Abstract:
A gas turbine engine has a turbine section, and a housing enclosing the turbine section, with a mount structure secured to the housing for mounting and including internal flow passages for delivering air to remote locations.

Description:
BACKGROUND 
     This application relates to a mount for a gas turbine engine. 
     Conventional gas turbine engines typically include a fan section, a compressor section and a turbine section. There are two generally 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 engines where the fan is directly driven. 
     Another known architecture includes a third spool with a third turbine 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 gas turbine engine has a turbine section, and a housing enclosing the turbine section, with a mount structure secured to the housing for mounting and including internal flow passages for delivering air to remote locations. 
     In another embodiment according to the previous embodiment, a turbine case is positioned downstream of the turbine section, and the mount structure is secured to an outer surface of the turbine case. 
     In another embodiment according to any of the previous embodiments, the mount structure is provided with a mount plate. 
     In another embodiment according to any of the previous embodiments, the mount plate is connected to the mount ring by a plurality of ribs. 
     In another embodiment according to any of the previous embodiments, the ribs are positioned on opposed circumferential sides of the mount plate. 
     In another embodiment according to any of the previous embodiments, air supply passages are also formed within the mount plate and at least one of the ribs. 
     In another embodiment according to any of the previous embodiments, air supply passages are also formed within the mount plate. 
     In another embodiment according to any of the previous embodiments, the air from the mount structure is delivered to a gear reduction for driving a propulsor. 
     In another embodiment according to any of the previous embodiments, the gear reduction is driven by a propulsor turbine, with the propulsor turbine downstream of the turbine section. 
     In another embodiment according to any of the previous embodiments, the mount structure is mounted intermediate the turbine section and the propulsor. 
     In another embodiment according to any of the previous embodiments, the air from the mount structure is delivered to a pitch control mechanism for a propeller. 
     In another embodiment according to any of the previous embodiments, the mount structure is mounted intermediate the turbine section and the propeller. 
     In another embodiment according to any of the previous embodiments, air is supplied into a port in the mount structure. 
     In another embodiment according to any of the previous embodiments, air is supplied into the port in the mount structure from a forward location on the gas turbine engine. 
     In another embodiment according to any of the previous embodiments, air from the mount structure is delivered to an environmental control system for an aircraft. 
     In another embodiment according to any of the previous embodiments, air is supplied into a port in the mount structure from a forward location on the gas turbine engine. 
     In another embodiment according to any of the previous embodiments, air from the mount structure is delivered to an environmental control system for an aircraft. 
     In another embodiment according to any of the previous embodiments, the mount structure is mounted intermediate the turbine section and a propulsor. 
     In another embodiment according to any of the previous embodiments, a first compressor rotor is upstream of a second compressor rotor with a ratio of a pressure ratio across the first compressor rotor to a pressure ratio across the second compressor rotor being greater than or equal to about 2.0. 
     In another embodiment according to any of the previous embodiments, the ratio is less than or equal to about 8. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a three spool gas turbine engine. 
         FIG. 2A  schematically shows an engine. 
         FIG. 2B  shows an engine and mount. 
         FIG. 3  shows details of the engine mount. 
         FIG. 4  shows further details. 
     
    
    
     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  that 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: 
     more work is performed by the low pressure compressor rotor  30  than by the high pressure compressor rotor  26 ; 
     the combination of a first pressure ratio through the low pressure compressor rotor  30  and a second pressure ratio through the 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; in the illustrated embodiment, 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; and 
     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  rotates 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  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  99  connecting the propulsor turbine to the propeller  36 . 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 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. 
     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. 
     With the relatively 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 . While element  60  is described as a “ring,” it should be understood that other shapes would come within the teachings of this application. 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  60  and includes a plurality of ribs, one of which,  100 , is illustrated in  FIG. 2A . An aircraft body  184  is shown schematically and is secured to the plate  64 . As shown, the ring  60  has air passages  201  leading to the gear reduction  200 . Cooling air may be supplied with the ring  60  being utilized as part of the cooling air supply. Further, a separate passage  202  extends from the ring  60  to a pitch control mechanism  203 . Again, the air may be utilized for cooling the pitch control mechanism  203 . As known, the pitch control mechanism  203  may allow changing the pitch of the propeller blades  36 . 
     As shown in  FIG. 2B , there are pairs of ribs  100  and  101  extending in opposed lateral directions and fixed between the plate  64  and the ring  60 . 
     As shown in  FIG. 3 , the ring  60  includes pairs of ribs  100  and  101  connecting the ring  60  to the plate  64 . The ring  60  includes ports  104  which may receive cooling air, and the air may be circulated within hollow passages within the ring  60 , the ribs  100  and  101 , and the plate  64 . The air is circulated to be adjacent to passages, such as the passages  201  or  202  as shown in  FIG. 2A . Thus, the ring  60  and plate  64  not only provide mounting structure, they also eliminate the need for additional plumbing to route air across the system. 
       FIG. 4  shows an embodiment wherein an air supply  302  receives air from a forward end of the engine, passes that air through a passage  304  and to a passage  300  leading into the port  104 , such that the air may circulate through the ring  60 . In this way, air is supplied into the ring  60 , and then may be distributed as mentioned above to various components such as the gear reduction  200 , or the pitch control mechanism  203 . 
     A cabin air supply system is shown schematically at  FIG. 4 , and may be for use on an aircraft receiving the engine. Again, this air may pass through the ring  60  on its way to the cabin air supply system  308   
     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.