Abstract:
A gas turbine engine according to an example of the present disclosure includes, among other things, a fan configured to deliver airflow to a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The fan has a fan diameter, Dfan, and the low pressure turbine section has a turbine diameter, Dturb.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application is a continuation of U.S. patent application Ser. No. 14/428,049, which was filed on Mar. 15, 2015, which is a National Stage Entry of PCT Application No. PCT/US2013/025470, filed on Feb. 10, 2013, which claims priority to U.S. Provisional Application Ser. No. 61/708,288, filed Oct. 1, 2012. 
     
    
     BACKGROUND 
       [0002]    A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
         [0003]    The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the 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 includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction. 
         [0004]    A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly 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 geared architectures have improved propulsive efficiency, 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 fan configured to deliver airflow to a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The fan and the low pressure turbine section are configured to rotate at a common speed and in a common direction. The gas turbine engine has a bypass ratio of greater than about 10. The fan has a fan diameter, Dfan, and the low pressure turbine section has a turbine diameter, Dturb. The fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65. 
         [0007]    In a further embodiment of any of the foregoing embodiments, the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section. 
         [0008]    In a further embodiment of any of the foregoing embodiments, the fan includes a plurality of fan blades, and the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades. 
         [0009]    In a further embodiment of any of the foregoing embodiments, the high pressure turbine section includes two stages. 
         [0010]    In a further embodiment of any of the foregoing embodiments, the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section. 
         [0011]    In a further embodiment of any of the foregoing embodiments, the low pressure turbine section has a greater number of stages than the low pressure compressor section. 
         [0012]    In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.45. 
         [0013]    In a further embodiment of any of the foregoing embodiments, the fan includes fewer than 20 fan blades. 
         [0014]    In a further embodiment of any of the foregoing embodiments, the low pressure turbine section has a pressure ratio of greater than about 5. 
         [0015]    A gas turbine engine according to an example of the present disclosure includes a fan having fewer than 20 fan blades situated at an inlet of a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The low pressure turbine section has a greater number of stages than the low pressure compressor section. The fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65. 
         [0016]    In a further embodiment of any of the foregoing embodiments, the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section. 
         [0017]    In a further embodiment of any of the foregoing embodiments, the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades. 
         [0018]    In a further embodiment of any of the foregoing embodiments, the high pressure turbine section includes two stages. 
         [0019]    In a further embodiment of any of the foregoing embodiments, the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section. 
         [0020]    In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.5. 
         [0021]    In a further embodiment of any of the foregoing embodiments, the gas turbine engine has a bypass ratio of greater than about 10. 
         [0022]    In a further embodiment of any of the foregoing embodiments, the low pressure turbine has a pressure ratio of greater than about 5. 
         [0023]    A method of designing a gas turbine engine according to an example of the present disclosure includes providing a fan configured to deliver airflow to a bypass duct, and providing a core engine configured to rotate the fan. The core engine include a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to rotate the fan at a common speed and in a common direction. The fan has a fan diameter, Dfan, the low pressure turbine section has a diameter, Dturb, and the fan diameter Dfan and the low turbine section diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65. The fan is configured to deliver a portion of air into the core engine, and a portion of air into the bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the core engine, is greater than 10. 
         [0024]    In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.5 and the low pressure turbine has a pressure ratio of greater than about 5. 
         [0025]    A further embodiment of any of the foregoing embodiments includes a low pressure compressor section driven by the low pressure turbine section. The low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section, and the high pressure turbine section includes two stages. 
         [0026]    In a featured embodiment, a gas turbine engine has a propulsor including a fan and a fan drive geared architecture. The fan defines a fan diameter. A gas generator includes a fan drive turbine, which drives the fan through the fan drive geared architecture. The fan drive turbine has a diameter less than 0.50 the size of the fan diameter. 
         [0027]    In another embodiment according to the previous embodiment, the diameter of the fan drive turbine is greater than 0.30 the size of the fan diameter. 
         [0028]    In another embodiment according to any of the previous embodiments, the diameter of the fan drive turbine is between about 0.35 and about 0.45 the size of the fan diameter. 
         [0029]    In another embodiment according to any of the previous embodiments, the fan drive turbine further comprises a high pressure turbine located upstream of the low pressure turbine. 
         [0030]    In another embodiment according to any of the previous embodiments, the fan drive turbine comprises a low pressure turbine. 
         [0031]    In another embodiment according to any of the previous embodiments, a compressor section has a low pressure compressor driven by the low pressure turbine and a combustor in fluid communication with the compressor section. 
         [0032]    In another embodiment according to any of the previous embodiments, a first shaft connects the low pressure turbine, low pressure compressor, and the fan drive geared architecture. 
         [0033]    In another embodiment according to any of the previous embodiments, the fan drive geared architecture comprises an epicyclic gear box. 
         [0034]    In another embodiment according to any of the previous embodiments, the diameter of the fan drive turbine is defined by an outer case surface of the fan drive turbine. 
         [0035]    In another embodiment according to any of the previous embodiments, the fan diameter is defined by an outer peripheral surface of the fan blades. 
         [0036]    In another embodiment according to any of the previous embodiments, an engine case surrounds the gas generator. The engine case includes at least one pylon mount interface for attachment to a pylon mounted underneath a wing. 
         [0037]    In another featured embodiment, a gas turbine engine has a propulsor including a fan and a fan drive geared architecture. The fan defines a fan diameter. A gas generator includes a fan drive turbine, which drives the fan through the fan drive geared architecture. The fan drive turbine has a diameter between about 0.35 and about 0.45 the size of the fan diameter. 
         [0038]    In another embodiment according to the previous embodiment, the fan drive geared architecture has a gear reduction ratio of greater than about 2.3 
         [0039]    In another embodiment according to any of the previous embodiments, the fan drive geared architecture comprises an epicyclic gear box. 
         [0040]    In another embodiment according to any of the previous embodiments, a compressor section has at least a first compressor and a second compressor, a combustor in fluid communication with the compressor section, and at least one additional turbine. A first shaft connects the fan drive turbine and the first compressor and a second shaft connects the second compressor and the one additional turbine. 
         [0041]    In another embodiment according to any of the previous embodiments, the second shaft rotates at a faster speed than the first shaft. 
         [0042]    In another embodiment according to any of the previous embodiments, the fan drive turbine comprises a low pressure turbine and the one additional turbine comprises a high pressure turbine. 
         [0043]    In another embodiment according to any of the previous embodiments, the fan drive geared architecture couples the first shaft to the fan at a location upstream of the compressor section. 
         [0044]    In another embodiment according to any of the previous embodiments, an engine case surrounds the gas generator. The engine case includes at least one pylon mount interface for attachment to a pylon mounted underneath a wing. 
         [0045]    In another embodiment according to any of the previous embodiments, the pylon mount interface comprises at least a front mount beam and a rear mount beam located aft of the front mount beam. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0046]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0047]      FIG. 1  schematically illustrates a geared turbofan engine embodiment. 
           [0048]      FIG. 2  schematically illustrates a direct drive turbine engine embodiment. 
           [0049]      FIG. 3  shows a side view of a geared turbofan embodiment in one example mounting configuration. 
           [0050]      FIG. 4  shows an end view of  FIG. 3  in an aft direction looking forward. 
       
    
    
     DETAILED DESCRIPTION 
       [0051]      FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
         [0052]    Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, 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; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
         [0053]    The example 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. 
         [0054]    The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such as a geared architecture  48 , 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 (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
         [0055]    A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
         [0056]    The example low pressure turbine  46  has a pressure ratio that is greater than about  5 . The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
         [0057]    A mid-turbine frame  58  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
         [0058]    The core airflow C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes vanes  60 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  60  of the mid-turbine frame  58  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  58 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
         [0059]    The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
         [0060]    In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten ( 10 : 1 ) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
         [0061]    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 (‘TSFCT’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
         [0062]    “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 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
         [0063]    “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. 
         [0064]    The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section  22  includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about  6  turbine rotors schematically indicated at  34 . In another non-limiting example embodiment the low pressure turbine  46  includes about 3 turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
         [0065]    The configuration shown in  FIG. 2  is a direct drive turbine engine  25 . The direct drive turbine engine  25  includes 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′ while the compressor section  24 ′ draws air in along a core flow path C′ where air is compressed and communicated to the combustor section  26 ′. In the combustor section  26 ′, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28 ′ where energy is extracted and utilized to drive the fan section  22 ′ and the compressor section  24 ′. 
         [0066]    The direct drive turbine engine  25  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 via several bearing systems  38 ′. The low speed spool  30 ′ generally includes an inner shaft that connects a fan  42 ′ having a plurality of blades and a low pressure (or first) compressor section  44 ′ to a low pressure (or first) turbine section  46 ′. The inner shaft or low speed spool  30 ′ directly drives the fan  42 ′, that is, the fan  42 ′ and low pressure turbine section  46 ′ are driven at the same speed. The high-speed spool  32 ′ includes an outer shaft that interconnects a high pressure (or second) compressor section  52 ′ and a high pressure (or second) turbine section  54 ′. The inner shaft and the outer shaft are concentric and rotate via the bearing systems  38 ′ about the engine central longitudinal axis A′. 
         [0067]    In the direct drive configuration shown in  FIG. 2 , a fan drive turbine directly drives the fan section  22 ′, i.e. there is no geared architecture in this configuration. In  FIG. 2 , the fan drive turbine comprises the low pressure turbine  46 ′ which is coupled to directly drive the fan  42 ′. 
         [0068]    The geared architecture configuration has increased efficiency that enables the use and fabrication of a smaller low pressure turbine  46  both in diameter and in the number or overall stages as compared to the direct drive turbine engine  25  ( FIG. 2 ), which must rotate at a less efficient speed. 
         [0069]    Moreover, the smaller, more efficient low pressure turbine  46  of the geared turbofan engine  20  enables alternate and more efficient mounting configurations. Space limitations for wing mounted engines result from a minimum distance between a bottom of an engine and the runway. Larger landing gear components can be utilized to raise the aircraft and thereby the engine relative to the runway, but larger landing gear components are not a desirable option due to significant weight penalties. Accordingly, as the propulsor fan section  22  grows in size, the mounting options decrease. For engines having the same fan section diameter, the fan drive turbine section of the direct drive engine  25  ( FIG. 2 ) is much larger than the fan drive turbine section of a geared turbofan engine  20  ( FIG. 1 ). 
         [0070]    This difference becomes significant when defining a mounting configuration for the engine. The core engine section including the fan drive turbine section can be mounted under the wing, with the fan section extending forward of the wing. The larger fan drive turbine section of a direct drive turbine requires that the engine centerline be spaced a further distance from a bottom surface of the wing as compared to a centerline of a geared turbofan engine with the smaller more efficient fan drive turbine. Even modest reductions in this spacing can enable significant weight savings in smaller landing gear lengths and structures. 
         [0071]    The example geared turbofan engine  20  includes a fan diameter  62  ( FIG. 1 ) and an example direct drive engine  25  includes a fan diameter  64  ( FIG. 2 ). In one example configuration, both the fan diameter  62  of the geared turbofan engine  20  and the fan diameter  64  of the direct drive turbine engine  25  are of a common size. Further, in this example, the fan pressure ratio and overall pressure ratio through the core are the same. When these fan diameters  62 ,  64  and pressure ratios are the same, the geared turbofan engine  20  includes a fan drive turbine diameter  66  ( FIG. 1 ) that is much smaller than a diameter  68  ( FIG. 2 ) of the fan drive turbine for the direct drive engine  25 . In one example, for a common fan diameter, the fan drive turbine is about 0.35 to about 0.45 the diameter  62  of the fan  42 , wherein a corresponding direct drive engine  25  would include a fan drive turbine between about 0.50 and 0.65 the diameter  64  of the fan  42 ′. 
         [0072]      FIGS. 3-4  show the geared turbofan engine  20  in one example mount configuration. A front mount beam  70  and a rear mount beam  72  are used to connect the engine case  74  to a pylon  76  that is mounted underneath a wing. One relatively important dimension, indicated at  80 , is the distance between a bottom surface  82  of the wing and an outermost surface  84  of the fan drive turbine section, that is, low pressure turbine  46 . For a fan diameter  64  ( FIG. 2 ) that is the same as the fan diameter  62  for the geared turbofan engine  20  in  FIGS. 3-4 , the fan drive turbine section, that is, low pressure turbine  46 ′, would have a comparatively greater size as indicated by an outermost surface  78  of the low pressure turbine  46 ′. The increased turbine size for the direct drive configuration decreases the wing clearance dimension  80 ′ when compared to the dimension  80  of the geared turbo fan engine  20 . 
         [0073]    Thus, the significance of the difference in size of the two different fan drive turbine sections is illustrated with the required spacing of the critical dimension  80 ′ for a direct drive turbine indicated between the outermost surface  78 , shown by the dashed lines, and the bottom surface  82  of the wing. Accordingly, the size of the fan  42 ′ for a direct drive turbine engine  25  is limited by the size of the fan drive turbine, i.e. the size of the low pressure turbine  46 ′. As such, the geared turbofan engine  20  with the smaller more efficient fan drive turbine, i.e. low pressure turbine  46 , can provide a larger fan in the same space, and/or enable a fan size not possible in a direct drive gas turbine engine  25 . 
         [0074]    Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.