Abstract:
A gas turbine engine has a first shaft including a first turbine rotor, and a second shaft including a second turbine rotor disposed downstream of the first turbine rotor. A third shaft includes a propulsor turbine positioned downstream of the second turbine rotor for driving a propeller. A mount ring is secured between the second turbine rotor and the propeller.

Description:
BACKGROUND 
       [0001]    This application relates to a two spool gas generator for a gas turbine engine and a propulsor drive. 
         [0002]    Conventional gas turbine engines typically include a fan section, a compressor section and a turbine section. There are two general known architectures. In one architecture, a low speed spool includes a low pressure turbine driving a low pressure compressor and also driving a fan. A gear reduction may be placed between the spool and the fan in some applications. There are also direct drive engines. 
         [0003]    Another known architecture includes a third spool with a third turbine being positioned downstream of the low pressure turbine and driving the fan. The three spools have shafts connecting a turbine to the driven element, and the three shafts are mounted about each other. 
         [0004]    All of these architectures raise challenges. 
       SUMMARY 
       [0005]    In a featured embodiment, a gas turbine engine has a first shaft including a first turbine rotor, and a second shaft including a second turbine rotor disposed downstream of the first turbine rotor. A third shaft includes a propulsor turbine positioned downstream of the second turbine rotor for driving a propeller. A mount ring is secured between the second turbine rotor and the propeller. 
         [0006]    In another embodiment according to the previous embodiment, a turbine case is positioned intermediate the second turbine rotor and the propeller. The mount ring is secured to an outer surface of the turbine case. 
         [0007]    In another embodiment according to any of the previous embodiments, the propulsor turbine is mounted within the turbine case. 
         [0008]    In another embodiment according to any of the previous embodiments, the mount ring is provided with a mount plate. 
         [0009]    In another embodiment according to any of the previous embodiments, the mount plate is connected to the mount ring by a plurality of pivotally connected links. 
         [0010]    In another embodiment according to any of the previous embodiments, two of the links are positioned on opposed circumferential sides of the mount plate. 
         [0011]    In another embodiment according to any of the previous embodiments, the mount plate is pivotally attached to the mount ring. 
         [0012]    In another embodiment according to any of the previous embodiments, a torque link is pivotally connected to the mount plate and to the mount ring. 
         [0013]    In another embodiment according to any of the previous embodiments, the first turbine rotor drives a first compressor rotor through the first shaft, and the second turbine rotor drives a second compressor rotor through the second shaft. 
         [0014]    In another embodiment according to any of the previous embodiments, the second compressor rotor has a first overall pressure ratio. The first compressor rotor has a second overall pressure ratio, with the ratio of the first overall pressure ratio to the second overall pressure ratio being greater than or equal to about 2.0. 
         [0015]    In another embodiment according to any of the previous embodiments, the ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 3.0. 
         [0016]    In another embodiment according to any of the previous embodiments, the ratio of the first overall pressure ratio to the second overall pressure ratio is less than or equal to about 8.0. 
         [0017]    In another embodiment according to any of the previous embodiments, the first turbine rotor includes a single turbine stage. 
         [0018]    In another embodiment according to any of the previous embodiments, the second turbine rotor includes two stages. 
         [0019]    In another embodiment according to any of the previous embodiments, the second compressor rotor includes eight stages. 
         [0020]    In another embodiment according to any of the previous embodiments, the first compressor rotor includes six stages. 
         [0021]    In another embodiment according to any of the previous embodiments, the first compressor rotor includes six stages. 
         [0022]    In another embodiment according to any of the previous embodiments, wherein said second compressor rotor includes eight stages. 
         [0023]    These and other features may be best understood from the following drawings and specification. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0024]      FIG. 1  schematically shows a three spool gas turbine engine. 
           [0025]      FIG. 2A  schematically shows an engine. 
           [0026]      FIG. 2B  shows an engine mount structure. 
           [0027]      FIG. 3  shows details of the engine mount. 
       
    
    
     DETAILED DESCRIPTION 
       [0028]    A gas turbine engine  19  is schematically illustrated in  FIG. 1 . A core engine, or gas generator  20 , includes high speed shaft  21  is part of a high speed spool along with a high pressure turbine rotor  28  and a high pressure compressor rotor  26 . A combustion section  24  is positioned intermediate the high pressure compressor rotor  26  and the high pressure turbine rotor  28 . A shaft  22  of a low pressure spool connects a low pressure compressor rotor  30  to a low pressure turbine rotor  32 . 
         [0029]    Engine  19  also includes a free turbine  34  is shown positioned downstream of the low pressure turbine rotor  32  and serves to drive a propeller  36 . 
         [0030]    Various embodiments are within the scope of the disclosed engine. These include embodiments in which: 
         [0031]    a good deal more work is done by the low pressure compressor rotor  30  than by the high pressure compressor rotor  26 ; 
         [0032]    the combination of the low pressure compressor rotor  30  and high pressure compressor rotor  26  provides an overall pressure ratio equal to or above about 30; 
         [0033]    the low pressure compressor rotor  30  includes eight stages and has a pressure ratio at cruise conditions of 14.5; in this embodiment, the high pressure compressor rotor  26  had six stages and an overall pressure ratio of 3.6 at cruise; 
         [0034]    a ratio of the low pressure compressor pressure ratio to the high pressure compressor ratio is greater than or equal to about 2.0, and less than or equal to about 8.0; 
         [0035]    more narrowly, the ratio of the two pressure ratios is between or equal to about 3.0 and less than or equal to about 8; and 
         [0036]    even more narrowly, the ratio of the two pressure ratios is greater than about 3.5. 
         [0037]    In the above embodiments, the high pressure compressor rotor  26  will rotate at slower speeds than in the prior art. If the pressure ratio through the fan and low pressure compressor are not modified, this could result in a somewhat reduced overall pressure ratio. The mechanical requirements for the high pressure spool, in any event, are relaxed. 
         [0038]    With the lower compressor, the high pressure turbine rotor  28  may include a single stage. In addition, the low pressure turbine rotor  32  may include two stages. 
         [0039]    By moving more of the work to the low pressure compressor rotor  30 , there is less work being done at the high pressure compressor rotor  26 . In addition, the temperature at the exit of the high pressure compressor rotor  26  may be higher than is the case in the prior art, without undue challenges in maintaining the operation. 
         [0040]    Variable vanes are less necessary for the high pressure compressor rotor  26  since it is doing less work. Moreover, the overall core size of the combined compressor rotors  30  and  26  is reduced compared to the prior art. 
         [0041]    The engine  19  has what may be called a propulsor turbine  34  which is axially downstream of the low pressure turbine rotor  32 . Further, the high pressure spool radially surrounds the low pressure spool, but neither of the spools surrounds the propulsor turbine, nor the shaft  100  connecting the propulsor turbine to the propeller  36 . In this sense, the propulsor rotor is separate from the gas generator portion of the engine. 
         [0042]    The disclosed engine architecture creates a smaller core engine and yields higher overall pressure ratios and, therefore, better fuel consumption. Further, uncoupling the low pressure turbine  32  from driving prop  36  enables it to run at a lower compressor surge margin, which also increases efficiency. Moreover, shaft diameters can be decreased and, in particular, for the diameter of the low pressure shafts as it is no longer necessary to drive the prop  36  through that shaft. 
         [0043]    In the prior art, the ratio of the low pressure compressor pressure ratio to the high pressure compressor ratio was generally closer to 0.1 to 0.5. Known three spool engines have a ratio of the low pressure compressor pressure ratio to the high pressure compressor ratio of between 0.9 and 3.0. 
         [0044]    With the very small diameter core engine  20 , there will be challenges in mounting the engine  19  to an aircraft. In particular, if the engine  19  was mounted as in the prior art, at front and rear locations, there would be challenges from so-called “backbone bending” due to the small diameter. Thus, as shown in  FIG. 2A , a mount ring  60  is secured to a turbine case  70  that is downstream of the core engine  20 . The turbine case  70  may also receive the propulsor turbine  34  and the gear reduction  200 . The propellers  36  are downstream and beyond the turbine case. The ring  60  supplies the sole mount plane for the engine  19 . A plate  64  extends forwardly from the ring and includes a plurality of struts, one of which,  62 , is illustrated in  FIG. 2A . An aircraft body  84  is shown schematically and is secured to the plate  64 . 
         [0045]    As shown in  FIG. 2B , there is a pair of struts  62  extending in opposed lateral directions and pivotally connect between the plate  64  and the ring  60 . 
         [0046]    As shown in  FIG. 3 , the plate  64  is secured to aircraft body at  84 . The ring  60  has an inner surface  71  that will surround the turbine case  70  and be secured to the turbine case. Pivot point  74  and  75  also secure a torque link between the plate  64  and the ring  60 . Both struts  62  are shown pivotally attached at  63  to the plate  64  and pivotally attached at  65  to the ring  60 . Further, the plate  64  is itself pivotally attached at  80  to the ring. The ring  60  and plate  64  provide a cantilever mount for the engine  19 . 
         [0047]    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 disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.