Abstract:
A gas turbine engine comprises a fan and a turbine having a fan drive rotor. There is also a second turbine rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive rotor. The fan drive rotor has a number of turbine blades in at least one of a plurality of rows of the fan drive rotor, and the turbine blades operate at least some of the time at a rotational speed. The number of turbine blades in the at least one row and the rotational speed are such that the following formula holds true for the at least one row of the fan drive turbine: (number of blades×speed)/60≧5500 Hz. The rotational speed is in revolutions per minute. A method of designing a gas turbine engine, and a turbine module are also disclosed.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application is a continuation-in-part of U.S. patent application Ser. No. 14/248,386, filed Apr. 4, 2014, which was a continuation-in-part of International Application No. PCT/US2013/020724 filed Jan. 9, 2013 which claims priority to United States Provisional Application No. 61/592,643, filed Jan. 31, 2012. U.S. patent application Ser. No. 14/248,386 further claims priority to U.S. Provisional Application No. 61/884,660 filed Sep. 30, 2013. 
     
    
     BACKGROUND 
       [0002]    This application relates to the design of a turbine which can be operated to produce noise to which human hearing is less sensitive. 
         [0003]    Gas turbine engines are known, and typically include a fan delivering air into a compressor. The air is compressed in the compressor and delivered downstream into a combustor section where it was mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving the turbine rotors to rotate. 
         [0004]    Typically, there is a high pressure turbine rotor, and a low pressure turbine rotor. Each of the turbine rotors includes a number of rows of turbine blades which rotate with the rotor. Typically interspersed between the rows of turbine blades are vanes. 
         [0005]    The low pressure turbine can be a significant noise source, as noise is produced by fluid dynamic interaction between the blade rows and the vane rows. These interactions produce tones at a blade passage frequency of each of the low pressure turbine stages, and their harmonics. 
         [0006]    The noise can often be in a frequency range to which humans are very sensitive. To mitigate this problem, in the past, a vane-to-blade ratio of the fan drive turbine has been controlled to be above a certain number. As an example, a vane-to-blade ratio may be selected to be 1.5 or greater, to prevent a fundamental blade passage tone from propagating to the far field. This is known as acoustic “cut-off.” 
         [0007]    However, acoustically cut-off designs may come at the expense of increased weight and reduced aerodynamic efficiency. Stated another way, if limited to a particular vane to blade ratio, the designer may be restricted from selecting such a ratio based upon other characteristics of the intended engine. 
         [0008]    Historically, the low pressure turbine has driven both a low pressure compressor section and a fan section. More recently, a gear reduction has been provided such that the fan and low pressure compressor can be driven at different speeds. 
       SUMMARY 
       [0009]    In a featured embodiment, a gas turbine engine comprises a fan and a turbine having a fan drive rotor. There is also a second turbine rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive rotor. The fan drive rotor has a number of turbine blades in at least one of a plurality of rows of the fan drive rotor, and the turbine blades operate at least some of the time at a rotational speed. The number of turbine blades in the at least one row and the rotational speed are such that the following formula holds true for the at least one row of the fan drive turbine: (number of blades×speed)/60≧5500 Hz. The rotational speed is in revolutions per minute. 
         [0010]    In another embodiment according to the previous embodiment, the formula results in a number greater than or equal to 6000 Hz. 
         [0011]    In another embodiment according to any of the previous embodiments, the gas turbine engine is rated to produce 15,000 pounds of thrust or more. 
         [0012]    In another embodiment according to any of the previous embodiments, the formula holds true for the majority of blade rows of the fan drive rotor. 
         [0013]    In another embodiment according to any of the previous embodiments, the rotational speed is an approach speed. 
         [0014]    In another embodiment according to any of the previous embodiments, the turbine section has a higher pressure turbine rotor and a lower pressure turbine rotor. The fan drive rotor is the lower pressure turbine rotor and the second turbine rotor is the higher pressure turbine rotor. 
         [0015]    In another embodiment according to any of the previous embodiments, there is a third turbine rotor, with the fan drive turbine being a most downstream of the three turbine rotors. 
         [0016]    In another featured embodiment, a method of designing a gas turbine engine comprises the steps of including a gear reduction between a fan drive turbine rotor and a fan. A number of blades in at least one row of the fan drive turbine rotor is selected, in combination with a rotational speed of the fan drive turbine rotor, such that the following formula holds true for the at least one row of the fan drive turbine rotor: (number of blades× 33  speed)/60≧5500 Hz. The rotational speed is in revolutions per minute. A second turbine rotor is included. 
         [0017]    In another embodiment according to the previous embodiment, the formula results in a number greater than or equal to 6000. 
         [0018]    In another embodiment according to any of the previous embodiments, the gas turbine engine is rated to produce 15,000 pounds of thrust or more. 
         [0019]    In another embodiment according to any of the previous embodiments, the formula holds true for the majority of the blade rows of the fan drive turbine. 
         [0020]    In another embodiment according to any of the previous embodiments, the rotational speed is an approach speed. 
         [0021]    In another embodiment according to any of the previous embodiments, a turbine section includes a higher pressure turbine rotor and a lower pressure turbine rotor. The fan drive turbine rotor is the lower pressure turbine rotor. 
         [0022]    In another featured embodiment, a turbine module comprises a fan drive rotor having a first blade row that includes a number of blades. The first blade row is capable of rotating at a rotational speed, so that when measuring the rotational speed in revolutions per minute: (number of blades x the rotational speed)/60≧5500 Hz. 
         [0023]    In another embodiment according to any of the previous embodiments, the formula results in a number greater than or equal to 6000. 
         [0024]    In another embodiment according to any of the previous embodiments, the formula holds true for the majority of blade rows of the fan drive rotor. 
         [0025]    In another embodiment according to any of the previous embodiments, the rotational speed is an approach speed. 
         [0026]    In another embodiment according to any of the previous embodiments, there being a higher pressure turbine rotor and a lower pressure turbine rotor, and the fan drive rotor is the lower pressure turbine rotor. 
         [0027]    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. 
         [0028]    These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0029]      FIG. 1  shows a gas turbine engine. 
           [0030]      FIG. 2  shows another embodiment. 
           [0031]      FIG. 3  shows yet another embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0032]      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), or an intermediate spool, among other systems or features. The fan section  22  drives air along a bypass flowpath B while the compressor section  24  drives air along a core flowpath 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. 
         [0033]    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. 
         [0034]    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 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 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 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. 
         [0035]    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. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
         [0036]    The terms “low” and “high” as applied to speed or pressure for the spools, compressors and turbines are of course relative to each other. That is, the low speed spool operates at a lower speed than the high speed spool, and the low pressure sections operate at lower pressure than the high pressures sections. 
         [0037]    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 star system, a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 or greater than about 2.5:1. 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. The low pressure turbine  46  pressure ratio is a ratio of the pressure measured at inlet of low pressure turbine  46  to the pressure at the outlet of the low pressure turbine  46  (prior to an exhaust nozzle). 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. 
         [0038]    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 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 1.50 and, in some embodiments, 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 [(Tambient 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. 
         [0039]    The use of the gear reduction between the low pressure turbine spool and the fan allows an increase of speed to the low pressure compressor. In the past, the speed of the low pressure turbine has been somewhat limited in that the fan speed cannot be unduly high. The maximum fan speed is at its outer tip, and in larger engines, the fan diameter is much larger than it may be in lower power engines. However, a gear reduction may be used to free the designer from compromising low pressure turbine speed in order not to have unduly high fan speeds. 
         [0040]    It has been discovered that a careful design between the number of rotating blades, and the rotational speed of the low pressure turbine can be selected to result in noise frequencies that are less sensitive to human hearing. 
         [0041]    A formula has been developed as follows: 
         [0000]      (blade count×rotational speed)/(60 seconds/minute)≧4000 Hz.
 
         [0042]    That is, the number of rotating blades in any low pressure turbine stage, multiplied by the rotational speed of the low pressure turbine (in revolutions per minute), divided by 60 seconds per minute (to put the amount per second, or Hertz) should be greater than or equal to 4000 Hz. In one embodiment, the amount is above 5500 Hz. And, in another embodiment, the amount is above about 6000 Hz. 
         [0043]    The operational speed of the low pressure turbine as utilized in the formula should correspond to the engine operating conditions at each noise certification point currently defined in Part 36 or the Federal Airworthiness Regulations. More particularly, the rotational speed may be taken as an approach certification point as currently defined in Part 36 of the Federal Airworthiness Regulations. For purposes of this application and its claims, the term “approach speed” equates to this certification point. 
         [0044]    Although the above formula only needs to apply to one row of blades in the low pressure turbine  26 , in one embodiment, all of the rows in the low pressure turbine meet the above formula. In another embodiment, the majority of the blade rows in the low pressure turbine meet the above formula. 
         [0045]    This will result in operational noise to which human hearing will be less sensitive. 
         [0046]    In embodiments, it may be that the formula can result in a range of greater than or equal to 4000 Hz, and moving higher. Thus, by carefully designing the number of blades and controlling the operational speed of the low pressure turbine (and a worker of ordinary skill in the art would recognize how to control this speed) one can assure that the noise frequencies produced by the low pressure turbine are of less concern to humans. 
         [0047]    This invention is most applicable to jet engines rated to produce 15,000 pounds of thrust or more and with bypass ratios greater than about 8.0. 
         [0048]      FIG. 2  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 . 
         [0049]      FIG. 3  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. 
         [0050]    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.