Gas turbine engine with separate core and propulsion unit

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.

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.

DETAILED DESCRIPTION

The engine20generally includes a low speed spool30and a high speed spool32mounted for rotation about an engine central longitudinal axis C relative to an engine static structure36.

The core airflow is compressed by the low pressure compressor44then the high pressure compressor52, mixed and burned with fuel in the combustor56, then expanded over the high pressure turbine54and low pressure turbine46. The turbines46,54rotationally drive the respective low speed spool30and high speed spool32in response to the expansion.

An aircraft wing352is shown with the gas turbine engine20mounted somewhat forwardly of the engine. A pylon351mounts the gas turbine engine to the wing352. 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 tank400, shown schematically, is mounted. Thus, this has somewhat limited the mounting of gas turbine engines in the past.

As can be appreciated fromFIG. 1, the turbines, compressors are all inter-related and rotate on a common axis with the extending spools30/32. As can be appreciated fromFIG. 1, it is somewhat difficult to remove the turbine, compressors, or fans separately from one another for service.

FIG. 2shows an inventive engine100. Air at114approaches a fan rotor111which is driven to rotate with a fan hub110. A shaft120is driven through a gear reduction118, which is in turn driven by a shaft125. Shaft125is driven by a free turbine127. A duct310communicates products of combustion from a core engine130that includes low pressure turbine170, a high pressure turbine160, a combustor155, and a compressor section including a high pressure compressor150and a low pressure compressor145. A spool165rotates the low pressure spool while a spool175rotates the high pressure spool.

An inlet duct195communicates air from a turning duct185into the low pressure compressor145. An opening190takes air from a main duct105. A pylon200mounts the engine100to an aircraft wing352.

A centerline X of core engine130incorporating the compressor sections145,150, combustor155and compressor sections160and170is offset by an angle B from a center line C of the shaft120/125. Thus, the fan rotor111rotates about axis C while the core engine130rotates 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 theFIG. 2embodiment, and for moving the burst zone A, the angle B should be greater than zero.

As further shown, a strut210extends from the pylon200and mounts to the duct310.

In the engine100, rather than delivering air into a core airflow at a fan side of the engine, all of the air is delivered into the duct105. A propulsion unit including the free turbine127, gear reduction118, and fan rotor111deliver this air beyond struts116, and to an outlet410of a cowl411. This provides the bulk of the propulsion for the engine. The inlet190into the turning duct185takes a portion of the air and delivers it into the inlet195for the compressor145. The air is compressed, delivered into the higher compressor section150, into the combustion section155, and across turbines160and170, which in turn drive the compressors150and145. Air downstream of the turbine section170passes through the duct310, and is driven across the free turbine127. The free turbine127drives gear reduction118to in turn cause the fan blades111to rotate.

Air downstream of the free turbine section127passes back outwardly and into the duct105through openings in struts116.

As can be appreciated fromFIG. 2, since the core engine130is 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 wing352. 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 wing352. That is, if the angle B has at least a component extending toward the wing352from 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 inFIG. 3, there are a plurality of struts116delivering air back into the duct105. Generally the struts which deliver air into the duct are not aligned with the opening190into the turning duct185.

An embodiment600is shown schematically inFIG. 4. As shown, a core engine608may communicate gas flow from an inlet duct606, through a compressor and turbine section as shown inFIG. 2. Products of the combustion downstream of the turbine sections in the core engine608pass into a connecting duct610, and then across a free turbine612. The free turbine612may drive the fan rotor602. The outlet gas from the free turbine612may be directed through the struts614and into a main duct604. As shown in this Figure, there is a separate propulsion unit including the free turbine612and fan rotor602. 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 engine600, 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 fan602is positioned at an inlet end of a main air duct604. The free turbine is between the inlet end and the core engine608relative 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.”