rotor blades, respectively. In other exemplary embodiments, the first plurality of turbine rotor blades  106  and/or the second plurality of turbine rotor blades  108  may include any other suitable number of stages of turbine rotor blades, such as two stages, four stages, etc., and further that in certain exemplary embodiments, the turbine  104  may additionally include one or more stages of stator vanes. 
     Moreover, for the embodiment depicted, the gas turbine engine further includes a gearbox  122  and a spool  124 , with the first plurality of turbine rotor blades  106  and the second plurality of turbine rotor blades  108  rotatable with one another through the gearbox  122  and drivingly coupled to the spool  124 . In at least certain exemplary embodiments, the spool  124  may be configured as, e.g., the exemplary low pressure spool  36  described above with reference to  FIG. 1 . It should be appreciated, however, that in other exemplary embodiments, the spool  124  may be any other spool (e.g., a high pressure spool, an intermediate spool, etc.). Additionally, the exemplary turbine section  100  further includes stationary frame member, and more specifically a turbine center frame  126  and a turbine rear frame  128 . 
     Referring particularly to the gearbox  122 , it will be appreciated that the gearbox  122  generally includes a casing  130 . The casing  130  is depicted in phantom in  FIG. 2  for clarity. The exemplary gearbox  122  depicted further includes a first gear  132  coupled to the first plurality of turbine rotor blades  106 , a second gear  134  coupled to the second plurality of turbine rotor blades  108 , and an intermediate gear  136  positioned between the first gear  132  and second gear  134  and coupled to a turbine frame, or rather to the turbine center frame  126 . The intermediate gear  136  is rotatably coupled to the turbine center frame  126 . In such a manner, it will be appreciated that for the embodiment depicted, the intermediate gear  136  remains stationary in the circumferential direction C, such that the first plurality of turbine rotor blades  106  are configured to rotate in an opposite circumferential direction than the second plurality of turbine rotor blades  108 . 
     More specifically, referring briefly to  FIG. 3 , an orientation of the first plurality of turbine rotor blades  106  and the second plurality of turbine rotor blades  108  is generally provided. As shown, the embodiment of  FIG. 3  depicts a first stage of turbine rotor blades  106 A of the first plurality of turbine rotor blades  106  and a first stage of turbine rotor blades  108 A of the second plurality of turbine rotor blades  108 . In the embodiment shown, the first plurality of turbine rotor blades  106  are configured to rotate in a first circumferential direction C 1 , while the second plurality of turbine rotor blades  108  are configured to rotate in a second circumferential direction C 2 . It should be understood that the first circumferential direction C 1  and the second circumferential direction C 2  as used and described herein are intended to denote directions relative to one another. Therefore, the first circumferential direction C 1  may refer to a clockwise rotation (viewed from downstream looking upstream) and the second circumferential direction C 2  may refer to a counter-clockwise rotation (viewed from downstream looking upstream). Alternatively, the first circumferential direction C 1  may refer to a counter-clockwise rotation (viewed from downstream looking upstream) and the second circumferential direction C 2  may refer to a clockwise rotation (viewed from downstream looking upstream). 
     Referring still to  FIG. 3 , it will further be appreciated that for the embodiment depicted, each turbine rotor blade  106 A of the first plurality of turbine rotor blades  106  includes an airfoil  138 , and similarly, each turbine rotor blade  108 A of the second plurality of turbine rotor blades  108  includes an airfoil  140 . The airfoils  138  each define an exit angle  142 , and similarly the airfoils  140  each define an exit angle  144 . Further, the airfoils  138 ,  140  may each further include a suction side  192  and a pressure side  194 . The exit angles  142 ,  144  of the airfoils  138 ,  140 , respectively, as well as the pressure and suction sides (not labeled) of such airfoils  138 ,  140 , respectively, may cause the first plurality of turbine rotor blades  106  and second plurality of turbine rotor blades  108  to rotate in the first and second circumferential directions C 1 , C 2 , respectively. It will be appreciated, however, that in other embodiments, the airfoils  138 ,  140  may have any other suitable configuration. 
     Referring back to  FIG. 2 , and as previously noted, the exemplary gearbox  122  depicted generally includes the first gear  132  coupled to the first plurality of turbine rotor blades  106 , the second gear  134  coupled to the second plurality of turbine rotor blades  108 , and the intermediate gear  136  positioned between the first gear  132  and second gear  134  and coupled to a turbine frame, or rather to the turbine center frame  126 . More specifically, the first plurality of turbine rotor blades  106  is coupled to the first gear  132  of the gearbox  122  through a first support member  146 , the second plurality of turbine rotor blades  108  is coupled to the second gear  134  of the gearbox  122  through a second support member  148 , and the intermediate gear  136  of the gearbox  122  is coupled to the turbine center frame through a frame member  150 . Further, it will be appreciated that for the embodiment depicted, the second support member  148 , the second gear  134  of the gearbox  122 , or both are connected to the spool  124 . 
     Further, the intermediate gear  136  defines an axis of rotation  152  and is rotatably coupled to the frame member  150  such that it may rotate about the axis of rotation  152 . For the embodiment depicted, the axis of rotation  152  defines an angle with the radial direction R less than about seventy-five (75) degrees. More specifically, for the embodiment shown, the angle the axis of rotation  152  defines with the radial direction R is about zero (0) degrees (i.e., less than or equal to ten (10) degrees; notably, the angle is not depicted or labeled in  FIG. 2  because the angle is zero degrees; cf.  FIG. 8 ). In such a manner, it will be appreciated that the intermediate gear  136  defines a forward end  154  and an aft end  156  generally along the axial direction A of the gas turbine engine. The first gear  132  of the gearbox  122  meshes with the intermediate gear  136  at the forward end  154  and at the second gear  134  meshes with the intermediate gear  136  at the aft end  156 . 
     Further, still, as is depicted in phantom, the first gear  132  and intermediate gear  136  together define a first intersection line  158  where the first gear  132  meshes with the intermediate gear  136 . Similarly, the second gear  134  and the intermediate gear  136  together define a second intersection line  160  where the second gear  134  meshes with the intermediate gear  136 . For the embodiment shown, the first intersection line  158  defines a first intersection angle  162  less than about seventy-five (75) degrees with the radial direction R (notably, the first intersection angle  162  is not depicted in  FIG. 2  since the angle is equal to zero; cf.  FIG. 5 ), and more specifically the first intersection angle  162  defined by the first intersection line  158  with the radial direction R is about zero (0) degrees. Similarly, the second intersection line  160  defines a second intersection angle  164  less than about seventy-five (75) degrees with the radial direction R (again, the second intersection angle  164  is not depicted in  FIG. 2  since the angle is equal to zero; cf.  FIG. 5 ), and more specifically, the second intersection angle  164  defined by the second intersection line  160  with the radial direction R is also about zero (0) degrees. As such, it will be appreciated that for the embodiment of  FIG. 2 , the first intersection angle  162 , the second intersection angle  164 , and the angle the axis of rotation  152  defines with the radial direction R are each equal to about zero (0) degrees. 
     Referring now also to  FIG. 4 , a perspective, cross-sectional view of the exemplary gearbox  122  of  FIG. 2  is depicted. As is shown, the intermediate gear  136  of the gearbox  122  is a first intermediate gear  136 A of a plurality of intermediate gears  136 . The plurality of intermediate gears  136  are spaced along the circumferential direction C of the gas turbine engine within the gearbox  122 . Each of the plurality of intermediate gears  136  may be configured in a similar manner to the intermediate gear  136  described above with reference to  FIG. 2 . For example, each of the plurality of intermediate gears  136  depicted is positioned between the first gear  132  and the second gear  134 , and is rotatably coupled to the frame member  150 , defining an axis of rotation  152  about which it is configured to rotate. For the embodiment shown, the axis of rotation  152  of each of the plurality of intermediate gears  136  defines an angle with the radial direction R less than about seventy-five (75) degrees, and more specifically, equal to about zero (0) degrees. (Notably, the gears  132 ,  134 ,  136  are depicted without gear teeth for clarity. It will be appreciated that any suitable gear teeth configuration may be utilized.) 
     For the embodiment depicted, the gearbox  122  includes between two and eight intermediate gears  136 , and more specifically, gearbox  122  includes six intermediate gears  136  spaced along the circumferential direction C. However, in other embodiments, the gearbox  122  may include any other suitable number of intermediate gears  136 . Moreover, it will be appreciated that for the embodiment depicted, the second support member  148  extends continuously from the second plurality of turbine rotor blades  108 , to the second gear  134  of the gearbox  122 , and to the spool  124 . It will be appreciated, however, that in other embodiments, the second support member  148 , the spool  124 , or both, may include one or more joints or breaks to facilitate installation of the gearbox  122 . By way of example, in certain exemplary embodiments, the second support member  148  may include a joint proximate the second gear  134 , proximate the spool  124 , or both. 
     Inclusion of a gearbox  122  configured in such a manner may allow for the gearbox  122  to occupy a smaller radial space inward of the turbine  104 , as compared to, e.g., a traditional planetary gearbox. Such may therefore allow for a turbine having a smaller diameter, saving weight and cost and increasing an overall efficiency. 
     It will be appreciated, however, that the exemplary turbine  104  and gearbox  122  depicted in  FIGS. 2 and 4  are provided by way of example only. For example, in other embodiments, although not depicted, the turbine  104 , the gearbox  122 , or both may include one or more bearing assemblies for facilitating rotation of the various components therein. 
     Moreover, it will be appreciated that although the exemplary intermediate gear(s)  136  of the gearbox  122  depicted in  FIGS. 2 and 4 , and discussed above, are depicted as a single gear rotatable about its axis of rotation  152 , in other embodiments, the intermediate gear  136 , as well as the first gear  132  and second gear  134 , may have any other suitable configuration. 
     For example, referring now briefly to  FIG. 5 , providing a schematic view of a turbine  104  and gearbox  122  in accordance with another exemplary embodiment of the present disclosure, the intermediate gear  136  is configured as a bevel gear. More specifically, the intermediate gear  136  defines an axis of rotation  152 , a first intersection line  158  where the intermediate gear  136  meshes with first gear  132 , and a second intersection line  160  where the intermediate gear  136  meshes with the second gear  134 . The first intersection line  158  defines a first intersection angle  162  with the axis of rotation  152  (and the radial direction R for the embodiment depicted, as the axis of rotation  152  is parallel to the radial direction R) greater than zero (0) degrees and less than 90 degrees, such as greater than fifteen (15) degrees and less than seventy-five (75) degrees, such as greater than thirty (30) degrees and less than sixty (60) degrees, such as about forty-five (45) degrees. Similarly, the second intersection line  160  defines a second intersection angle  164  with the axis of rotation  152  (and the radial direction R for the embodiment depicted, as the axis of rotation  152  is parallel to the radial direction R) substantially equal to the angle  162  defined between the first intersection line  158  and the axis of rotation  152  (e.g., greater than zero (0) degrees and less than 90 degrees, such as greater than fifteen (15) degrees and less than seventy-five (75) degrees, such as greater than thirty (30) degrees and less than sixty (60) degrees, such as about forty-five (45) degrees). Such a configuration may assist with maintaining the first plurality of turbine rotor blades  106  and the second plurality of turbine rotor blades  108  constrained along the radial direction R during operation of the gas turbine engine. 
     Further, in still other exemplary embodiments, the intermediate gear  136  may have any other suitable configuration. For example, referring now to  FIGS. 6 and 7 , schematic views of turbines  104  and gearboxes  122  in accordance with other exemplary embodiments of the present disclosure are provided. Referring particularly to  FIG. 6 , for the embodiment depicted, the intermediate gear  136  is configured as a compound gear, the compound gear including a radially outer gear  166  and a radially inner gear  168  rotatable with one another, and more specifically, fixed to one another. For the embodiment shown, the radially outer gear  166  meshes with the first gear  132  and the radially inner gear  168  meshes with the second gear  134 . The radially outer gear  166  defines a first diameter  170  and the radially inner gear  168  defines a second diameter  172 . Varying the first diameter  170  and the second diameter  172  may allow for varying a gear ratio between the first plurality of turbine rotor blades  106  and the second plurality of turbine rotor blades  108 . For the embodiment shown, the first diameter  170  is greater than the second diameter  172 , such that the first plurality of turbine rotor blades  106  may rotate more quickly than the second plurality of turbine rotor blades  108 . Notably, however, in other embodiments, such as the exemplary embodiment of  FIG. 7 , the second diameter  172  may be greater than the first diameter  170 , such that the second plurality of turbine rotor blades  108  may rotate more quickly than the first plurality of turbine rotor blades  106 . 
     Further, still, in other embodiments, the gearbox  122  may have still other configurations. For example, referring now to  FIG. 8 , providing an schematic view of a turbine  104  and gearbox  122  in accordance with yet another exemplary embodiment of the present disclosure, it will be appreciated that the intermediate gear  136  is tilted, such that an angle  174  the axis of rotation  152  defines with the radial direction R is not equal to zero (0). For example, for the embodiment shown, the angle  174  of the axis of rotation  152  defines with the radial direction R may be less than about seventy-five (75) degrees, such as less than about thirty (30) degrees, such as greater than about ten (10) degrees. Notably, although the intermediate gear  136  is depicted as being tilted forward in the embodiment of  FIG. 8 , in other exemplary embodiments the intermediate gear  136  may alternatively be tilted aft. 
     Moreover, as with the embodiment of, e.g.,  FIG. 5 , described above, the intermediate gear  136  defines a first intersection line  158  where the intermediate gear  136  meshes with first gear  132 , and a second intersection line  160  meshes with the second gear  134 . The first intersection line  158  defines an angle  174  with the axis of rotation  152  greater than zero (0) degrees and less than about ninety (90) degrees, such as greater than about fifteen (15) degrees and less than about seventy-five (75) degrees, and similarly the second intersection line  160  defines an angle  174  with the axis of rotation  152  substantially equal to the angle  174  defined between the first intersection line  158  and the axis of rotation  152 . However, since the axis of rotation  152  is not parallel to the radial direction R, a first intersection angle  162  defined between the first intersection line  158  and the radial direction R is not equal to a second intersection angle  164  defined between the second intersection line  160  and the radial direction R. 
     It will be appreciated that by tilting the intermediate gear  136 , the gearbox  122  may effectively change a gear ratio between the first gear  132  and second gear  134 , and more specifically, between the first plurality of turbine rotor blades  106  and the second plurality. For example, for the embodiment shown, the second plurality of turbine rotor blades  108  may be configured to rotate more quickly than the first plurality of turbine rotor blades  106 . However, in other embodiments, the configuration may be switched, such that the first plurality of turbine rotor blades  106  is configured to rotate more quickly the second plurality of turbine rotor blades  108 . 
     It will further be appreciated that in other exemplary embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, although the disclosure herein refers to a gearbox within a turbine, in other exemplary embodiments, the gearbox may be positioned within a compressor. With such an exemplary embodiment, the first plurality of rotor blades may instead be a first plurality of compressor rotor blades, and the second plurality of rotor blades may instead be a second plurality of compressor rotor blades. Other configurations are contemplated as well. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.