Abstract:
Please replace the abstract with the following rewritten abstract. No new matter has been added. 
     An example gas turbine engine includes, among other things, a geared architecture rotatably coupling a fan drive shaft to an engine fan, the geared architecture having a speed reduction ratio that is greater than or equal to 2.4. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to primary stream exhaust velocity, is approximately in a range of 0.75 to 0.90.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application is the U.S. National Phase of PCT/US2013/022402, filed Jan. 21, 2013. 
     
    
     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-temperature exhaust gas flow. The high-temperature 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 low inner shaft. 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. 
         [0004]    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 
       [0005]    A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a geared architecture rotatably coupling a fan drive shaft to an engine fan, the geared architecture having a speed reduction ratio that is greater than or equal to 2.4. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to primary stream exhaust velocity, is approximately in a range of 0.75 to 0.90. 
         [0006]    In a further non-limiting embodiment of the foregoing gas turbine engine, the engine may be configured so that the Exhaust Velocity Ratio is in the range when cruising at 35,000 feet and when operating at a 0.80 Mach number cruise power condition. 
         [0007]    In a further non-limiting embodiment of either of the foregoing gas turbine engines, a Fan Pressure Ratio for the engine may be less than 1.45 at 35,000 feet and when operating at a 0.80 Mach number cruise power condition. 
         [0008]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, a Bypass Ratio of the engine may be greater than 8.0. 
         [0009]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the engine may be configured so that the Exhaust Velocity Ratio is in the range when the fan stream exhaust velocity is less than 1175 feet per second. 
         [0010]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the speed reduction ratio is less than or equal to 4.2. 
         [0011]    A gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, a geared architecture rotatably coupling a fan drive shaft to a fan of a gas turbine engine, the geared architecture having a speed reduction ratio that is greater than or equal to 2.4, a fan stream exhaust of the gas turbine engine, and a primary stream exhaust of the gas turbine engine. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to primary stream exhaust velocity, is approximately in a range of 0.75 to 0.90. 
         [0012]    In a further non-limiting embodiment of the foregoing gas turbine engine, the engine may be configured so that the Exhaust Velocity Ratio is in the range when cruising at 35,000 feet and when operating at a 0.80 Mach number cruise power condition. 
         [0013]    In a further non-limiting embodiment of either of the foregoing gas turbine engines, a Fan Pressure Ratio for the engine may be less than 1.45 at 35,000 feet and when operating at a 0.80 Mach number cruise power condition. 
         [0014]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, a Bypass Ratio of the engine may be greater than 8.0. 
         [0015]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the engine may be configured so that the Exhaust Velocity Ratio is in the range when the fan stream exhaust velocity is less than 1175 feet per second. 
         [0016]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the speed reduction ratio is less than or equal to 4.2. 
         [0017]    A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a geared architecture rotatably coupling a fan drive shaft to an engine fan, the geared architecture having a speed reduction ratio that is less than or equal to 4.2. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to primary stream exhaust velocity, is approximately in a range of 0.75 to 0.90. 
         [0018]    In a further non-limiting embodiment of the foregoing gas turbine engine, the engine may be configured so that the Exhaust Velocity Ratio is in the range when cruising at 35,000 feet and when operating at a 0.80 Mach number cruise power condition. 
         [0019]    In a further non-limiting embodiment of either of the foregoing gas turbine engines, a Fan Pressure Ratio for the engine may be less than 1.45 at 35,000 feet and when operating at a 0.80 Mach number cruise power condition. 
         [0020]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, a Bypass Ratio of the engine may be greater than 8.0. 
         [0021]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the engine may be configured so that the Exhaust Velocity Ratio is in the range when the fan stream exhaust velocity is less than 1175 feet per second. 
         [0022]    In a further non-limiting embodiment of any of the forgoing gas turbine engines, the speed reduction ratio is greater than or equal to 2.4. 
         [0023]    A gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, a geared architecture rotatably coupling a fan drive shaft to a fan of a gas turbine engine, the geared architecture having a speed reduction ratio that is less than or equal to 4.2, a fan stream exhaust of the gas turbine engine, and a primary stream exhaust of the gas turbine engine. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to primary stream exhaust velocity, is approximately in a range of 0.75 to 0.90. 
         [0024]    In a further non-limiting embodiment of the foregoing gas turbine engine, the engine may be configured so that the Exhaust Velocity Ratio is in the range when cruising at 35,000 feet and when operating at a 0.80 Mach number cruise power condition. 
         [0025]    In a further non-limiting embodiment of either of the foregoing gas turbine engines, a Fan Pressure Ratio for the engine may be less than 1.45 at 35,000 feet and when operating at a 0.80 Mach number cruise power condition. 
         [0026]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, a Bypass Ratio of the engine may be greater than 8.0. 
         [0027]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the engine may be configured so that the Exhaust Velocity Ratio is in the range when the fan stream exhaust velocity is less than 1175 feet per second. 
         [0028]    In a further non-limiting embodiment of any of the forgoing gas turbine engines, the speed reduction ratio is greater than or equal to 2.4. 
         [0029]    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. 
     
    
     
       DESCRIPTION OF THE FIGURES 
         [0030]    The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the detailed description. The figures that accompany the detailed description can be briefly described as follows: 
           [0031]      FIG. 1  shows a section view of an example gas turbine engine. 
           [0032]      FIG. 2  shows an example embodiment of the gas turbine engine of  FIG. 1 . 
       
    
    
     DETAILED DESCRIPTION 
       [0033]      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 temperature 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 . 
         [0034]    Although the disclosed non-limiting embodiment depicts a gas turbine gas turbine engine, it should be understood that the concepts described herein are not limited to use with gas turbines 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. 
         [0035]    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. 
         [0036]    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. 
         [0037]    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. 
         [0038]    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. 
         [0039]    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 . 
         [0040]    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 temperature 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 may 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. 
         [0041]    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. 
         [0042]    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. 
         [0043]    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 pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
         [0044]    “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. 
         [0045]    “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. 
         [0046]    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 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 in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
         [0047]    Referring now to  FIG. 2 , an engine  62  is a variation of the engine  20 . The engine  62  has a fan stream exhaust  64  and a primary stream exhaust  68 . Generally, the bypass flow B exits through the fan stream exhaust  64 , and the core flow C exits through the primary stream exhaust  68 . 
         [0048]    The fan stream exhaust  64  is provided between a nacelle nozzle  72  (at an aft area of a nacelle  74 ) and an engine casing  78 . The primary stream exhaust  68  is provided between a casing nozzle  82  (at an aft area of the casing  78 ) and a tailcone  84 . Flow through the primary stream exhaust  68  has been expanded through the low pressure turbine  46 . 
         [0049]    During operation, a ratio of a velocity of flow through the fan stream exhaust  64  to a velocity of flow through the primary stream exhaust  68 , termed the “Engine Exhaust Stream Velocity Ratio,” (or the “Exhaust Velocity Ratio”) is in a range from approximately 0.75 to 0.90. Particularly in geared engine designs having a speed reduction ratio of from 2.4 to 4.2, an Exhaust Velocity Ratio in the stated, desired range has been found to reduce overall fuel consumption compared to engines having this relationship falling outside of this range. 
         [0050]    The geometries of the engine  62  and the nacelle  74  could be selected to achieve the stated range for the Exhaust Velocity Ratio during cruise operation. The fan pressure ratio, total fan inlet flow, and bypass ratio could be selected to achieve the desired Exhaust Velocity Ratio. 
         [0051]    Changes to the fan pressure ratio could be achieved by changing the geometry of the blades of the fan  42 . The fan stream nozzle throat area is the minimum flow area at the exit of the fan nozzle  72 . The primary stream nozzle throat area is the minimum flow area at the exit of the primary nozzle  82  over the tail cone  84 . These areas could be designed for values to achieve a selected total fan flow and bypass ratio. Selection of a combination of these geometries would cause the engine  62  to operate at the desired Exhaust Velocity Ratio. 
         [0052]    Notably, an engine designed to operate within the stated envelope for the Exhaust Velocity Ratio falls within the scope of the disclosure, even if the engine is not continuously operating within that envelop. A person having skill in this art and the benefit of this disclosure could calculate, for example, an engine exhaust stream velocity ratio during a particular operating condition based on the designed fan stream exhaust and other parameters. 
         [0053]    In one example, the engine  62  exhibits a relationship of a fan stream exhaust velocity to primary stream exhaust velocity within this range when the engine  62  is cruising at 35,000 feet and operating at a 0.80 Mach number cruise power condition. Probes  86  and  90  may be located at or near the fan stream exhaust  64  and the primary stream exhaust  68  to measure the respective pressure and temperature of the flows, from which exhaust velocities can be determined in order to verify that the designed fan and primary stream exhausts result in the desired ratio. 
         [0054]    One characteristic of the engine  62  is that a fan pressure ratio of the engine  62  is less than 1.45 when the engine  62  is cruising at 35,000 feet and operating at a 0.80 Mach number cruise power condition. 
         [0055]    Another characteristic of the engine  62  is that a designed bypass ratio of the engine  62  is greater than 8.0. Flow need not be actively moving through the engine  62  for the engine  62  to have a designed bypass ratio that is greater than 8.0. 
         [0056]    Yet another characteristic of the engine  62  is that the geared architecture  48  has a speed reduction ratio of from 2.4 to 4.2. 
         [0057]    In one example, the fan stream exhaust velocity is less than 1175 ft/s (358 m/s) when the engine is cruising at 35,000 feet and operating at a 0.80 Mach number cruise power condition. 
         [0058]    Features of the disclosed examples include a fan stream to primary stream exhaust velocity relationship that advantageously results in reduced fuel consumption by improving propulsive efficiency and overall engine efficiency. 
         [0059]    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 this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.