Abstract:
A method of configuring a plurality of gas turbine engines includes the steps of configuring each of the engines with respective ones of a plurality of propulsors. Each propulsor includes a propulsor turbine and one of a fan and a propeller. Each of the engines is configured with respective ones of a plurality of substantially mutually alike gas generators, with the respective propulsor turbine driven by products of combustion downstream of the gas generator.

Description:
BACKGROUND 
     This application relates to a two spool gas generator for creating a family of gas turbine engines having different propulsor drives. 
     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. 
     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. 
     All of these architectures raise challenges. 
     SUMMARY 
     In a featured embodiment, a method of configuring a plurality of gas turbine engines includes the steps of configuring each of the engines with respective ones of a plurality of propulsors. Each propulsor includes a propulsor turbine, and one of a fan and a propeller. Each of the engines is configured with respective ones of a plurality of substantially mutually alike gas generators, with the respective propulsor turbine driven by products of combustion downstream of the gas generator. 
     In another embodiment according to the previous embodiment, the gas generators each have a compressor section with a first and a second compressor rotor, and a turbine section with a first and second turbine rotor. The propulsor turbine is downstream of the second turbine rotor. 
     In another embodiment according to any of the previous embodiments, the second compressor rotor has a first overall pressure ratio, and the first compressor rotor has a second overall pressure ratio. A ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0. 
     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. 
     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. 
     In another embodiment according to any of the previous embodiments, the first turbine rotor includes a single turbine stage. 
     In another embodiment according to any of the previous embodiments, the second turbine rotor includes two stages. 
     In another embodiment according to any of the previous embodiments, the second compressor rotor includes eight stages. 
     In another embodiment according to any of the previous embodiments, the first compressor rotor includes six stages. 
     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. 
     In another featured embodiment, a family of gas turbine engines has substantially mutually alike gas generators. A plurality of propulsor turbines are each driven by products of combustion downstream of one of the gas generators, with at least one of the plurality of propulsor turbines driving a fan and another of the plurality of propulsor turbines driving a propeller. 
     In another embodiment according to the previous embodiment, the gas generators each have a compressor section with a first and a second compressor rotor, and a turbine section with a first and second turbine rotor. The propulsor turbine is downstream of the second turbine rotor. 
     In another embodiment according to any of the previous embodiments, the second compressor rotor has a first overall pressure ratio, and the first compressor rotor has a second overall pressure ratio. A ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0. 
     In another embodiment according to any of the previous embodiments, the high pressure turbine includes a single turbine stage. 
     In another embodiment according to any of the previous embodiments, the low pressure compressor rotor includes eight stages. 
     In another embodiment according to any of the previous embodiments, the first compressor rotor includes six stages. 
     In another embodiment according to any of the previous embodiments, the propulsor turbine drives a fan located at an upstream end to supply a free airstream to the second compressor rotor. 
     In another embodiment according to any of the previous embodiments, the fan rotates about a first axis, and the first and second compressor rotors. The first and second turbine rotors rotate about a second axis. The first and second axes are non-parallel. 
     In another embodiment according to any of the previous embodiments, the propulsor turbine drive a plurality of propellers. 
     In another featured embodiment, a gas turbine engine has a first shaft connecting a first compressor rotor to be driven by a first turbine rotor. A second shaft connects a second compressor rotor to be driven by a second turbine rotor. The second compressor rotor is upstream of the first compressor rotor, and the first turbine rotor is upstream of the second turbine rotor. The second compressor rotor has a first overall pressure ratio, and the first compressor rotor has a second overall pressure ratio. A ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0. A propulsor turbine operatively connects to drive a propeller through a third shaft, with the propulsor turbine is positioned to be downstream of the first turbine rotor. A mid-turbine frame includes a bearing supporting a downstream end of the first shaft. The mid-turbine frame is positioned intermediate the second turbine rotor, and the propulsor turbine. An intermediate case includes a bearing supporting each of the first and second shafts. An inlet case is positioned upstream of the second compressor rotor, and includes a bearing that supports the first shaft, and a turbine exhaust case that receives the propulsor turbine. The inlet case further includes bearings supporting the third shaft. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a three spool gas turbine engine. 
         FIG. 2A  shows a second embodiment. 
         FIG. 2B  shows a possible detail of the second embodiment. 
         FIG. 3  shows further details of the first embodiment. 
         FIG. 4  shows further details of the first embodiment. 
     
    
    
     DETAILED DESCRIPTION 
     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 . 
     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 . 
     Various embodiments are within the scope of the disclosed engine. These include embodiments in which: 
     a good deal more work is done by the low pressure compressor rotor  30  than by the high pressure compressor rotor  26 ; 
     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; 
     the low pressure compressor rotor  30  includes eight stages and has a pressure ratio at cruise conditions of 14.5; 
     the high pressure compressor rotor  26  had six stages and an overall pressure ratio of 3.6 at cruise; 
     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; 
     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; 
     even more narrowly, the ratio of the two pressure ratios is greater than about 3.5. 
     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. 
     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. 
     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. 
     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. 
     The engine  60  as shown in  FIG. 2  includes a two spool core engine  120  including a low pressure compressor rotor  30 , a low pressure turbine rotor  32 , a high pressure compressor rotor  26 , and a high pressure turbine rotor  28 , and a combustor  24  as in the prior embodiments. This core engine  120  is a so called “reverse flow” engine meaning that the compressor  30 / 26  is spaced further into the engine than is the turbine  28 / 32 . Air downstream of the fan rotor  62  passes into a bypass duct  64 , and toward an exit  65 . However, a core inlet duct  66  catches a portion of this air and turns it to the low pressure compressor  30 . The air is compressed in the compressor rotors  30  and  26 , combusted in combustor  24 , and products of this combustion pass downstream over the turbine rotors  28  and  32 . The products of combustion downstream of the turbine rotor  32  pass over a fan drive turbine  74 . Then, the products of combustion exit through an exit duct  76  back into the bypass duct  64  (downstream of inlet  66  such that hot gas is not re-ingested into the core inlet  66 ), and toward the exit  65 . A gear reduction  63  may be placed between the fan drive turbine  74  and fan  62 . 
     The core engine  120 , as utilized in the engine  60 , may have characteristics similar to those described above with regard to the core engine  20 . 
     The engines  19  and  60  are similar in that they have what may be called a propulsor turbine ( 34  or  74 ) 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 surround the propulsor turbine, nor the shaft  100  connecting the propulsor turbine to the propellers  36  or fan  62 . In this sense, the propulsor rotor is separate from the gas generator portion of the engine. 
     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 a fan  62  or 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 fan  62  or prop  36  through that shaft. 
     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. 
     A disclosed method, and a family of gas turbine engines, utilize the common gas generator or two spool core including the low pressure turbine  32 , high pressure turbine  28 , combustor  24 , high pressure compressor  26 , and low pressure compressor  30 . Once these components have been designed, they can be utilized to create any number of gas turbine engines having a distinct free or propulsor turbine driving a propulsor that may be a fan or a propeller. The present invention, thus, allows a dramatic reduction in the design, development, test and manufacturing cost for creating a family of gas turbine engines having different propulsor arrangements. 
     As shown in  FIG. 2B , the reverse core gas generator may rotate about an axis Y, while the fan  62  may rotate about an axis X which is non-parallel to axis Y. This allows the overall length of the engine  60  to be reduced. As shown schematically in  FIG. 2B , an aircraft wing  200  may mount the engine  60 . 
       FIG. 3  shows further features of the gas generator  20  which includes the propulsor turbine  34  driving a gear reduction  310  that in turn drives the shaft  100  to drive propellers  36 . 
       FIG. 4  shows further details of the engine  19 . An inlet case  220  may include a bearing  221  supporting the shaft  22  at a forward end. An intermediate case  222  may include a bearing  223  supporting the shaft  22 , and another bearing  225  supporting the shaft  21 . A free intershaft bearing  229  may support both shafts  21  and  22 . A mid-turbine frame  230  may be positioned downstream of the lower pressure turbine  32 , and include a bearing  231  providing an end mount for the shaft  22 . 
     A turbine exhaust case  300  may mount the propulsor turbine  34 . The turbine exhaust case may include a plurality of bearings  301  and  302  supporting the shaft  100 . 
     Because the propulsor turbine  34 , and propeller  36  are configured as one unit, they can stay mounted to the aircraft while the gas generator  20  is removed. Due to the pressure ratio split of the gas generator  20 , the high spool is very small and lightweight, enabling the use of the intershaft bearing  229  between the high and low spool at an aft end of the gas generator  20 . Because an inner shaft bearing  231  is utilized, the mid-turbine frame  230  may be moved aft of the low pressure turbine  32 , into a cooler environment, which in turn improves cost and life. The front of the low pressure compressor  30  includes bearing  221 , supported by the inlet case  220 , so that the low pressure compressor  30  is straddle mounted. Straddle mounting of the low pressure compressor  30  improves control over blade shift clearances and further improves engine efficiency. 
     Once the two spool core engine or gas generator  20 / 120  has been designed, it can be utilized generally identically to create a family of gas turbine engines having distinct free or propulsor turbines driving distinct propulsors. Although two embodiments of the family of gas turbine engines can be provided are disclosed, a worker of ordinary skill in the art would recognize any number of other arrangements that could be provided given the power of this method. 
     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.