Patent Description:
Prior art gas turbine engines comprising a gear reduction are disclosed in <CIT>, <CIT>, <CIT>, <CIT>, <CIT> and <CIT>.

In one aspect of the present invention, a gas turbine engine is provided according to claim <NUM>.

In an embodiment, the first low spool support bearing is supported by an engine static structure located axially forward of the low pressure compressor.

In another embodiment according to any of the previous embodiments, the engine static structure is located axially forward of the low pressure compressor is a front center body.

In another embodiment according to any of the previous embodiments, the second low spool support bearing is supported by an engine static structure located axially aft of the low pressure compressor.

In another embodiment according to any of the previous embodiments, the engine static structure is located axially aft of the low pressure compressor is an intermediate case.

In another embodiment according to any of the previous embodiments, the intermediate case at least partially defines a portion of a core flow path through the gas turbine engine fluidly downstream of the low pressure compressor and fluidly upstream of the high pressure compressor.

In another embodiment according to any of the previous embodiments, the intermediate case includes at least one structural support strut spanning the core flow path.

In another embodiment according to any of the previous embodiments, an inner race of the first low spool support bearing is configured to rotate with the low spool. An outer race of the first low spool support bearing is fixed to the front center body. An inner race of the second low spool support bearing is configured to rotate with the low spool. An outer race of the second low spool support bearing is fixed to the intermediate case.

In another embodiment according to any of the previous embodiments, a mid-turbine frame is located axially between the high pressure turbine and the low pressure turbine and supports an axially aft end of the high spool.

In another embodiment according to any of the previous embodiments, a pair of low spool support bearings is located axially aft of the low pressure turbine. The low spool is unsupported by the mid-turbine frame.

In another embodiment according to any of the previous embodiments, an aft end of the high spool is supported by a bearing system engaging a diffuser case.

In another embodiment according to any of the previous embodiments, the fan section includes a fan drive shaft in driving engagement with the fan. A pair of fan shaft support bearing supports the fan drive shaft relative to the front center body. The carrier is in driving engagement with the fan drive shaft.

In another embodiment according to any of the previous embodiments, the low pressure compressor includes at least <NUM> stages and no more than <NUM> stages. The high pressure compressor includes more stages than the low pressure compressor.

The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description, The drawings that accompany the detailed description can be briefly described as follows.

The fan section <NUM> may include a single-stage fan <NUM> having a plurality of fan blades <NUM>. The fan blades <NUM> may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan <NUM> drives air along a bypass flow path B in a bypass duct <NUM> defined within a housing <NUM> such as a fan case or nacelle, and also drives air along a core flow path C for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. A splitter <NUM> aft of the fan <NUM> divides the air between the bypass flow path B and the core flow path C. The housing <NUM> may surround the fan <NUM> to establish an outer diameter of the bypass duct <NUM>. The splitter <NUM> may establish an inner diameter of the bypass duct <NUM>. The engine <NUM> may incorporate a variable area nozzle for varying an exit area of the bypass flow path B and/or a thrust reverser for generating reverse thrust.

The low speed spool <NUM> generally includes an inner shaft <NUM> that interconnects, a low pressure compressor <NUM> and a low pressure turbine <NUM>. The inner shaft <NUM> is connected to the fan <NUM> through a speed change mechanism, which in accordance with the invention is a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The inner shaft <NUM> may interconnect the low pressure compressor <NUM> and low pressure turbine <NUM> such that the low pressure compressor <NUM> and low pressure turbine <NUM> are rotatable at a common speed and in a common direction. In other embodiments, the low pressure turbine <NUM> drives both the fan <NUM> and low pressure compressor <NUM> through the geared architecture <NUM> such that the fan <NUM> and low pressure compressor <NUM> are rotatable at a common speed. Alternatively, the low pressure compressor <NUM> includes a forward hub 45A and an aft hub 45B driven by the inner shaft <NUM>.

The high speed spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure compressor <NUM> and a high pressure turbine <NUM>. In the illustrated example, the mid-turbine frame <NUM> only includes a bearing system <NUM> that supports the high spool <NUM> and the mid-turbine frame <NUM> does not support the low speed spool <NUM>. Additionally, a pair of bearing systems 38E are located adjacent a downstream end of the low speed spool <NUM> adjacent an exhaust outlet of the gas turbine engine to support the low speed spool <NUM>. Furthermore, a bearing assembly 38C can be located radially inward from the combustor <NUM> and supported by a diffuser case and be used in place of or in addition to the bearing system <NUM> associated with the mid-turbine frame <NUM>.

The low pressure compressor <NUM>, high pressure compressor <NUM>, high pressure turbine <NUM> and low pressure turbine <NUM> each include one or more stages having a row of rotatable airfoils. Each stage may include a row of vanes adjacent the rotatable airfoils. The rotatable airfoils are schematically indicated at <NUM>, and the vanes are schematically indicated at <NUM>. In one example, the low pressure compressor <NUM> includes at least <NUM> stages and no more than <NUM> stages and in another example, the low pressure compressor <NUM> includes at least <NUM> stages and no more than <NUM> stages. In both examples, the high pressure compressor <NUM> includes more stages than the low pressure compressor.

The engine <NUM> may be a high-bypass geared aircraft engine. The bypass ratio can be greater than or equal to <NUM> and less than or equal to about <NUM>, or more narrowly can be less than or equal to <NUM>. The geared architecture <NUM> may be an epicyclic gear train, such as a planetary gear system or a star gear system. The epicyclic gear train may include a sun gear, a ring gear, a plurality of intermediate gears meshing with the sun gear and ring gear, and a carrier that supports the intermediate gears. With the planetary gear system, the ring gear is fixed from rotation relative to the engine static structure <NUM> and the carrier rotates with the fan <NUM>. With the star gear system, the carrier is fixed from rotation relative to the engine static structure <NUM> and the ring gear rotates with the fan <NUM>. The sun gear may provide an input to the gear train. The ring gear (e.g., star gear system) or carrier (e.g., planetary gear system) may provide an output of the gear train to drive the fan <NUM>. A gear reduction ratio may be greater than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, and in some embodiments the gear reduction ratio is greater than or equal to <NUM>. The gear reduction ratio may be less than or equal to <NUM> or <NUM>. The fan diameter is significantly larger than that of the low pressure compressor <NUM>. The low pressure turbine <NUM> can have a pressure ratio that is greater than or equal to <NUM> and in some embodiments is greater than or equal to <NUM>. All of these parameters are measured at the cruise condition described below.

The fan section <NUM> of the engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>,<NUM> meters), The flight condition of <NUM> Mach and <NUM>,<NUM> ft (<NUM>,<NUM> meters), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The engine parameters described above, and those in the next paragraph are measured at this condition unless otherwise specified.

"Fan pressure ratio" is the pressure ratio across the fan blade <NUM> alone, without a Fan Exit Guide Vane ("FEGV") system. A distance is established in a radial direction between the inner and outer diameters of the bypass duct <NUM> at an axial position corresponding to a leading edge of the splitter <NUM> relative to the engine central longitudinal axis A. The fan pressure ratio is a spanwise average of the pressure ratios measured across the fan blade <NUM> alone over radial positions corresponding to the distance. The fan pressure ratio can be less than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, such as between <NUM> and <NUM>. "Corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]<NUM> (where °R = <NUM>/<NUM> x K). The corrected fan tip speed can be less than or equal to <NUM> ft / second (<NUM> meters/second), and can be greater than or equal to <NUM> ft / second (<NUM> meters/second).

As shown in <FIG>, the low speed spool <NUM> is supported by a number of bearing systems <NUM>. In particular, an axially forward end of the low speed spool <NUM> adjacent the geared architecture <NUM> is supported by a first low spool support bearing 38A and a second low spool support bearing 38B. The first bearing 38A is located axially between the low pressure compressor <NUM> and the geared architecture <NUM>. A location of attachment of the first bearing 38A with the low speed spool <NUM> is located axially upstream of a location of attachment of the low pressure compressor <NUM> to the low speed spool <NUM> and the location of engagement of the first bearing 38A is axially downstream of the geared architecture <NUM>. Furthermore, a location of attachment of the second bearing 38B with the low speed spool <NUM> is located axially downstream of the location of attachment of the low pressure compressor <NUM> with low speed spool <NUM> and axially upstream of the high pressure compressor <NUM> and the high speed spool <NUM>.

In this disclosure, axial and radial directions are in relation to the engine axis A unless stated otherwise. Additionally, axially upstream and downstream directions are in relation to a direction of flow of air through the core flow path C unless stated otherwise.

In the illustrated example, the first low spool support bearing 38A is supported by the engine static structure <NUM> located axially forward of the low pressure compressor <NUM>. In the illustrated example, the first bearing 38A is supported by a front center body 36A of the engine static structure <NUM>. An inner race of the first bearing 38A is configured to rotate with the low speed spool <NUM> and an outer race of the first bearing 38A is fixed relative to the front center body 36A. The front center body 36A provides structure support to a front of the gas turbine engine <NUM> forward of the low pressure compressor <NUM>. The front center body 36A can include structural vanes and/or struts <NUM> that pass through the core flow path C upstream of the low pressure compressor <NUM>.

In addition to supporting the first bearing 38A, the front center body 36A provides structural support for a pair of fan shaft support bearings 38F that support a fan drive shaft <NUM>. The fan bearings 38F each include an inner race that is configured to rotate with the fan drive shaft <NUM> and an outer race fixed relative to the front center body 36A of the static structure <NUM>. The fan drive shaft <NUM> is also in driving engagement with by an output of the geared architecture <NUM>.

As shown in <FIG>, the geared architecture <NUM> provides an output through a carrier <NUM> that is configured to rotate with the fan drive shaft <NUM>. The carrier <NUM> supports multiple planet gears <NUM> supported for rotation on bearings <NUM>, such as journal bearings, relative to the carrier <NUM>. The planet gears <NUM> also surround a sun gear <NUM> that is in driving engagement with the low pressure turbine <NUM> through the low speed spool <NUM>. A ring gear <NUM> surrounds the planet gears <NUM> and is fixed from rotating relative to the front center body 36A of the engine static structure <NUM>.

The second bearing 38B is supported by an intermediate case 36B of the engine static structure <NUM>. The intermediate case is located axially aft of the low pressure compressor <NUM> and axially forward of the high pressure compressor <NUM> and the high speed spool <NUM>. An inner race of the second bearing 38B is configured to rotate with the low speed spool <NUM> and an outer race of the second bearing 38B is fixed relative to the intermediate case 36B.

The intermediate case 36B at least partially define a boundary of the core flow path C fluidly downstream of the low pressure compressor <NUM> and upstream of the high pressure compressor <NUM>. The intermediate case 36B also includes at least one structural support strut <NUM> radially spanning the core flow path C. The structure strut <NUM> can also be surrounded by an airfoil, such as an inlet guide vane, to turn air leaving the fan <NUM> and entering the core flow path C. A contour of the airfoil is determined based on a rotational direction of the fan <NUM> which determines a direction of movement of the air entering the core flow path C.

Although the different non-limiting examples are illustrated as having specific components, the examples of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting examples in combination with features or components from any of the other non-limiting examples.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a fan section (<NUM>) including a fan (<NUM>) with fan blades (<NUM>), wherein said fan section (<NUM>) drives air along a bypass flow path (B) in a bypass duct (<NUM>);
a gear reduction (<NUM>) in driving engagement with the fan (<NUM>), wherein the gear reduction (<NUM>) is a planetary gear system (<NUM>);
a low spool (<NUM>) including a low pressure turbine (<NUM>) driving a low pressure compressor (<NUM>) and driving the gear reduction (<NUM>) to drive the fan (<NUM>) at a speed slower than the low pressure turbine (<NUM>);
a high spool (<NUM>) including a high pressure turbine (<NUM>) driving a high pressure compressor (<NUM>); and
a second low spool support bearing (38B) located axially between the low pressure compressor (<NUM>) and the high pressure compressor (<NUM>),
wherein the planetary gear system (<NUM>) includes a ring gear (<NUM>) fixed from rotating relative to an engine static structure (<NUM>), a sun gear (<NUM>) in driving engagement with an input from the low spool (<NUM>), and a carrier (<NUM>) supporting multiple planet gears (<NUM>) surrounding the sun gear (<NUM>), the planet gears (<NUM>) supported for rotation on bearings (<NUM>) relative to the carrier (<NUM>), and wherein the planetary gear system (<NUM>) provides an output through the carrier (<NUM>) that is configured to rotate to drive the fan (<NUM>),
characterized in that it further comprises a first low spool support bearing (38A) located axially between the low pressure compressor (<NUM>) and the gear reduction (<NUM>).