Patent Publication Number: US-2016222815-A1

Title: High Efficiency Geared Turbofan

Description:
RELATED APPLICATION 
     This application claims priority to provisional application 61/885145 filed on 1 Oct. 2013. 
    
    
     BACKGROUND OF THE INVENTION 
     This application relates to a geared turbofan gas turbine engine wherein the turbine efficiency is increased compared to the prior art. 
     Gas turbine engines as known include a fan delivering air into a compressor section. The compressor compresses the air and delivers the air into a combustor section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. 
     Historically, in one type of gas turbine engine there were two turbine rotors and two compressor rotors. A lower pressure turbine rotor drove a lower pressure compressor rotor and also drove the fan rotor at a single rate of speed. In another type of gas turbine engine, there were three turbine rotors with the first two turbines driving two compressor sections and a turbine rotor driving the fan rotor directly with no, or minimal attached compressor stages. 
     More recently, it has been proposed to include a gear reduction between the low pressure compressor and the fan rotor. 
     This has provided improvements in the operation of a gas turbine engine. 
     SUMMARY OF THE INVENTION 
     In a featured embodiment, a gas turbine engine comprises a fan drive turbine. The fan drive turbine drives the fan through a gear reduction; A change in enthalpy is defined across the gas turbine engine. The change in enthalpy divided by a speed of the fan drive turbine squared is less than or equal to about 1.8. An axial component of gases approaching an upstream most blade of the fan drive turbine divided by the speed of the fan drive turbine is equal to or less than about 0.9. 
     In another embodiment according to the previous embodiment, the fan drive turbine also drives ng a compressor rotor, and then drives the fan through the gear reduction such as the compressor rotor. The fan drive turbine rotates at a higher speed than the fan. 
     In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal to 2.6. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine has three to six stages. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine is at least partially manufactured from a directionally solidified material. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine includes at least one cooled blade. 
     In another embodiment according to any of the previous embodiments, the gear ratio is greater than or equal to 3.0. 
     In another embodiment according to any of the previous embodiments, the speed of the fan drive turbine is a mean line velocity measured in feet per second. The change in enthalpy is measured in joules. The axial component of the gases is measured in feet per second. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine is utilized in combination with at least two additional turbine rotors where each drive a compressor rotor. 
     In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal to 2.6. 
     In another embodiment according to any of the previous embodiments, the gear ratio is greater than or equal to 3.0. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine has three to six stages. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine is at least partially manufactured from a directionally solidified material. 
     In another embodiment according to any of the previous embodiments, at least one of the two additional turbine rotors is made at least in part from a single crystal material. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine includes at least one cooled blade. 
     In another embodiment according to any of the previous embodiments, the speed of the fan drive turbine is a mean line velocity measured in feet per second. The change in enthalpy is measured in joules. The axial component of the gases is measured in feet per second. 
     In another embodiment according to any of the previous embodiments, a gear ratio of the gear reduction is greater than or equal to 2.6. 
     In another embodiment according to any of the previous embodiments, the gear ratio is greater than or equal to 3.0. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine has three to six stages. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine is at least partially manufactured from a directionally solidified material. 
     In another embodiment according to any of the previous embodiments, the fan drive turbine includes ng at least one cooled blade. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a gas turbine engine. 
         FIG. 2  shows the boundaries for a high efficiency engine such as illustrated in  FIG. 1 . 
         FIG. 3  is a graph of work coefficient versus flow coefficient. 
         FIG. 4A  shows a low efficiency turbine rotor. 
         FIG. 4B  shows a higher efficiency turbine rotor. 
         FIG. 5  shows another embodiment. 
         FIG. 6  shows another embodiment. 
         FIG. 7  shows exemplary values for a plurality of quantities. 
     
    
    
     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 flow path B in a bypass duct defined within a nacelle  15 , 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 . Although depicted as a two-spool 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 two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. In another embodiment there are at least three turbines. Further, in another embodiment, there are at least two turbines in front of a fan drive turbine to reduce gas temperature entering the fan drive turbine. In another embodiment , the at least two turbines rotate independently. 
     The exemplary 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, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated 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 second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  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 mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . 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 mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     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  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 five. 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 five 5:1. In another embodiment, the bypass ratio is greater than 12.0. 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 or equal to about 2.6. 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 thrust set 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 burned per hour 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. 
       FIG. 2  is a graph of the quantities Δh/U 2  versus C x /U. As shown, turbine efficiency increases as one heads toward zero for both quantities. A low efficiency island is shown near the upper right corner of the graph and includes a turbine designed as shown in  FIG. 4A . The higher efficiency island is shown near the bottom left corner and would include turbine blades designed closer to that shown in  FIG. 4B . 
     As shown, by increasing U which is the speed of a fan drive turbine  46  driving the fan rotor  24 , turbine efficiency tends to be raised. 
     The turbine efficiency plot shown here includes the Ah which is a total enthalpy change. Enthalpy is a measure of the total energy of a thermodynamic system. It includes the system&#39;s internal energy or thermodynamic potential as well as its volume and pressure. The unit of measurement for enthalpy in the international system of units is the joule, but other historical conventional units are still in use, such as the British Thermal Unit and the calorie. 
     A work coefficient quantity is Δh/U 2  and is a parameter in which the dimensions are read out, such that it is dimension-less, and relates the turbine work to the mean wheel speed of the turbine U. The use of this work coefficient combined with a flow coefficient C x /U, shown in  FIG. 3 . The flow coefficient C x /U is a good indicator of a velocity triangle in a fan drive turbine. A low C x /U design can be characterized by high turning speed within an individual blade row and relatively low axial velocity. On the other hand, high C x /U designs tend toward low turning in the blade row and low camber airfoils. 
     As an example,  FIG. 4A  shows the nature of the low efficiency islands in  FIG. 2  (as shown by the general location  4 A in  FIG. 2 ) wherein the gas vector “A” coming off a stator  100  is poorly aligned to the tangential C t  and U directions as it approaches a rotor blade  101 . This inherently makes the gases approaching the rotor blade  101  more axial, which is undesirable and less tangential which is desirable. 
     On the other hand,  FIG. 4B  shows a high efficiency island, shown generally as  4 B in  FIG. 2 . Here, the stator  102  is designed in combination with a rotor  104  such that the gas coming off of the stator is closer to the tangential C t  and U directions. This inherently improves the presentation of velocity energy to the rotating blade  104 . That is, the U component is greater than in  FIG. 4A . Further, the great increase in U raises the stress in all areas of the fan drive turbine. If high efficiency and high U are desired, this may necessitate the use of materials with higher strength at a given temperature. The designer may also cool the turbine to increase the allowable stress with a weaker class of material. 
       FIG. 5  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 . 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 rotor  216 . This engine may operate as discussed with regard to  FIGS. 2, 3, 4B and 7 , as may the engine  20  of  FIG. 1 . 
     In addition, a directionally solidified material may be utilized for the fan drive turbine  46  or even a more temperature capable material such as a single crystal material or an internally cooled blade. Such materials may be used for at least one blade in the particular turbine. As shown in  FIG. 6 , a turbine blade  150  for use in the engine of  FIG. 1 or 5 , can have internal cooling passages  151 , as known. 
     Gear ratios equal to or above 2.6 for the gear reduction  48  may be utilized. In embodiments, the gear ratio may be equal to or above 3.0. 
     The fan rotor  42  may be sized to maximize the bypass ratio while minimizing fuel burn. 
       FIG. 7  shows a sample of turbine engines and a number of quantities. 
     The fan drive turbine in these example engines may have three to five low pressure turbine stages. 
     In the above referenced system, the C x /U quantity averages out to 0.49. C x  is an average axial velocity taken in feet/second and U is the rotor speed at a mean line velocity in feet per second. The stage loading quantity dh/U 2  average equals 1.27. Both of the flow coefficient and work coefficient quantities are non-dimensional. What is referred to here as dh/U 2  is really gJdh/U 2  where g equals 32.2 feet pounds per minute/second squared per lbf and J equals 778-feet lbf/btu. dh equals a change in specific enthalpy across the turbine measured in btu/lbm and U equals a rotor speed at a mean radius in feet per second. 
     As can be appreciated from  FIG. 2  and by the boundary  300 , Applicant has designed gas turbine engines wherein a fan drive turbine drives a fan through a gear reduction, wherein a change in enthalpy divided by a speed of the fan drive turbine squared is less than or equal to about 1.8. In addition, an axial component of the gas approaching the upstream most blade of the fan drive turbine divided by the speed of the fan drive turbine is equal to or less than about 0.9. 
     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.