Abstract:
A gas turbine engine includes a propulsion unit mounted to rotate about a first axis, and a core engine mounted to rotate about a second axis, and wherein the first and second axes are non-parallel. A gas turbine engine includes a propulsion unit driven by a free turbine which is adjacent to the propulsion unit and an associated fan, and having a gas generator core engine including a compressor, combustor and turbine section. A method is also disclosed.

Description:
BACKGROUND OF THE INVENTION 
     This application relates to a gas turbine engine, wherein a core engine is mounted separately from a propulsion unit. 
     Gas turbine engines are known, and have typically included a fan delivering a portion of air into a bypass duct, and a second portion of air into a core flow leading into a compressor section. The air is compressed in the compressor and delivered downstream into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass across turbine rotors which are driven to rotate, and in turn rotate the compressor and fan section. Historically one turbine section drove both a compressor stage and a fan at the same speed. More recently it has been proposed to incorporate a gear reduction such as the fan can rotate at slower speeds than the compressor stage. With this arrangement, the outer diameter of the fan can increase, and the outer diameter of the turbine and compressor sections can decrease. 
     Historically, the fan and compressors have been mounted coaxially, and have been driven by turbines that are at a rear end of the engine, with the fan and compressor at a forward end. It has typically not been possible to service any portion of the engine, without removing the concentrically rotating turbines, compressors and fan as a combined unit. At a minimum, service is made complex by the inter-relationships of these sections. 
     Another challenge with mounting gas turbine engines relates to the so called “disk burst zone.” This zone is an area where broken pieces from a core engine could be driven. 
     The disk burst zone extends for approximately 30° about the last stage of the gas turbine engine. The gas turbine engine is typically mounted to an aircraft wing through a pylon. The aircraft wing also includes a fuel tank. There is a limitation on the mounting of current gas turbine engines in that the disk burst zone cannot extend through the fuel tank. Thus, gas turbine engines have typically been necessarily been mounted somewhat forwardly on the aircraft wing. 
     SUMMARY OF THE INVENTION 
     In a featured embodiment, a gas turbine engine has a propulsion unit including a fan, and a free turbine connected to drive the fan about a first axis. A core engine includes at least a compressor, a combustion section, and a turbine. The core engine turbine is connected to drive the compressor. The compressor and the core engine turbine rotate about a second axis. The first and second axes are non-parallel to each other. 
     In another embodiment according to the previous embodiment, a gear reduction is between the free turbine and fan. 
     In another embodiment according to the previous embodiment, an angle is defined between the first and second axes. The angle has a component that extends in a direction that will approach an aircraft wing that is to mount the gas turbine engine. 
     In another embodiment according to the previous embodiment, an angle is defined between the first and second axes. A range of the angle is greater than zero and less than or equal to about 90°. 
     In another embodiment according to the previous embodiment, the fan delivers air into a main duct. The main duct has an inlet tapping a portion of the air from the main duct into a turning duct which feeds air into the compressor. 
     In another embodiment according to the previous embodiment, the turning duct generally reverses a direction of flow of air from the main duct into the compressor. 
     In another embodiment according to the previous embodiment, an outlet of gas downstream of the free turbine extends back into the main duct. 
     In another embodiment according to the previous embodiment, the outlet extends into the main duct through struts extending across the main duct. 
     In another embodiment according to the previous embodiment, the struts are positioned upstream of the location where the turning duct taps air from the main duct. 
     In another embodiment according to the previous embodiment, the struts which have the outlet of gas downstream of the free turbine are circumferentially spaced from the inlet into the turning duct. 
     In another embodiment according to the previous embodiment, a connecting duct connects the core engine turbine to the free turbine. 
     In another embodiment according to the previous embodiment, the connecting duct is a mount location for mounting the core engine to an aircraft. 
     In another embodiment according to the previous embodiment, there are two turbine stages and two compressor stages in the core engine. 
     In another embodiment according to the previous embodiment, the propulsion unit is positioned such that its free turbine and fan are in a forward end of the gas turbine engine. The core engine is spaced rearwardly, and is separate from the propulsion unit. 
     In another featured embodiment, an aircraft has a wing and a pylon mounting a gas turbine engine to the wing. The gas turbine engine includes a propulsion unit including a fan, and a free turbine connected to drive the fan about a first axis, a core engine including at least the compressor, a combustion section, and a turbine. The core engine turbine is connected to drive the compressor. The compressor and core engine turbine rotate about a second axis. The first and second axis are non-parallel to each other. 
     In another embodiment according to the previous embodiment, an angle may be defined between the first and second axes. The angle has a component extending in a direction that will approach an aircraft wing which is to mount the gas turbine engine. 
     In another embodiment according to the previous embodiment, a connecting duct connects the core engine turbine to the free turbine. 
     In another embodiment according to the previous embodiment, the connecting duct is a mount location for mounting the core engine to the wing. 
     In another embodiment according to the previous embodiment, a strut extends from the pylon to be connected to the connecting duct. 
     In another embodiment according to the previous embodiment, there are two turbine stages and two compressor stages in the core engine. 
     In another featured embodiment, a gas turbine engine has a propulsion unit including a fan, and a free turbine connected to drive the fan. A core engine includes at least a compressor, a combustion section and a turbine. The core engine turbine is connected to drive the compressor. The compressor and core engine turbine are positioned toward an outlet end of the gas turbine engine relative to the propulsion unit. The core engine is separate from the propulsion unit. 
     In another embodiment according to the previous embodiment, the core engine compressor receives air from a main air duct. The fan delivers air into the main air duct. The air delivered into the core engine compressor is compressed, passed into the combustion section, and products of combustion pass over turbine rotors heading in a direction back toward the fan. The free turbine receives the products of combustion downstream of the core engine turbine. A connecting duct connects the core engine to the free turbine. 
     In another embodiment according to the previous embodiment, the core engine rotates on an axis which is co-linear with a rotation axis of the free turbine and fan. 
     In another embodiment according to the previous embodiment, the fan is positioned at an inlet end of a main air duct. The free turbine is positioned between the inlet end and core engine relative to an axial dimension extending along a rotational axis of the fan, and from the inlet end toward an outlet end of the main duct. 
     In another featured embodiment, a method of mounting a gas turbine engine to an aircraft wing includes providing a propulsion unit including a fan and a free turbine connected to drive the fan about a first axis, and connecting a core engine to the free turbine. The core engine includes at least a compressor, a combustion section, and a turbine. The core engine turbine is connected to drive the compressor. The compressor and core engine turbine rotate about a second axis. The first and second axes are non-parallel to each other. The second axis is selected to move a disk burst zone forwardly relative to an aircraft wing such that a gas turbine engine incorporating the propulsion unit and core engine can be mounted further rearwardly on the aircraft wing. 
     In another embodiment according to the previous embodiment, an angle between the first and second axes is selected to control the desired amount of movement of the disk burst zone. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a prior art gas turbine engine. 
         FIG. 2  is a cross-sectional view of an inventive gas turbine engine. 
         FIG. 3  is a partial view of a portion of the  FIG. 2  engine. 
         FIG. 4  is an alternative embodiment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a known 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 flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a 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 turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     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 C relative to an engine static structure  36 . 
     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 . The inner shaft  40  and the outer shaft  50  are concentric and rotate about the engine central longitudinal axis C 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 turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     An aircraft wing  352  is shown with the gas turbine engine  20  mounted somewhat forwardly of the engine. A pylon  351  mounts the gas turbine engine to the wing  352 . As shown, a disk burst zone A extends for about 30° across an exit point of the gas turbine engine. This is an area where portions of the gas turbine engine which may fracture, such as portions of the rotor disks, could fly outwardly and damage the wing, as an example. A limitation on the design of where to mount a gas turbine engine is that the disk burse zone A cannot extend across the area where a fuel tank  400 , shown schematically, is mounted. Thus, this has somewhat limited the mounting of gas turbine engines in the past. 
     As can be appreciated from  FIG. 1 , the turbines, compressors are all inter-related and rotate on a common axis with the extending spools  30 / 32 . As can be appreciated from  FIG. 1 , it is somewhat difficult to remove the turbine, compressors, or fans separately from one another for service. 
       FIG. 2  shows an inventive engine  100 . Air at  114  approaches a fan rotor  111  which is driven to rotate with a fan hub  110 . A shaft  120  is driven through a gear reduction  118 , which is in turn driven by a shaft  125 . Shaft  125  is driven by a free turbine  127 . A duct  310  communicates products of combustion from a core engine  130  that includes low pressure turbine  170 , a high pressure turbine  160 , a combustor  155 , and a compressor section including a high pressure compressor  150  and a low pressure compressor  145 . A spool  165  rotates the low pressure spool while a spool  175  rotates the high pressure spool. 
     An inlet duct  195  communicates air from a turning duct  185  into the low pressure compressor  145 . An opening  190  takes air from a main duct  105 . A pylon  200  mounts the engine  100  to an aircraft wing  352 . 
     A centerline X of core engine  130  incorporating the compressor sections  145 ,  150 , combustor  155  and compressor sections  160  and  170  is offset by an angle B from a center line C of the shaft  120 / 125 . Thus, the fan rotor  111  rotates about axis C while the core engine  130  rotates about an axis X, which is offset by an angle B. The angle B may be some non-zero angle, or as described below, may be zero in at least some embodiments. In embodiments which position the core engine to be offset, the angle B may be greater than zero and less than or equal to about 90°. Note other angles can be utilized. The burst zone features are maximized across this range. 
     For purposes of the  FIG. 2  embodiment, and for moving the burst zone A, the angle B should be greater than zero. 
     As further shown, a strut  210  extends from the pylon  200  and mounts to the duct  310 . 
     In the engine  100 , rather than delivering air into a core airflow at a fan side of the engine, all of the air is delivered into the duct  105 . A propulsion unit including the free turbine  127 , gear reduction  118 , and fan rotor  111  deliver this air beyond struts  116 , and to an outlet  410  of a cowl  411 . This provides the bulk of the propulsion for the engine. The inlet  190  into the turning duct  185  takes a portion of the air and delivers it into the inlet  195  for the compressor  145 . The air is compressed, delivered into the higher compressor section  150 , into the combustion section  155 , and across turbines  160  and  170 , which in turn drive the compressors  150  and  145 . Air downstream of the turbine section  170  passes through the duct  310 , and is driven across the free turbine  127 . The free turbine  127  drives gear reduction  118  to in turn cause the fan blades  111  to rotate. 
     Air downstream of the free turbine section  127  passes back outwardly and into the duct  105  through openings in struts  116 . 
     As can be appreciated from  FIG. 2 , since the core engine  130  is mounted at an axis which is non-parallel to the axis C, the disk burst zone A is shifted, or angled, forwardly away from the wing  352 . Now, the engine may be mounted further rearwardly underneath the wing than has been the case in the prior art. Essentially, a core engine, mounted at an axis which is non-parallel to the axis of a propulsion unit C would achieve this benefit whenever the axis X is mounted to extend toward the wing  352 . That is, if the angle B has at least a component extending toward the wing  352  from the propulsion unit drive axis C, then this forward movement of the disk burst zone A will be achieved. The amount of movement can be controlled by changing the size of the angle B. A method of selecting the angle B to position to disk burst zone A such that the engine can be mounted further rearwardly under the wing would also be apparent from the above disclosure. 
     As can be appreciated in  FIG. 3 , there are a plurality of struts  116  delivering air back into the duct  105 . Generally the struts which deliver air into the duct are not aligned with the opening  190  into the turning duct  185 . 
     An embodiment  600  is shown schematically in  FIG. 4 . As shown, a core engine  608  may communicate gas flow from an inlet duct  606 , through a compressor and turbine section as shown in  FIG. 2 . Products of the combustion downstream of the turbine sections in the core engine  608  pass into a connecting duct  610 , and then across a free turbine  612 . The free turbine  612  may drive the fan rotor  602 . The outlet gas from the free turbine  612  may be directed through the struts  614  and into a main duct  604 . As shown in this Figure, there is a separate propulsion unit including the free turbine  612  and fan rotor  602 . This may also include a gear reduction in some embodiments. The separate propulsion unit is positioned forward or toward the inlet of the gas turbine engine  600 , while the core engine is spaced rearwardly of the propulsion unit, and is separate from the propulsion unit. With this embodiment, servicing of the core engine relative to the propulsion unit is simplified compared to the prior art. 
     The fan  602  is positioned at an inlet end of a main air duct  604 . The free turbine is between the inlet end and the core engine  608  relative to an axial dimension extending along a rotational axis of the fan, and from the inlet end toward an outlet end of the main duct. 
     Further modifications which can flow given the separate propulsion unit and core engines, and in particular, the ability to provide modular engines, are disclosed in co-pending U.S. patent application Ser. No. 13/370,743, filed on even date herewith and entitled “Gas Turbine Engine With Modular Cores and Propulsion Unit.” 
     Although an embodiment of this invention has been disclosed, a worker of ordinary skill in the 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.