Abstract:
A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section having a plurality of fan blades. The plurality of fan blades has a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath. A core engine includes a first turbine configured to drive a first compressor, and a fan drive turbine configured to drive the fan section. A method of designing a gas turbine engine is also disclosed.

Description:
CROSS REFERENCE TO RELATED APPLICATION 
       [0001]    This application is a continuation of U.S. patent application Ser. No. 14/496,574, filed Sep. 25, 2014, which is a divisional application of Ser. No. 14/087,471, filed Nov. 22, 2013, and entitled “Geared Turbofan Engine Gearbox Arrangement,” the disclosure of which is incorporated by reference herein in its entirety as if set forth at length. 
     
    
     BACKGROUND 
       [0002]    The disclosure relates to turbofan engines. 
         [0003]    Gas turbine engines and similar structures feature a number of subassemblies mounted for rotation relative to a fixed case structure. Such engines typically have a number of main bearings reacting radial and/or thrust loads. Examples of such bearings are rolling element bearings such as ball bearings and roller bearings. Typically such bearings all react radial loads. Some such bearings also react axial (thrust) loads (either unidirectionally or bidirectionally). Ball bearings typically react thrust loads bidirectionally. However, if the inner race is configured to engage just one longitudinal side of the balls while the outer race engages the other longitudinal side, the ball bearing will react thrust unidirectionally. 
         [0004]    Tapered roller bearings typically react thrust unidirectionally. Two oppositely-directed tapered roller bearings may be paired or “duplexed” to react thrust bidirectionally. An example is found in the fan shaft bearings of U.S. Patent Application Publication 2011/0123326A1, which is incorporated herein by reference in its entirety and which is entitled “Bellows Preload and Centering Spring for a Fan Drive Gear System”. 
         [0005]    U.S. Patent Application Publication 2013/0192198, which is incorporated herein by reference in its entirety and which entitled “Compressor Flowpath”, discloses a flowpath through a compressor having a low slope angle. 
         [0006]    For controlling aspects of the flowpath passing through the fan duct, some turbofan engines include controllable features such as variable fan blade pitch and variable area fan exhaust nozzles. U.S. Pat. No. 5,431,539, which is incorporated herein by reference in its entirety and which is entitled “Propeller Pitch Change Mechanism”, and U.S. Pat. No. 5,778,659, which is incorporated herein by reference in its entirety and which is entitled “Variable Area Fan Exhaust Nozzle Having Mechanically Separate Sleeve and Thrust Reverser Actuation System”, disclose respective such systems. 
         [0007]    Unless explicitly or implicitly indicated otherwise, the term “bearing” designates an entire bearing system (e.g., inner race, outer race and a circumferential array of rolling elements) rather than the individual rolling elements. The term “main bearing” designates a bearing used in a gas turbine engine to support the primary rotating structures within the engine that produce thrust. This is distinguished, for example, from an accessory bearing (which is a bearing that supports rotating structures that do not produce thrust such as the fuel pump or oil pump bearings in an accessory gearbox). 
         [0008]    There are several different factors influencing flowpath geometry at certain locations in the engine. Weight, material strength and aerodynamics influence desirable core flowpath radius at different locations within the compressor and turbine sections. As noted above, U.S. Patent Application Publication 2013/0192198 discloses certain advantageous aspects of flowpath geometry within a compressor. This, however, may be competing with considerations regarding the core flowpath elsewhere in the engine. For example, the presence of an actuation mechanism or variable pitch fan blades may mandate a relatively large hub diameter. Similarly, the presence of a drive gear system axially between the compressor and the fan may also cause relatively high core flowpath diameters. Normally, it may be desirable to minimize radial turning of the core flow between such high radius sections and a lower diameter compressor section downstream thereof. Of particular importance to flowpath geometry and overall engine efficiency, however, are the bearing arrangements used to support the various rotating structures; improvements in this area are, therefore, always of interest to the turbofan engine designer. 
       SUMMARY 
       [0009]    A gas turbine engine according to an example of the present disclosure includes a fan section having a plurality of fan blades. The plurality of fan blades has a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath, and a ratio of Rh to Rt being less than 0.4. A core engine includes a first turbine configured to drive a first compressor, and a fan drive turbine configured to drive the fan section. The fan drive turbine includes  5  stages, and the fan drive turbine includes fewer stages than the first compressor. 
         [0010]    A further embodiment of any of the foregoing embodiments includes a second compressor upstream of the first compressor. The second compressor includes a greater number of stages than the first turbine. 
         [0011]    In a further embodiment of any of the foregoing embodiments, the core engine defines a core flowpath and a longitudinal axis. A second inboard boundary of the core flowpath has a radius R 1  defined at a first stage of a second compressor and a radius R 2  defined at a splitter rim. Each radius R 1 , R 2  is defined relative to the longitudinal axis. The splitter rim is configured to guide flow into the core flowpath. A ratio of R 1  to R 2  is greater than 0.5. 
         [0012]    In a further embodiment of any of the foregoing embodiments, the ratio of R 1  to R 2  is between 0.55 and 1.0. 
         [0013]    In a further embodiment of any of the foregoing embodiments, the ratio of Rh to Rt is less than about 0.30. 
         [0014]    In a further embodiment of any of the foregoing embodiments, the first turbine includes two stages. 
         [0015]    A further embodiment of any of the foregoing embodiments includes a geared architecture configured to rotate the fan section at a different speed than the fan drive turbine. 
         [0016]    A gas turbine engine according to the example of the present disclosure includes a fan section having a plurality of fan blades configured to deliver airflow to a bypass duct, the plurality of fan blades having a peak tip radius Rt and an inboard leading edge radius Rh defined at a first inboard boundary of a first flowpath, and a ratio of Rh to Rt being less than 0.4. A core engine defines a core flowpath and a longitudinal axis. The core engine includes a first turbine configured to drive a first compressor, a second compressor, and a fan drive turbine configured to drive the fan section and arranged aft of the first turbine. A second inboard boundary of the core flowpath has a radius R 1  defined at a first stage of the second compressor and a radius R 2  defined at a splitter rim. Each radius R 1 , R 2  is defined relative to the longitudinal axis. The splitter rim is configured to guide flow into the core flowpath. A ratio of R 1  to R 2  is greater than 0.5. 
         [0017]    In a further embodiment of any of the foregoing embodiments, the ratio of R 1  to R 2  is between 0.55 and 1.0. 
         [0018]    In a further embodiment of any of the foregoing embodiments, the ratio of Rh to Rt is less than about 0.30. 
         [0019]    In a further embodiment of any of the foregoing embodiments, six or more turbine stages are arranged downstream of the first turbine. 
         [0020]    In a further embodiment of any of the foregoing embodiments, the first turbine includes two stages. 
         [0021]    A further embodiment of any of the foregoing embodiments includes a geared architecture configured to rotate the fan section at a different speed than the fan drive turbine. 
         [0022]    A method of designing a gas turbine engine according to an example of the present disclosure includes providing a fan section having a plurality of fan blades. The fan blades have a peak tip radius Rt and an inboard leading edge radius Rh defined at a first inboard boundary of a first flowpath. A ratio of Rh to Rt is less than 0.4. The method includes providing a core engine including a first turbine configured to drive a first compressor, a second compressor, and a fan drive turbine configured to drive the fan section. The fan drive turbine includes 5 stages. The first compressor includes a greater number of stages than the fan drive turbine, and the second compressor includes a greater number of stages than the first turbine. 
         [0023]    In a further embodiment of any of the foregoing embodiments, the ratio of Rh to Rt is less than about 0.30. 
         [0024]    In a further embodiment of any of the foregoing embodiments, the core engine defines a core flowpath and a longitudinal axis. A second inboard boundary of the core flowpath has a radius R 1  defined at a first stage of the second compressor and a radius R 2  defined at a splitter rim. Each radius R 1 , R 2  is defined relative to the longitudinal axis. The splitter rim is configured to guide flow into the core flowpath. A ratio of R 1  to R 2  is greater than 0.5. 
         [0025]    In a further embodiment of any of the foregoing embodiments, the ratio of Rh to Rt is less than about 0.30. 
         [0026]    In a further embodiment of any of the foregoing embodiments, the ratio of R 1  to R 2  is between 0.55 and 1.0. 
         [0027]    A further embodiment of any of the foregoing embodiments includes providing a geared architecture configured to rotate the fan section at a different speed than the fan drive turbine. 
         [0028]    In a further embodiment of any of the foregoing embodiments, the first turbine includes two stages. 
         [0029]    One aspect of the disclosure involves a three-spool turbofan engine comprising a fan having a plurality of blades. A transmission is configured to drive the fan. The fan blades have a peak tip radius RT. The fan blades have an inboard leading edge radius RH at an inboard boundary of the flowpath. A ratio of RH to RT is less than about 0.40. 
         [0030]    A further embodiment may additionally and/or alternatively include the engine being a three spool turbofan engine comprising a first spool comprising a first pressure turbine and a first shaft coupling the first pressure turbine to the transmission. A second spool comprises a second pressure turbine, a first compressor, and a second spool shaft coupling the second pressure turbine to the second spool compressor. A core spool comprises a third pressure turbine, a second compressor, and a core shaft coupling the third pressure turbine to the second compressor. A combustor is between the second compressor and the third pressure turbine. 
         [0031]    A further embodiment may additionally and/or alternatively include the first compressor having a rear hub engaging a bearing, said bearing engaging the first shaft, and another bearing engaging the first compressor and the first shaft. 
         [0032]    A further embodiment may additionally and/or alternatively include a ring gear of the transmission being mounted to rotate with the fan as a unit. 
         [0033]    A further embodiment may additionally and/or alternatively include each fan blade having a leading edge and a trailing edge. A splitter is positioned along a flowpath through the engine and having a leading rim separating a core branch of the flowpath from a bypass branch of the flowpath. An inboard boundary of the core flowpath has a radius RII at an axial position of the splitter rim and a radius RI at a leading stage of blades of the first compressor. A ratio of an axial length L 10  between the splitter rim and the leading stage of blades of the first compressor at the inboard boundary of the core flowpath to the radius RII is less than 1.2. 
         [0034]    A further embodiment may additionally and/or alternatively include each fan blade having a leading edge and a trailing edge. A splitter is positioned along a flowpath through the engine and having a leading rim separating a core branch of the flowpath from a bypass branch of the flowpath. An inboard boundary of the core flowpath has a radius RII at an axial position of the splitter rim and a radius RI at a leading stage of blades of the first compressor. A ratio of the radius RI to the radius RII is greater than 0.50. The ratio of the radius RI to the radius RII may be 0.55-1.0. 
         [0035]    A further embodiment may additionally and/or alternatively include the fan blades being non-variable. 
         [0036]    A further embodiment may additionally and/or alternatively include a variable fan nozzle. 
         [0037]    A further embodiment may additionally and/or alternatively include the engine having a plurality of main bearings. A first of said main bearings engages a static support and a forward hub of the second spool. A second of said main bearings engages the first shaft and the forward hub of the second spool. 
         [0038]    A further embodiment may additionally and/or alternatively include the first bearing and the second bearing be behind the transmission. 
         [0039]    A further embodiment may additionally and/or alternatively include a length between said first of said main bearings and a center of gravity of said rotor of the first compressor being less than half of a disk to disk overall length of the first compressor. 
         [0040]    A further embodiment may additionally and/or alternatively include a length between said first of said main bearings and a center of gravity of a rotor of the second spool compressor being less than a radius RI of the inboard boundary of the core flowpath at a leading stage of blades of the first compressor. 
         [0041]    A further embodiment may additionally and/or alternatively include the forward hub extending forward from a disk of the first compressor. 
         [0042]    A further embodiment may additionally and/or alternatively include the forward hub extending forward from a bore of the disk of the first compressor. 
         [0043]    A further embodiment may additionally and/or alternatively include the first compressor having at least one disk forward of said disk. 
         [0044]    A further embodiment may additionally and/or alternatively include the static support passing through said at least one disk forward of said disk. 
         [0045]    A further embodiment may additionally and/or alternatively include the first compressor having at least two disks forward of said disk. 
         [0046]    A further embodiment may additionally and/or alternatively include said at least one disk being forward of a centerplane of the second bearing. 
         [0047]    A further embodiment may additionally and/or alternatively include the first bearing and the second bearing being non thrust roller bearings. 
         [0048]    A further embodiment may additionally and/or alternatively include rollers of the first bearing and the second bearing being at least partially longitudinally overlapping. 
         [0049]    A further embodiment may additionally and/or alternatively include a separation of a transverse centerplane of the first bearing and a transverse centerplane of the second bearing being less than a radius (RB) of the first bearing. 
         [0050]    A further embodiment may additionally and/or alternatively include a first seal sealing the first bearing and a second seal sealing the second bearing to isolate a transmission compartment ahead of the first bearing and the second bearing from a region behind the first bearing and the second bearing. 
         [0051]    A further embodiment may additionally and/or alternatively include the transmission comprising: a sun gear mounted to rotate with the first shaft; a ring gear mounted to rotate with the fan; a plurality of intermediate gears between the sun gear and the ring gear; and a carrier holding the intermediate gears. 
         [0052]    A further embodiment may additionally and/or alternatively include a third of said main bearings being a thrust bearing engaging the first spool shaft. 
         [0053]    A further embodiment may additionally and/or alternatively include a fourth of said main bearings being a non thrust roller bearings bearing engaging an aft end of the first spool shaft. 
         [0054]    A further embodiment may additionally and/or alternatively include the core shaft engaging at least two of said main bearings, and wherein at least one of said at least two of said main bearings is a thrust bearing. 
         [0055]    A further embodiment may additionally and/or alternatively include the first pressure turbine having three to five blade stages. 
         [0056]    A further embodiment may additionally and/or alternatively include the second spool shaft engaging at least two of said main bearings, at least one of which is a thrust bearing. 
         [0057]    A further embodiment may additionally and/or alternatively include an inter shaft bearing axially locating the first spool shaft. 
         [0058]    A further embodiment may additionally and/or alternatively include the first spool shaft engaging at least three of said main bearings. 
         [0059]    A further embodiment may additionally and/or alternatively include the fan being a single stage fan. 
         [0060]    A further embodiment may additionally and/or alternatively include the ratio of RH to RT being less than about 0.30. 
         [0061]    Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
         [0062]    These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0063]      FIG. 1  is a schematic longitudinal sectional view of a first turbofan engine embodiment. 
           [0064]      FIG. 1A  is an enlarged view of a forward portion of the engine of  FIG. 1 . 
           [0065]      FIG. 2  is a schematic longitudinal sectional view of a second turbofan engine embodiment. 
           [0066]      FIG. 2A  is an enlarged view of a forward portion of the engine of  FIG. 2 . 
           [0067]      FIG. 3  is a schematic longitudinal sectional view of a third turbofan engine embodiment. 
           [0068]      FIG. 4  is a schematic longitudinal sectional view of a fourth turbofan engine embodiment. 
       
    
    
       [0069]    Like reference numbers and designations in the various drawings indicate like elements. 
       DETAILED DESCRIPTION 
       [0070]      FIG. 1  shows a turbofan engine  20  having a central longitudinal axis or centerline  500 . The engine has a structural case including a core case  22 . The exemplary structural case further comprises a fan case  24  connected to the core case by a circumferential array of struts  26  and surrounding a fan  28 . The core case and the fan case may have respective outboard aerodynamic nacelles (shown schematically as  31  and  32 ). 
         [0071]    The exemplary forward rim of the fan case is proximate an engine inlet  30  receiving an inlet flow  502  when the engine is operating. The inlet flow passes downstream through the fan  28  and divides into a core flow  504  passing inboard along a core flowpath  506  (core branch of a combined flowpath) within the core case and a bypass flow  508  passing outboard along a bypass flowpath  510  (bypass branch of a combined flowpath) between the core case  22  and the fan case  24 . 
         [0072]    The bypass flowpath extends to an outlet  38 . The exemplary outlet  38  is defined by a variable nozzle assembly  35 . The exemplary variable nozzle assembly  35  includes a movable member  36  having a downstream/trailing end  37  for defining the outlet between the end  37  and the core nacelle  31 . The exemplary member  36  may articulate between at least two conditions or positions. The exemplary articulation involves an axial translation between a forward/retracted condition or position shown in solid line and a rearward/extended condition or position shown in broken line with the numeral  36 ′. The translation may be driven by an actuator (not shown) (e.g., a hydraulic actuator). 
         [0073]    The core flow  504  (or a majority portion thereof allowing for bleeds, etc.) passes sequentially through one or more compressor sections, a combustor, and one or more turbine sections before exiting a core outlet  34 . In the exemplary engine the fan is a single-stage fan having a single stage of fan blades  40 . Each of the compressor and turbine sections may include one or more blade stages mounted to rotate as a unit about the centerline  500 . The blade stages may be alternatingly interspersed with vane stages. Each compressor section is co-spooled with an associated turbine section. From upstream to downstream along the core flowpath, the exemplary engine has two compressor sections  42  and  44 , the combustor  45 , and three turbine sections  46 ,  48 , and  50 . The fan and compressor sections (and their stages) progressively compress inlet air which passes into the combustor for combustion with fuel to generate gas of increased pressure which passes downstream through the turbine sections where the gas pressure is progressively reduced as work is extracted. The turbine section  46  operates at a pressure that is higher than the intermediate turbine  48  and the low turbine  50  and is often referred to as a high (or third) pressure turbine (HPT) or a core turbine. The HPT blade stages are connected via a shaft  52  (“high shaft” or “core shaft”) to the blade stages of the compressor section  44  to drive that compressor section (often referred to as a high pressure compressor (HPC) or core compressor) to form a high spool or core spool. 
         [0074]    The turbine section  48  operates at a pressure range that is intermediate to the low and high pressure sections  50  and  46 . The turbine section  48  is thus often referred to as an intermediate (or second) pressure turbine (IPT). The IPT blade stages are connected via a shaft  54  (“intermediate shaft”) to the compressor section  42  to drive that compressor section (often referred to as an intermediate pressure compressor (IPC)) to form an intermediate spool. 
         [0075]    The turbine section  50  operates at a low pressure range relative to the high pressure turbine  46  and the intermediate pressure turbine  48  and is thus often referred to as a low (or first) pressure turbine (LPT) or as a fan drive turbine. The LPT blade stages are connected via a shaft  56  (“low shaft”) to a transmission  60  (e.g., an epicyclic transmission, more particularly a geared system known as a fan drive gear system (FDGS)) to indirectly drive the fan  28  with a speed reduction. 
         [0076]    An exemplary high pressure turbine  46  is a single or double stage turbine assembly (although three or more HPT stages are possible); an exemplary intermediate stage turbine  48  is a single or double stage turbine assembly (although three or more IPT stages are possible); an exemplary low pressure turbine  50  is a multi-stage turbine such as, for example, one or more stages, or more specifically three to five stages (although one or two stages is also possible). 
         [0077]    The exemplary transmission  60  ( FIG. 1A ) comprises a central externally-toothed sun gear  80 . The sun gear  80  is encircled by an internally-toothed ring gear  82 . A number of externally toothed star or planet gears  84  are positioned between and enmeshed with the sun gear  80  and ring gear  82 . The star or planet gears  84  can be referred to as intermediate gears. A cage or carrier assembly  86  carries the intermediate gears via associated bearings  88  for rotation about respective bearing axes. The exemplary bearings  88  may be rolling element bearings (e.g., ball or roller bearings) or may be journal bearings having external circumferential surface portions closely accommodated within internal bore surfaces of the associated intermediate gears  84 . Regardless of the type, the bearings may be metallic (such as aluminum titanium, other metal, or an alloy of more than one metal), ceramic, composite, or other material. 
         [0078]    The exemplary carrier assembly  86  comprises a front plate (e.g., annular) in front of the gears and a rear plate (e.g., annular) behind the gears. These plates may be mechanically connected by the bearings  88  and/or by linking portions between adjacent intermediate gears. 
         [0079]    In the exemplary embodiment, a forward end of the low shaft  56  is coupled to the sun gear  80 . The exemplary low shaft  56  has a generally rigid main portion  100  and a flexible forward portion  102 . A forward end of the portion  102  may have a splined outer diameter (OD) surface interfitting with a splined inner diameter (ID) surface of the sun gear  80  to transmit rotation. 
         [0080]    The exemplary carrier assembly  86  is substantially non rotatably mounted relative to the engine case  22 . In the exemplary embodiment, the carrier assembly  86  is coupled to the case  22  via a compliant flexure  110  that allows at least small temporary radial and axial excursions and rotational excursions transverse to the centerline  500 . The exemplary flexure  110  carries a circumferential array of fingers  111  engaging the carrier  86  (e.g., between adjacent gears  84 ). A peripheral portion of the flexure  110  is mounted to the case to resist rotation about the centerline  500 . Thus, flexing of the flexure accommodates the small excursions mentioned above while holding the carrier against rotation about the centerline. 
         [0081]    The exemplary ring  82  is coupled to the fan  28  to rotate with the fan  28  as a unit. In the exemplary embodiment a rear hub  122  of a main fan shaft  120  connects the fan  28  to the ring gear  82 . 
         [0082]    The speed reduction ratio is determined by the ratio of diameters of the ring gear  82  to the sun gear  80 . This ratio will substantially determine the maximum number of intermediate gears  84  in a given ring. The actual number of intermediate gears  84  will be determined by stability and stress/load sharing considerations. An exemplary reduction is between about 2:1 and about 13:1. Although only one intermediate gear  84  is necessary, in exemplary embodiments, the number of intermediate gears  84  may be between about three and about eleven. An exemplary gear layout with fixed carrier is found in U.S. Patent Application Publication 2012/0251306A1, which is incorporated by reference herein in its entirety and which is entitled “Fan Rotor Support. In addition, although the exemplary transmission  60  is described as being of a “star” type configuration, other types of configurations (such as “planetary” systems) are within the scope of this invention. 
         [0083]    Thus, the exemplary engine  20  has four main rotating components (units) rotating about the centerline  500 : the core spool (including the high pressure turbine  46 , the high shaft  52 , and the high pressure compressor  44 ); the intermediate spool (including the intermediate pressure turbine  48 , the intermediate shaft  54 , and the intermediate pressure compressor  42 ); the low spool (including the low pressure turbine  50 , low shaft  56 , and the sun gear  80 ); and the fan assembly (including the fan  28  itself, the fan shaft  120 , and the ring gear  82 ). Each of these four things needs to be supported against: radial movement; overturning rotations transverse to the centerline  500 ; and thrust loads (parallel to the centerline  500 ). Radial and overturning movements are prevented by providing at least two main bearings engaging each of the four units. 
         [0084]    Each unit would have to also engage at least one thrust bearing. The nature of thrust loads applied to each unit will differ. Accordingly, the properties of the required thrust bearings may differ. For example, the fan  28  primarily experiences forward thrust and, therefore, the thrust bearings engaging the fan  28  may be configured to address forward thrust but need not necessarily address rearward thrusts of similar magnitudes, durations, etc. 
         [0085]    The  FIG. 1  embodiment has two main bearings  148 ,  150  along the fan shaft forward of the transmission  60 . Inboard, the inner race of each bearing  148 ,  150  engages a forward portion of the shaft  120  aft of the fan  28 . Outboard, the outer race of each bearing  148 ,  150  engages static structure of the case. The exemplary static structure comprises a support  152  extending inward and forward from a forward frame  154 . These two bearings  148 ,  150  thus prevent radial excursions and overturning moments which the fan  28  may produce during flight. 
         [0086]    To resist thrust loads, one or both of the bearings  148 ,  150  may be thrust bearings. In an exemplary embodiment, both are thrust bearings (schematically shown as ball bearings). Both may be thrust bearings because there may typically be no differential thermal loading (and thus thermal expansion) of the support  152  relative to the shaft  120  between these bearings. Where the two coupled structures are subject to differences in thermal expansion, it may be desirable to have only one bearing be a thrust bearing. 
         [0087]    In one alternative example of a single thrust bearing and a single non thrust bearing, the bearing  150  would be a straight roller bearing with longitudinal roller axes configured to only handle radial loads. The other bearing (i.e., the bearing  148 ) would be a thrust bearing. Due to the significance of forward thrust loads on the fan  28 , the bearing  148  may be biased to resist forward loads. The exemplary bearing  148  may then be a bidirectional ball bearing or a bidirectional tapered roller bearing (e.g., wherein the rollers have a forward taper and forwardly converging roller axes to preferentially handle the forward thrust loads). A similar bidirectional tapered roller bearing is shown in U.S. Pat. No. 6,464,401 of Allard, which is incorporated herein by reference in its entirety and which is entitled “High Load Capacity Bi Directional Tapered Roller Bearing”. Ball bearings are typically bidirectional thrust bearings. However, a unidirectional ball bearing may be formed by having at least one of the races contacting only a single longitudinal side of the balls. 
         [0088]    An exemplary bearing arrangement for supporting the remaining three units is discussed below. Various aspects of each of these may be independently implemented or all may be implemented in a given engine. 
         [0089]    The exemplary low shaft  56  is principally radially supported by a forward bearing  162 , an intermediate bearing  170 , and an aft bearing  172 . The exemplary forward bearing  162  is indirectly radially grounded to the case  22 . An exemplary indirect grounding (discussed further below) is via the intermediate spool and bearing  160 . The exemplary bearing  160  ( FIG. 1A ) is directly radially grounded to the case (e.g., by a bearing support  164  extending inward from a frame  154  aft of the support  152 ).  FIG. 1  also shows an inlet guide vane array  155  immediately upstream of the struts of the frame  154  and an outlet guide vane array  157  immediately downstream of the frame  154  and upstream of the leading compressor stage. In exemplary implementations, the vanes of the array  157  may be variable vanes. The exemplary array  155  is immediately downstream of a splitter  159  dividing the core flowpath from the bypass flowpath. 
         [0090]    The exemplary bearing  170  intervenes directly between the low spool and intermediate spool at an intermediate location. In the exemplary embodiment, it is indirectly radially grounded by the bearing  220 . The bearing  220  is directly radially grounded by a support  240  extending radially inward from a structural vane array (frame)  242  between the compressor sections  42  and  44 . 
         [0091]    The exemplary aft bearing  172  is directly radially grounded to the case  22  via a support  180  extending inward from a frame  182  extending across the core flowpath  504 . The exemplary support  180  is aft of the LPT  50  with the frame  182  being a turbine exhaust frame. Alternative implementations may shift the support  180  forward of the LPT  50  to engage an inter turbine frame  183  between the turbine sections  48  and  50 . 
         [0092]    In the exemplary embodiment, the bearings  162  and  172  are non thrust roller bearings (e.g., straight roller bearings). The bearing  170  serves as inter-shaft thrust bearing (e.g., a bidirectional ball bearing) having an inner race engaging an intermediate portion of the low shaft  56  and an outer race engaging the intermediate shaft  54  to indirectly axially ground the low shaft  56  to the case  22  via the intermediate shaft  54 . 
         [0093]    By locating the bearing  170  relatively axially close to the bearing  220 , the bearing  170  may also provide an intermediate location of radial grounding in addition to the forward and aft radial groundings provided by the bearings  162  and  172 . Alternative implementations might eliminate or reduce the amount of this radial grounding. In the  FIG. 1  example, the bearings  160  and  162  are stacked so close as to be partially axially overlapping (i.e., axial overlap of their rollers) to provide a high degree of radial support. 
         [0094]    In contrast, there is a slight non overlap forward shift of the bearing  170  relative to the bearing  220 . In the exemplary engine, the outer race of the bearing  170  engages a forwardly projecting support extending forward from a rear hub  174  of the compressor section  42 . The exemplary rear hub  174  extends from a bore  175  of one of the disks of the compressor section  42 . Slight flexing of the hub  174  and the outer bearing support  173  protruding therefrom may provide a little more radial compliance than associated with the forward bearing  162 . 
         [0095]    The intermediate spool is supported by forward bearing  160 , intermediate bearing  220 , and an aft bearing  230 . In an exemplary embodiment, forward bearing  160  is a non thrust roller bearing providing radial retention only. The inner race of the bearing  160  (and outer race of the bearing  162 ) are mounted along respective outer and inner faces of a hub or support  236  extending forward from the bore  237  of one of the disks of the compressor section  42  (e.g., the first (upstream most) disk). The exemplary intermediate bearing  220  is a bidirectional thrust bearing (e.g., ball bearing) directly radially and axially supporting/grounding the intermediate spool via the support  240  extending to the inter compressor frame  242  between the compressor sections  42  and  44 . The bearing  230  indirectly radially supports/grounds the intermediate spool by engaging the intermediate spool and the low spool. In the exemplary embodiment, the inner race of the bearing  230  engages a portion of the intermediate shaft aft of the turbine section  48  and the outer race of the bearing  230  engages a support extending forward from a hub  248  of the LPT  50 . The exemplary hub  248  extends forward from the bore of a disk (e.g., the last or downstream most disk) of the LPT. 
         [0096]    The radial loads on the intermediate spool at the bearing  230  will primarily be transmitted to the low shaft  56  and through an aft portion of the low shaft  56  to the bearing  172  and grounded by the support  180  and frame  182 . Axial (thrust) loads will pass through the bearing  220 . 
         [0097]    Thus, thrust loads on the low spool are transmitted via the shaft  56  through the bearing  170 , through the intervening portion of the intermediate shaft/spool, to the bearing  220 , and grounded back through the support  240 . 
         [0098]    The core spool may be fully directly supported by two bearings  250  and  260  of which at least one would be a thrust bearing. In the exemplary embodiment, the bearing  250  is a forward bearing grounding a forward portion of the core shaft ahead of the compressor section  44  to the inter compressor frame  242  via a support  270 . The aft bearing  260  grounds a portion of the core shaft intermediate the compressor section  44  and turbine section  46  via a support  272  extending to a combustor frame  274  ahead of the turbine section  46 . In alternative embodiments, this aft bearing  260  may be shifted aft of the turbine section  46  via a support (not shown) to an inter turbine frame  278  between the sections  46  and  48 . In the exemplary implementation, the bearing  250  is a thrust bearing (e.g., a bidirectional ball bearing with its inner race engaging the core spool and its outer race engaging the support  270 ). The exemplary bearing  260  is a straight roller bearing with its inner race engaging the core shaft  52  and its outer race engaging the support  272 . The exemplary support  270  extends to a rear portion of the frame  240  aft of the support  242 . The exemplary inner race of the bearing  250  is mounted to a hub or support extending forward from a bore of a disk (e.g., the upstream most disk) of the compressor section  44 . 
         [0099]      FIG. 1  further shows the transmission  60  as having a centerplane  516  and the gears as having a gear width WG and the fan blade array as having a centerplane  518 . From fore to aft, the bearings have respective centerplanes  520 ,  522 ,  524 ,  526 ,  528 ,  530 ,  532 ,  534 ,  536 , and  538 . 
         [0100]    As discussed above, an exemplary embodiment places the centerplanes  524  and  526  of the bearings  160  and  162  relatively close to each other so as to best transmit radial loads from the low shaft  56  to the case. An exemplary separation between the planes  524  and  526  ( FIG. 1A ) in such embodiments is less than the characteristic radius of the bearing  160  (e.g., radius RB relative to the axis  500  of the intersections of the individual rolling element axes with the bearing centerplane). In contrast, the exemplary embodiment has a greater separation between the centerplanes  528  and  530  of the bearings  170  and  220 . This may provide a greater radial compliance at the associated intermediate location. 
         [0101]      FIG. 1A  further shows a transmission compartment  286  containing the transmission  60 . Aftward, the transmission compartment is largely bounded by the support  164  and bearings  160  and  162 . Seals may be provided to seal the transmission compartment  286  from a region  288  (e.g., a compressor compartment) aft thereof. The exemplary seals comprise an outer seal  290  sealing between the static structure and the intermediate spool and an inner seal  292  sealing between the intermediate spool and the low spool. Exemplary seal  290  is held by a carrier  291 . An exemplary carrier  291  is formed as an inward and aftward extension of the support  164  holding the seal  290  in sliding/sealing engagement with the low spool (e.g., with an inner race of the bearing  160 ). Similarly, a seal carrier  293  carries the exemplary seal  292 . In the exemplary embodiment, the seal carrier  293  is mounted to or formed as a portion of the low shaft main portion  100  holding the seal  292  in sealing and sliding engagement with the intermediate spool (e.g., with an outer race of the bearing  162 ). In alternative implementations, the carrier and seal elements of one or both of the sealing systems may be reversed (e.g., the seal carrier  293  could be formed as a portion of the hub  236  holding the seal  292  in sliding/sealing engagement with the low spool). 
         [0102]      FIG. 2  shows an alternate embodiment  320  which may be otherwise similar to the engine  20  but which has a forward shift of its compressor section  42 ′ relative to the compressor section  42  of  FIG. 1 . The exemplary forward shift may be achieved by having the hub or support structure  236  ( FIG. 2A ) that cooperates with the bearings  160  and  162  extend forward from the bore  237 ′ of an intermediate disk of the compressor section  42 ′ in distinction to the extension from the upstream most disk of the compressor section  42 . In the exemplary engine  320 , the hub  236  ( FIG. 2A ) extends from the third disk leaving two disks and their associated blade stages thereahead. The exemplary shift shifts at least one disk stage forward of the bearings  160  and/or  162 . In this example, the longitudinal position of the first disk (e.g., measured by the centerplane of its web and/or bore) is shifted ahead of the centerplanes of the bearings  160  and  162 . An exemplary shift places the first disk ahead of both bearings  160  and  162  and the second disk ahead of only the bearing  162 . However, other locations and combinations are possible. 
         [0103]    A further characterization of the longitudinal compactness involves the relationship between the first disk and the transmission.  FIG. 2A  shows a centerplane  560  of the first disk  340 . The centerplane  560  is behind the gear centerplane  516  by a length LD. Exemplary LD is 2.0 times the gear width WG or less, more particularly, 1.5 times WG or less. Alternatively characterized, exemplary LD is 60% or less of the core flowpath inboard radius RI at the disk centerplane  560 , more particularly, 50% or less or 35% or less of RI. 
         [0104]    Yet alternatively characterized relative to such a core flowpath inboard radius RG at the gear centerplane  516 , exemplary LD is 50% of RG or less, more particularly, 40% or less or 30% or less. 
         [0105]    To further facilitate longitudinal compactness, relative to the engine  20 , the engine  320  axially shrinks the frame  154 ′ relative to the frame  150 . In this example, the frame  154 ′ and its associated struts replace both the frame  154  and its associated struts and the inlet guide vane array  155  ( FIG. 1A ). The guide vane array  157  ( FIG. 1A ) downstream of the struts is effectively shifted forward to become  157 ′. Along with the foreshortening of the frame  154 ′, the outboard periphery and mounting location of the support  164  is shifted forward and outward to become  164 ′. Thus, the exemplary support  164 ′ is shallower than support  164  and partially overarches the span of the transmission gears. Because of this overarching, the fingered flexure  110  is shifted to be mounted to a mounting feature (e.g., flange)  110 ′ along the support  164 ′. 
         [0106]      FIG. 3  shows yet a further embodiment  420  reflecting the variation discussed above wherein the bearing  260  is shifted aft of the high pressure turbine section  46 . Other variations might add a second intermediate spool. Other variations include unducted fans. Other variations include multi stage fans.  FIG. 4  shows an axially compact engine lacking the forward shift of the  FIG. 3  embodiment but, instead, being otherwise similar to  FIG. 1  from the leading compressor disk forward. LPC stage count, however, is reduced relative to the  FIG. 1  embodiment so that the rear end of the LPC is shifted forward relative to the  FIG. 1  embodiment as is the rear end of the LPC of the  FIG. 3  embodiment. 
         [0107]      FIG. 4  further shows various radial measurements. A fan tip radius is shown as RT. In this exemplary embodiment, the tip maximum radius is at the blade leading edge.  FIG. 4  further shows a characteristic hub radius RH. Exemplary RH is defined as the flowpath inboard or inner diameter (ID) radius at the fan blade leading edge.  FIG. 4  further shows a core inlet inner radius RII which is measured as the inboard or ID radius of the core flowpath at the axial position of the forward rim of the splitter  159 . A characteristic compressor inlet radius may be measured as the aforementioned RI. Alternatively, this radius may be measured at the leading edge of the associated upstreammost blade stage. These will typically be very close to each other. 
         [0108]      FIG. 4  further shows an axial length L 10  between the locations at which RII and RI are measured.  FIGS. 2 and 4  also label a length LCG between a centerplane of the bearing  160  which may represent the closest main bearing behind the FDGS, the forwardmost bearing intervening directly (between the intermediate spool and the case or both) and the transverse plane of the center of gravity of the intermediate pressure compressor rotor (e.g., ignoring the intermediate pressure shaft aft of the bearing  170  and ignoring the intermediate pressure turbine rotor). A characteristic intermediate pressure compressor length LIC is shown as the center to center axial distance between the leading and trailing disks. Particularly, with the forward shifting of the IPC of the embodiments of  FIGS. 2 and 3  but also with the foreshortening of the embodiment of  FIG. 4 , LCG may be reduced relative to  FIG. 1 . The exemplary LCG may be reduced to less than RII and even to less than RI. LCG may also be reduced to less than one half of LIC (e.g., as shown in  FIGS. 2 and 3 ). The particular reconfiguration of  FIGS. 2 and 3  helps bring the center of gravity close to the plane of the bearing  160  to maintain stability. This stability reduces the radial loads that must be reacted by the bearing  220 . Thus, the bearing  170  may be more specifically configured for reacting thrust loads with less capacity to react radial loads and may be lightened. 
         [0109]    The use of “first”, “second”, and the like in the following claims is for differentiation within the claim only and does not necessarily indicate relative or absolute importance or temporal order. Similarly, the identification in a claim of one element as “first” (or the like) does not preclude such “first” element from identifying an element that is referred to as “second” (or the like) in another claim or in the description. 
         [0110]    One or more embodiments have been described. Nevertheless, it will be understood that various modifications may be made. For example, when applied to an existing basic configuration, details of such configuration or its associated environment may influence details of particular implementations. Accordingly, other embodiments are within the scope of the following claims.