Patent Publication Number: US-9850821-B2

Title: Gas turbine engine with fan-tied inducer section

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application is a continuation-in-part of U.S. patent application Ser. No. 13/406,819, filed Feb. 28, 2012. 
    
    
     BACKGROUND OF THE INVENTION 
     This disclosure relates to a gas turbine engine with a fan-tied inducer section. 
     A typical jet engine has multiple shafts or spools that transmit torque between turbine and compressor sections of the engine. In one example, a low speed spool generally includes a low shaft that interconnects a fan, a low pressure compressor, and a low pressure turbine. 
     The low pressure turbine drives the low shaft, which drives the low pressure compressor. A geared architecture connects the low shaft to the fan. Air exiting the fan at the root has relatively low energy, which generates a swirling effect that makes it difficult to efficiently feed air into the low pressure compressor. 
     SUMMARY OF THE INVENTION 
     In a featured embodiment, a gas turbine engine comprises a shaft defining an axis of rotation. A speed reduction device is driven by the shaft. A fan includes a fan rotor driven by the speed reduction device. A compressor section comprises a plurality of compressor stages driven by the shaft. At least one inducer stage is positioned aft of the fan and coupled for rotation with the fan rotor. The at least one inducer stage is positioned upstream of the compressor section. 
     In another embodiment according to the previous embodiment, the at least one inducer stage comprises one or more inducer blades fixed for rotation with the fan rotor and a core inlet stator fixed to a non-rotating engine structure. 
     In another embodiment according to any of the previous embodiments, the core inlet stator is positioned radially outward of the inducer blades and is connected to a strut that is supported by at least one bearing such that the shaft can rotate relative to the strut. 
     In another embodiment according to any of the previous embodiments, the strut is further supported by a thrust bearing such that the fan rotor can rotate relative to the strut. 
     In another embodiment according to any of the previous embodiments, the at least one inducer stage comprises a plurality of inducer stages coupled to the fan rotor. 
     In another embodiment according to any of the previous embodiments, each inducer state comprises one or more blades fixed for rotation with the fan rotor and a core stator fixed to a non-rotating engine structure, and wherein the core stator includes one or more vanes. 
     In another embodiment according to any of the previous embodiments, the compressor section includes at least a first compressor driven by the shaft and a second compressor driven by a second shaft and which rotates faster than the first compressor. 
     In another embodiment according to any of the previous embodiments, the first compressor is positioned immediately aft of the at least one inducer stage. 
     In another embodiment according to any of the previous embodiments, the at least one inducer stage is positioned forward of the speed reduction device. 
     In another embodiment according to any of the previous embodiments, a turbine section includes at least a first turbine that is driven by the shaft and a second turbine that is driven by the second shaft and rotates faster than the first turbine 
     In another embodiment according to any of the previous embodiments, the core inlet stator is a variable vane. 
     In another embodiment according to any of the previous embodiments, the core inlet stator is heated for anti-icing. 
     In another embodiment according to any of the previous embodiments, the one or more inducer blades are heated for anti-icing. 
     In another embodiment according to any of the previous embodiments, the turbine section includes three turbine rotors, with the first turbine including a first turbine rotor that drives the fan and the first compressor, and with the second turbine including a second turbine rotor that drives the second compressor, and with a third turbine rotor driving a third compressor. 
     In another embodiment according to any of the previous embodiments, a turbine section with a high pressure turbine and a low pressure turbine, and with each of the low pressure turbine and the high pressure turbine driving a compressor rotor of the compressor section. 
     In another embodiment according to any of the previous embodiments, the speed reduction device is positioned intermediate the fan and the compressor rotor driven by the low pressure turbine. 
     In another embodiment according to any of the previous embodiments, the speed reduction device is positioned intermediate the low pressure turbine and the compressor rotor driven by the low pressure turbine. 
     In another embodiment according to any of the previous embodiments, the speed change device comprises a gearbox including a sun gear in meshing engagement with star gears and a ring gear in meshing engagement with the star gears, and wherein the ring gear drives the fan, and wherein the gearbox has a gear reduction ratio of greater than about 2.3. 
     In another embodiment according to any of the previous embodiments, the speed change device comprises a gearbox including a sun gear in meshing engagement with a plurality of planetary gears supported by a planet carrier and a ring gear in meshing engagement with the planet gears, and wherein the ring gear is fixed and the planet carrier provides input to the fan, and wherein the gearbox has a gear reduction ratio of greater than about 2.3. 
     In another featured embodiment, a gas turbine engine comprises a first shaft defining an axis of rotation. A speed change device is driven by the first shaft. A fan section includes at least one fan blade coupled to a fan rotor wherein the fan rotor is driven by the speed change device. A compressor section comprises at least a low pressure compressor driven by the first shaft and a high pressure compressor driven by a second shaft. A turbine section comprises at least a low pressure turbine configured to drive the first shaft and a high pressure turbine configured to drive the second shaft. At least one inducer stage is positioned aft of the fan blade and coupled for rotation with the fan rotor about the axis of rotation, and wherein the at least one inducer stage is positioned forward of the compressor section. 
     In another embodiment according to the previous embodiment, the low pressure compressor is positioned aft of the speed change device. 
     In another embodiment according to any of the previous embodiments, the at least one inducer stage comprises one or more inducer blades fixed for rotation with the fan rotor and a core inlet stator fixed to a non-rotating engine structure, and wherein the core inlet stator is positioned radially outward of the inducer blades and is connected to a strut that is supported by at least one bearing such that the first shaft can rotate relative to the strut. 
     In another embodiment according to any of the previous embodiments, the at least one inducer stage comprises a plurality of inducer stages coupled to the fan rotor, and wherein each inducer stage comprises one or more inducer blades fixed for rotation with the fan rotor and a core inlet stator fixed to a non-rotating engine structure, and with the core stator including one or more vanes. 
     In another embodiment according to any of the previous embodiments, the at least one inducer stage is positioned forward of the speed change device. 
     In another embodiment according to any of the previous embodiments, the low pressure turbine is one of at least three turbine rotors, and with the low pressure turbine driving the fan via a first turbine rotor. A second turbine rotor drives the high pressure compressor. A third turbine rotor drives a third compressor. 
     In another embodiment according to any of the previous embodiments, the high pressure turbine drives a compressor rotor of the high pressure compressor and the low pressure turbine drives a compressor rotor of the low pressure compressor. 
     In another embodiment according to any of the previous embodiments, the speed reduction device is positioned intermediate the fan and the compressor rotor driven by the low pressure turbine. 
     In another embodiment according to any of the previous embodiments, the speed reduction device is positioned intermediate the low pressure turbine and the compressor rotor driven by the low pressure turbine. 
     These and other features of this application will be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically illustrates a gas turbine engine. 
         FIG. 2  is one example of a schematic representation of an engine upper half including an inducer section coupled to a fan. 
         FIG. 3  is another example of a schematic representation of an engine upper half including an inducer coupled to the fan before the core inlet stator. 
         FIG. 4  is another example of a schematic representation of an engine upper half including an inducer section coupled to a fan. 
         FIG. 5A  is a schematic representation of a speed reduction device comprises comprising a star gear type gearbox 
         FIG. 5B  is a schematic representation of a speed reduction device comprising a planetary gear type gearbox. 
         FIG. 6  shows another embodiment. 
         FIG. 7  shows yet another embodiment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed reduction device, such as a geared architecture  48  for example, to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A, which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm per hour of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram deg R)/518.7)^0.5]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. 
     A gas turbine engine  60  shown in  FIG. 2  includes a two-spool turbofan as described above, which generally incorporates a fan section  22 , a compressor section  24  with high  52  and low  44  pressure compressors, and a turbine section  28  with high  54  and low  46  pressure turbines. In this configuration, the low pressure turbine  46  is comprised of a plurality of stages. In the example shown, the low pressure turbine  46  includes a first stage  62 , a second stage  64 , and a third stage  66 . The high pressure turbine  54  is comprised of a first stage  68  and a second stage  70  that are positioned forward of the plurality of stages  62 ,  64 ,  66  of the low pressure turbine  46 . 
     Each of the stages for the high  54  and low  46  pressure turbines includes a plurality of blades coupled to a respective rotor. In the example shown, blades  71  of the low pressure turbine  46  are coupled to a first rotor  72  and blades  74  of the high pressure turbine  54  are coupled to a second rotor  76 . The first rotor  72  is configured to drive the low shaft  40  and the second rotor  76  is configured to drive the high shaft  50 . Each stage of the high  54  and low  46  pressure turbines also includes a plurality of vanes (not shown) interspersed with the blades where the vanes are mounted to a static engine structure  36 . 
     The high pressure compressor  52  is comprised of first  80 , second  82 , third  84 , and fourth  86  stages. The low pressure compressor  44  is comprised of first  88 , second  90 , third  92 , and fourth  94  stages that are positioned forward of the plurality of stages  80 ,  82 ,  84 ,  86  of the high pressure compressor  52 . Each of the stages of the high pressure compressor  52  includes a plurality of blades  96  that are coupled to a rotor  98  that is driven by the high shaft  50 . Each of the stages of the low pressure compressor  44  is comprised of blades  100  that are coupled to a rotor  102  that is driven by the low shaft  40 . Each stage of the high  52  and low  44  pressure compressors also includes a plurality of vanes (not shown) interspersed with the blades where the vanes are mounted to a static engine structure  36 . Various bearings  38  rotatably support the high  50  and low  40  shafts as known. 
     The fan section  22  includes a fan  42  that is driven by the geared architecture  48 . The fan  42  is comprised of a plurality of fan blades  104  that are coupled to a fan rotor  106  for rotation about the axis. The geared architecture  48  couples the low shaft  40  to the fan rotor  106  such that the fan rotor  106  rotates at a lower speed than the low shaft  40 . The geared architecture  48  is an epicyclic gear arrangement that includes a plurality of star or planet gears driven by a sun gear fixed for rotation with the low shaft  40 . The star gears drive a ring gear that is configured to drive the fan rotor  106 . 
     The engine  60  also includes an inducer section  110  that comprises a fan-tied compressor stage, i.e. the inducer section is an additional compressor stage that is connected to the fan rotor  106 . The inducer section  110  serves to efficiently feed the low pressure compressor  44  a more controlled/stabilized air flow. Alternatively, the inducer may allow the manufacturer to make one engine model without the inducer and use the exact same core on an engine model with the inducer. This approach would be a way of improving manufacturing efficiencies since the engine core with the inducer could in turn be used at higher thrust while maintaining the same peak core temperatures due to the higher air flows provided by the additional inducer stage or stages. In this way most of the engine part numbers would be common between the two models, thereby eliminating duplication in engineering work, development work, toolings, and other savings. 
     In one example, the inducer section  110  includes at least one inducer stage  112  that is driven by the fan rotor  106 . The inducer stage  112  comprises one or more blade rows  114  fixed for rotation with the fan rotor  106  and a core stator structure  116  fixed to the non-rotating static engine structure  36 . The core stator structure  116  is configured to facilitate reducing swirl coming off of the fan and diffusing the air flow. The core stator structure  116  includes one or more vanes  118  fixed to the static engine structure  36 . Additional support for the stator structure  116  is provided by a connection to a strut  120 . A bearing  122  rotatably supports the shaft  40  for rotation relative to the strut  120  and stator structure  116 . A thrust bearing  148  also provides support for the fan and inducer assembly. 
     In one example shown in  FIG. 2 , the core stator structure  116  is positioned axially between the fan  42  and the blades  114  of the inducer stage  112 . 
     In another example shown in  FIG. 3 , the core stator structure  116  is positioned aft of the blades  114  of the inducer stage  112 . 
     In another example shown in  FIG. 4 , the at least one inducer stage comprises a plurality of inducer stages  112 ,  130  coupled to the fan rotor  106 .  FIG. 4  shows an example configuration having a first inducer stage  112  and a second inducer stage  130  that is configured similarly to the first inducer stage  112 . The second inducer stage  130  comprises one or more blades  132  fixed for rotation with the fan rotor  106  and a core stator structure  116  fixed to the static engine structure  36  as described above. The core stator structure  116  can optionally include a second set of vanes  134  positioned aft of the blades  118  of the first inducer stage  112  and forward of the blades  132  of the second inducer stage  130 . Optionally the vanes could be positioned respectively aft of each set of blades in a manner similar to that shown in  FIG. 3 . 
     In one example, the geared architecture  48  comprises a gearbox.  FIGS. 5A and 5B  shows two different examples of gearboxes.  FIG. 5A  shows a gearbox that comprises a star gear configuration. This configuration includes a sun gear  150  in meshing engagement with star gears  152  and a ring gear  154  in meshing engagement with the star gears  152 . The sun gear  150  is driven by the shaft  40  and the ring gear  154  at drives the fan rotor  106 . 
       FIG. 5B  shows a gearbox that comprises a planetary gear configuration. This configuration includes a sun gear  160  in meshing engagement with a plurality of planetary gears  162  supported by a planet carrier  164  and a ring gear  166  in meshing engagement with the planet gears  162 . The ring gear  166  is fixed to a static structure  36  and the planet carrier  164  provides input to the fan rotor  106 . 
     In either of these configurations, the gearbox defines a gearbox axial center-plane P (see  FIG. 2-4 ). 
     In one example, the low pressure compressor  44  is positioned immediately aft of the inducer section  110  and the inducer section  110  is positioned forward of the gearbox axial center-plane P. 
     In one example, the low pressure compressor  44  is positioned aft of the gearbox axial center-plane P. 
     The various configurations described above provide a geared turbofan with a slow turning, fan-tied auxiliary compressor stage or stages, and a separate higher speed low pressure compressor/low pressure turbine that are tied to a common single shaft. The fan-tied low pressure compressor stage or stages provide an inducer that is connected to the fan rotor immediately aft of the fan. This provides several benefits. 
     For example, the configurations disclosed above improve engine operability by fractionally reducing the pressure rise required of the higher speed low pressure compressor and moving that fractional pressure rise to the lower speed fan rotor utilizing the associated inducer stage. Moving low pressure compressor stages from the higher speed low pressure compressor shaft to the slower rotating fan rotor improves engine operability by reducing the inertia of the higher speed low pressure compressor and turbine, which are tied to the same single shaft. Relative to a configuration that does not have an inducer section, the inertia of the stages in an inducer configuration is decreased by a factor of 1/GR2 where GR2 is the square of the speed reduction ratio of the gear. For example, if the gear ratio GR is 2, locating a stage as fan-tied reduces the inertia of that low compressor stage by the factor of 0.25 relative to locating the stage within the higher speed low pressure compressor. 
     In another example, one or more of the core inlet stator vanes  118  could be a variable vane that rotates along a spanwise axis (as schematically indicated by dashed arrow  170  in  FIG. 2 ) in order to better align the airfoil to the input flow. This would be especially desirable in reducing takeoff peak temperatures in the core. 
     In another example, the core inlet stator vanes  118  and/or the inducer blades  114 ,  132  could be heated for anti-icing conditions. A heating device is schematically shown at  174  in  FIG. 4 . 
     In one example, the fan  42  and inducer section  110  turn in the same direction as the fan drive turbine using a speed reduction device or a gearbox of the planet type as shown in  FIG. 5B . In this configuration the sun gear  160  provides the input torque, the ring gear  166  is fixed to the supporting structure  36  and the carrier  164  of the gears  162  between the sun  160  and the ring gear  166  is connected to the fan hub  106  and provides the rotational torque required by the fan  42 . 
     Further, for configurations where the gear is a star gear configuration that results in the fan rotating in an opposite direction to the rotational direction of the input shaft from the turbine (see  FIG. 5A ), the fan-tied low pressure compressor enables even more pressure to be addressed by the fan rotor  106  as the fan-tied low pressure compressor more easily accommodates more work being done by the fan rotor  106  than the counter rotating high speed low pressure compressor does without the presence of a fan-tied compressor stage. This also increases the supercharging temperature of the high speed low pressure compressor and, thus, results in a lower tip Mach number for the first rotor of the high speed low pressure compressor, resulting improved efficiency. 
       FIG. 6  shows an embodiment  200 , wherein there is a fan drive turbine  208  driving a shaft  206  to in turn drive a fan rotor  202 . A gear reduction  204  may be positioned between the fan drive turbine  208  and the fan rotor  202 . This gear reduction  204  may be structured and operate like the gear reduction disclosed above. A compressor rotor  210  is driven by an intermediate pressure turbine  212 , and a second stage compressor rotor  214  is driven by a turbine rotor  216 . A combustion section  218  is positioned intermediate the compressor rotor  214  and the turbine section  216 . 
       FIG. 7  shows yet another embodiment  300  wherein a fan rotor  302  and a first stage compressor  304  rotate at a common speed. The gear reduction  306  (which may be structured as disclosed above) is intermediate the compressor rotor  304  and a shaft  308  which is driven by a low pressure turbine section. 
     Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.