Abstract:
A gas turbine engine according to an example of the present disclosure includes a drive turbine configured to drive a fan section, a combustor section located axially upstream of the drive turbine, and a speed change mechanism located axially downstream of the combustor section and axially upstream of the drive turbine. An output of the speed change mechanism connects to the fan.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/919,831, which was filed on Dec. 23, 2013 and is incorporated herein by reference. 
     
    
     BACKGROUND 
       [0002]    A gas turbine engine may include a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is typically compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow typically expands through the turbine section to drive the compressor and the fan section. Among other variations, the compressor section can include low and high pressure compressors, and the turbine section can include low and high pressure turbines. 
         [0003]    Typically, a high pressure turbine drives a high pressure compressor through an outer shaft to form a high spool, and a low pressure turbine drives a low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the inner shaft. A direct drive gas turbine engine may include a fan section driven by the low spool such that a low pressure compressor, low pressure turbine, and fan section rotate at a common speed in a common direction. 
         [0004]    A speed reduction device, which may be a fan drive gear system or other mechanism, may be utilized to drive the fan section such that the fan section rotates at a different speed than the turbine section. This allows for an overall increase in propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the speed reduction device that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds. 
         [0005]    Although gas turbine engines utilizing speed change mechanisms are generally known to be capable of improved propulsive efficiency relative to conventional engines, gas turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies. 
       SUMMARY 
       [0006]    A gas turbine engine according to an example of the present disclosure includes a drive turbine configured to drive a fan section, a combustor section located axially upstream of the drive turbine, and a speed change mechanism located axially downstream of the combustor section and axially upstream of the drive turbine. An output of the speed change mechanism connects to the fan. 
         [0007]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism is an epicyclical gearbox. 
         [0008]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism includes a ring gear connected to a fan drive shaft. 
         [0009]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism includes a planetary carrier connected to a fan drive shaft. 
         [0010]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism includes a sun gear connected to the drive turbine. 
         [0011]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism is a planetary gear system. 
         [0012]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism is a star gear system. 
         [0013]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism is located immediately upstream of the drive turbine. 
         [0014]    A further embodiment of any of the foregoing embodiments includes a fan drive shaft connected to the fan and a low pressure compressor connected to the fan drive shaft. 
         [0015]    A further embodiment of any of the foregoing embodiments includes a first compressor immediately downstream of the fan section and immediately upstream of the combustor section. 
         [0016]    A further embodiment of any of the foregoing embodiments includes a high pressure turbine and an intermediate turbine, wherein the speed change mechanism is located immediately downstream of the intermediate turbine and upstream of the drive turbine. 
         [0017]    In a further embodiment of any of the foregoing embodiments, the high pressure turbine is axially upstream of the intermediate turbine. 
         [0018]    A further embodiment of any of the foregoing embodiments includes a low pressure compressor and a high pressure compressor, wherein the low pressure compressor is connected with the intermediate turbine and the high pressure compressor is connected with the high pressure turbine. 
         [0019]    In a further embodiment of any of the foregoing embodiments, the gas turbine engine is a three spool gas turbine engine. 
         [0020]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism is an epicyclical gearbox. 
         [0021]    A method of operating a gas turbine engine according to an example of the present disclosure includes rotating a fan drive turbine to create a first rotational speed and reducing the first rotational speed output to a second rotational speed axially downstream of a second turbine section. 
         [0022]    In a further embodiment of any of the foregoing embodiments, a speed change mechanism reduces the first rotational speed of the fan drive turbine to the second rotational speed. 
         [0023]    In a further embodiment of any of the foregoing embodiments, the speed change mechanism is an epicyclical gearbox. 
         [0024]    A further embodiment of any of the foregoing embodiments includes rotating a fan and a compressor at the second rotational speed. 
         [0025]    In a further embodiment of any of the foregoing embodiments, the gas turbine engine is a three spool gas turbine engine. 
         [0026]    The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]      FIG. 1  illustrates a schematic view of a gas turbine engine according to an example embodiment. 
           [0028]      FIG. 2  illustrates a schematic view of a gas turbine engine according to another example embodiment. 
           [0029]      FIG. 3  illustrates a schematic view of a gas turbine according to yet another example embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0030]      FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26 , and a turbine section  28 . The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle or housing  21 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . 
         [0031]    The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  40  mounted for rotation about an engine central longitudinal axis A. The low speed spool  30  generally includes an inner shaft  31 . The inner shaft  31  interconnects a fan  32 , a first (or low) pressure compressor  34  with a first (or low) pressure turbine  36  through a speed change mechanism, which in the exemplary gas turbine engine  20  is illustrated as a geared architecture  38 . In one example, the low pressure turbine  36  is attached to a sun gear of the geared architecture  38  and the fan  32  and the low pressure compressor  34  are attached to a ring gear of the geared architecture  38  via shaft  31 , which extends through shaft  42 . 
         [0032]    A high speed spool  40  includes an outer shaft  42  that interconnects a second (or high) pressure compressor  44  and a second (or high) pressure turbine  46 . A combustor  48  is arranged in the exemplary gas turbine engine  20  axially between the high pressure compressor  44  and the high pressure turbine  46 . The inner shaft  31  and the outer shaft  42  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0033]    The low pressure turbine  36  rotates at a first rotational speed and the geared architecture  38  reduces the first rotational speed to a second rotational speed axially downstream of the high pressure turbine  46 . The fan  32  and the low pressure compressor  34  both rotate at the second rotational speed. The high pressure compressor  44  and the high pressure turbine  46  rotate at a third rotational speed different from the first and second rotational speed. 
         [0034]    The core airflow C is compressed by the low pressure compressor  34  and the high pressure compressor  44 , mixed and burned with fuel in the combustor  48  then expanded over the high pressure turbine  46  and the low pressure turbine  36 . The turbines  46  and  36  rotationally drive the respective high speed spool  40  and low speed spool  30  in response to the expansion. 
         [0035]    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 about ten (10), the geared architecture  38  is an epicyclic gear train such as a planetary gear system, a star gear system, or other gear system. The gear reduction ratio of greater than about 2.3 at the low pressure turbine  36 . 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 (&#39;TSFC&#39;)”—is the industry standard parameter of lbm 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 ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. The Low Corrected Fan Tip Speed according to another non-limiting embodiment of the example gas turbine engine  20  is less than about 1400 ft/second. 
         [0036]    An overall pressure ratio is a pressure ratio between a leading edge of a fan blade of the fan  32  and the trailing edge of the compressor section  24 . In one non-limiting embodiment, the overall pressure ratio is greater than approximately 30. In another non-limiting embodiment, the overall pressure ratio is less than approximately 70. 
         [0037]      FIG. 2  illustrates another example gas turbine engine  120 . The gas turbine engine  120  is substantially similar to the gas turbine engine  20  of  FIG. 1  except where shown in  FIG. 2  or discussed below. The example gas turbine engine  120  is disclosed herein as a two spool turbofan that generally incorporates a fan section  122 , a compressor section  124 , a combustor section  126 , and a turbine section  128 . 
         [0038]    The exemplary gas turbine engine  120  generally includes a low speed spool  130  and a high speed spool  140  mounted for rotation about an engine central longitudinal axis A. The low speed spool  130  generally includes an inner shaft  131 . The inner shaft  131  interconnects a fan  132  and a low pressure turbine  136  through a speed change mechanism  138 , such as an epicyclical gearbox that drives the fan  132  via a ring gear. The high speed spool  140  includes an outer shaft  142  that interconnects a high pressure compressor  144  and a high pressure turbine  146 . 
         [0039]    A combustor  148  is arranged in the exemplary gas turbine engine  120  between the high pressure compressor  144  and the high pressure turbine  146 . The inner shaft  131  and the outer shaft  142  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0040]    The low pressure turbine  136  rotates at a first rotational speed and the geared architecture  138  reduces the first rotational speed to a second rotational speed axially downstream of the high pressure turbine  146 . The fan  132  rotates at the second rotational speed. The high pressure compressor  144  and the high pressure turbine  146  rotate at a third rotational speed different from the first and second rotational speed. 
         [0041]      FIG. 3  illustrates yet another example gas turbine engine  220 . The gas turbine engine  220  is generally the same as the gas turbine engine shown in  FIG. 1  except where shown in  FIG. 3  or discussed below. The gas turbine engine  220  is disclosed herein as a three spool turbofan that generally incorporates a fan section  222 , a compressor section  224 , a combustor section  226 , and a turbine section  228 . 
         [0042]    The exemplary gas turbine engine  220  generally includes a low speed spool  230 , an intermediate spool  260 , and a high speed spool  240  mounted for rotation about an engine central longitudinal axis A. The low speed spool  230  generally includes an inner shaft  231  that interconnects a fan  232  and a speed change mechanism  238 , such as an epicyclic gearbox that drives the fan  132  via a ring gear. The intermediate spool  260  generally includes an intermediate shaft  262  that interconnects a low pressure compressor  234  and an intermediate pressure turbine  264 . The high speed spool  240  includes an outer shaft  242  that interconnects a high pressure compressor  244  and a high pressure turbine  246 . 
         [0043]    A combustor  256  is arranged in the exemplary gas turbine engine  220  between the high pressure compressor  244  and the high pressure turbine  246 . The inner shaft  231 , the intermediate shaft  262 , and the outer shaft  242  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0044]    A low pressure turbine  236  rotates at a first rotational speed and the geared architecture  238  reduces the first rotational speed to a second rotational speed axially downstream of the intermediate turbine  264 . The fan  232  rotates at the second rotational speed. The intermediate shaft  262  rotates at a third rotational speed different than the first or second rotational speed. The outer shaft  242  rotates at a fourth rotational speed different than the first, second, and third rotational speeds. 
         [0045]    The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. For example, in each of the foregoing embodiments, the ring gear of the speed change mechanism  38 ,  138 ,  238  (which would apply in the context of a star epicyclic gearbox) could be replaced with a carrier (which would apply in the context of a planetary epicyclic gearbox). The scope of legal protection given to this disclosure can only be determined by studying the following claims.