Patent Description:
In a known configuration, a support structure includes a ring of stationary aerofoils (stators) that are located near the air intake of the engine core. The support structure extends between and supports a fan bearing assembly and a compressor bearing assembly respectively.

A front portion of a conventional support structure is conical in shape so as to define an internal volume within the support structure, i.e. around the fan shaft. As such, the bearing support structure may house engine components for which access is needed for assembly, periodic maintenance and servicing.

In order to access the engine components, a front portion of the support structure is removed. This requires a releasable connection in the forward end of the support structure between the fan bearing assembly and the stator, typically at a large diameter, larger than the gearbox. In the case that very large radial loads are transmitted from the fan into the support structure, movement may occur in the releasable connection (i.e. a bolt may slip or unwind) and therefore compromise the joint's structural duty and ability to locate components. This problem is exacerbated on modern fans which tend to be large in diameter with a low count of wide chord blades, e.g. as may be used in a geared turbofan engine, due to the high loads which may be experienced, particularly during a fan blade-off scenario. Fan blade-off load management may be further complicated in a geared turbofan architecture as typical fusing systems, which fail so as to allow the now out-of-balance fan set to rotate about its new centre of gravity, may not be practical as the space required could detrimentally effect the architectural design, e.g. cause the gearbox design to be very large and heavy.

It is possible to design a bolted joint on the fan load path that can withstand loads generated during a fan blade-off scenario. However such a joint relies on generating sufficient frictional force in the bolted joint to prevent slippage. Any slippage will provide a mechanism to unwind the bolt. Coefficient of friction is difficult to predict and so must be assumed to be low. The resulting joints are very large and heavy, which is contrary to the general aim of weight reduction and greater fuel efficiency within the aerospace sector.

A specific additional function of a bearing support structure containing an epicyclic gearing mechanism, such as typically used in a geared turbofan, is to react the torque imposed on the static "ring" gear (Item <NUM> in <FIG>). The torque reaction load travels through the support structure to the engine casings and mounts. With the torque reaction load being passed into the support structure, this may further drive the design of the releasable connection to also be capable of transferring this torque. Again a suitable bolted joint will become larger and heavier to accommodate this.

It is an aim of the present disclosure to provide a support structure configuration that addresses one of more the aforementioned problems or at least provides a useful alternative to known support structure configurations.

<CIT> relates to a method for assembling a turbofan engine comprising coupling a bearing assembly and a shaft as a unit to a bearing support. A transmission and the fan shaft are installed to a front frame assembly. The bearing assembly and shaft are installed as a unit so that the shaft engages a central gear of the transmission and the bearing support engages the front frame assembly. A fan rotor bearing support structure forward of the transmission comprises a forward web and an aft web which join at a root of an outboard/aft mounting flange, The outboard/aft mounting flange is radially outboard of the periphery of the transmission.

<CIT> relates to a geared architecture for a gas turbine engine comprising an output shaft for connection with a fan, an input shaft and a gearbox connecting the input shaft with the output shaft. The gearbox has a forward planet carrier plate supported by a forward radially fixed bearing structure and a rearward planet carrier plate supported by a rearward radially fixed bearing structure.

<CIT> relates to a bearing support for a rotor of an aircraft turbine engine that includes a front bearing and a front bearing support and a bearing strut for attaching the front bearing to the aircraft turbine engine support structure.

<CIT> relates to a gas turbine engine including an aft bearing assembly including at least two bearings that is positioned at least partially in an aft sump and supports a shaft in the turbine section.

<CIT> relates to a gas turbine engine having modules including a gear reduction module and a bearing mounting assembly.

<CIT> relates to a mounting system for a planetary geartrain in a gas turbine engine that comprises a support strut, a deflection flange and a deflection limiter.

The present invention provides a gas turbine engine as set out in the appended claims.

The first bearing assembly may comprise a fan bearing assembly and/or the second bearing assembly may comprise a compressor bearing assembly.

The first section and/or the second section may be substantially conical in form.

The gearbox may be mounted radially inside an inner end of the plurality of stator vanes and/or within the axial extent of the first section and/or the second section.

The first section may comprise a support for a gearbox output bearing and/or the second section may comprise a support for a gearbox input bearing.

The bearing support structure may comprise an array of at least twenty stator vanes that may be circumferentially spaced about the longitudinal axis.

Arrangements of the present invention are particularly beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine comprises a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The gearbox is a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than <NUM>, for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>, for example on the order of or at least <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from <NUM> or <NUM> to <NUM>. In some arrangements, the gear ratio may be outside these ranges.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or <NUM>% span position, to a tip at a <NUM>% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: <NUM>, <NUM>, <NUM><NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from <NUM> to <NUM>. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: <NUM>, <NUM>, <NUM>, <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches) cm, or <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches) or <NUM> (around <NUM> inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from <NUM> to <NUM> or <NUM> to <NUM>.

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than <NUM> rpm, for example less than <NUM> rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from <NUM> to <NUM> (for example <NUM> to <NUM> or <NUM> to <NUM>) may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from <NUM> to <NUM> may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades <NUM> on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip<NUM>, where dH is the enthalpy rise (for example the <NUM>-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> (all units in this paragraph being Jkg-<NUM>K-<NUM>/(ms-<NUM>)<NUM>). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from <NUM> to <NUM> or <NUM> to <NUM>.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from <NUM> to <NUM>, or <NUM> to <NUM>, or <NUM> to <NUM>. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine core. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from <NUM> to <NUM>.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: <NUM> Nkg-'s, <NUM> Nkg-'s, <NUM> Nkg-'s, <NUM> Nkg-'s, <NUM> Nkg-'s, <NUM> Nkg-<NUM>s or <NUM> Nkg-'s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from <NUM> Nkg-'s to <NUM> Nkg-<NUM>s, or <NUM> Nkg-<NUM>s to <NUM> Nkg-'s. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to <NUM> kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus <NUM> degrees C (ambient pressure <NUM>. 3kPa, temperature <NUM> degrees C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from <NUM> to <NUM>. The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> fan blades.

As used herein, cruise conditions may mean cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example on the order of Mach <NUM>, on the order of Mach <NUM> or in the range of from <NUM> to <NUM>. Any single speed within these ranges may be the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach <NUM> or above Mach <NUM>.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions at an altitude that is in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM> (around <NUM> ft), for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> (around <NUM> ft) to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example on the order of <NUM>. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to: a forward Mach number of <NUM>; a pressure of <NUM> Pa; and a temperature of -<NUM> degrees C. Purely by way of further example, the cruise conditions may correspond to: a forward Mach number of <NUM>; a pressure of <NUM> Pa; and a temperature of -<NUM> degrees C (which may be standard atmospheric conditions at <NUM> ft).

As used anywhere herein, "cruise" or "cruise conditions" may mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach Number, environmental conditions and thrust requirement) for which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to have optimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example <NUM> or <NUM>) gas turbine engine may be mounted in order to provide propulsive thrust.

The planet carrier <NUM> is coupled via linkages <NUM> to the fan <NUM> in order to drive its rotation about the principal rotational axis <NUM>. Radially outwardly of the planet gears <NUM> and intermeshing therewith is an annulus or ring gear <NUM> that is coupled, via linkages <NUM>, to a stationary supporting structure or stator <NUM>.

Accordingly, the present invention extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

Optionally, the gearbox may drive additional components (e. g the intermediate pressure compressor and/or a booster compressor).

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in <FIG> has a split flow nozzle <NUM>, <NUM> meaning that the flow through the bypass duct <NUM> has its own nozzle <NUM> that is separate to and radially outside the engine core nozzle <NUM>. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct <NUM> and the flow through the core <NUM> are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example.

The geometry of the gas turbine engine <NUM>, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the principal rotational axis <NUM>), a radial direction (in the bottom-to-top direction in <FIG>), and a circumferential direction (perpendicular to the page in the <FIG> view).

The following invention concerns a support structure (indicated generally as structure <NUM> in <FIG> and <FIG>) located behind the fan <NUM>, i.e. axially between the fan and the low pressure compressor <NUM>. The support structure <NUM> supports bearings, to be described in further detail below, on a stator array <NUM>.

In general terms, the support structure supports both a fan bearing, i.e. a bearing for the fan shaft. Additionally the support structure may support a compressor bearing, i.e. a bearing for a rotor shaft of the compressor <NUM>.

The stator array <NUM> is conventionally referred to as the engine section stator.

The stator array is referred to herein as comprising a plurality of stator vanes, e.g. of aerofoil cross section extending into the core airflow A flow path upstream of the low pressure compressor <NUM>.

A support structure <NUM> is shown in <FIG>. The support structure <NUM> is located in the compressor region of the gas turbine engine <NUM>. The support structure is located radially inside of the annular intake A of the engine core <NUM>. The bearing support structure is located generally radially inward of the stator <NUM>, i.e. between the radially inner end of the stator <NUM> and the principal rotational axis <NUM>.

The bearing support structure <NUM> is annular in form, disposed about the principal rotational axis <NUM>.

The bearing support structure is generally located axially between, and supporting, a bearing assembly of the fan <NUM> and a bearing assembly of a compressor.

The support structure <NUM> provides a housing for an internal area <NUM> or compartment of the engine <NUM>.

The support structure <NUM> comprises a first section <NUM>. The first section <NUM> is located forward, or upstream, of the support structure <NUM>. The first section <NUM> comprises a wall portion <NUM> extending between a radially inner forward end <NUM> and the stator <NUM>. The wall portion <NUM> is angled obliquely with respect to the principal rotational axis <NUM>, i.e. is rearwardly slanted or leaning, towards the stator <NUM>. The wall portion <NUM> may be angled between <NUM> and <NUM> degrees with respect to the principal rotational axis <NUM>.

The first section <NUM> joins with the stator <NUM> at a rear/outer end <NUM> thereof.

The wall portion <NUM> may comprise a curved portion. The wall portion <NUM> may be curved at a forward portion <NUM> thereof. The wall portion <NUM> may comprise a linear portion. The wall portion <NUM> may comprise a substantially linear/straight portion, e.g. at a central and/or rear portion <NUM> thereof. The wall section may comprise one or more of: a linear portion, a curved portion, or a polygonal portion and combinations thereof.

The first section <NUM> may be annular in form, e.g. so as to comprise a conical shape. The first section <NUM> may comprise a truncated conical (frustoconical) shape. The first section <NUM> may be substantially rotationally symmetric about the principal rotational axis <NUM> (i.e. the wall portion <NUM> is substantially the same throughout rotation about the principal rotational axis <NUM>).

The first section <NUM> comprises a support <NUM> for a fan bearing <NUM>. The support <NUM> may be disposed at the forward portion <NUM> of the support structure and/or at the forward end <NUM> thereof. The support <NUM> is connected to a bearing <NUM> for rotationally supporting the fan <NUM>. The support <NUM> may be connected to or integral with the outer race of the bearing <NUM>.

The stator <NUM> is disposed radially outward from the first section <NUM>. The stator <NUM> is disposed axially rearward from the first section. The stator <NUM> is connected to the rear and/or radially-inner portion <NUM> of the first section <NUM> and carries the first section <NUM>.

The stator <NUM> is integrally formed with the first section <NUM> in this example. The stator <NUM> and the first section <NUM> are manufactured as an integral, single piece, unitary or monolithic component. The stator <NUM> and the first section <NUM> may be manufactured as a single casting to form a single integral piece, i.e. an annular/ring piece. The stator <NUM> and the first section <NUM> may be manufactured using an additive layer manufacturing technique to form a single integral piece. The stator <NUM> and the first section <NUM> may be manufactured using one or more pre-pregs and cured to form an integral piece.

In other examples, the stator <NUM> and first section could be formed as a fabrication of cast, forged or ALM portions. The relevant portions may be bonded, welded or fused together so as to form a unitary structure which is indivisible without damage to the structure.

In other examples, e.g. as shown in <FIG>, the stator <NUM> and the first section <NUM> may not be integral but have bolted joint <NUM>, e.g. to make the relatively-large assembly easier to manufacture. The bolted joint <NUM> would be part way along the wall of the first section <NUM>, i.e. the front cone. The joint <NUM> is a low diameter joint relative to the joint/interface <NUM> on the rear cone, to be described below. In such examples, the interface/joint in the front section <NUM> is not at a radial height sufficient to allow assembly, servicing and maintenance of the gearbox. That is to say the interface may be inaccessible from the front and/or closer to the axis (e.g. at a lower radial height) than the interface <NUM>.

This bolted joint <NUM> may not be on the torque path, e.g. making it easier to design to accommodate ultimate events, such as a fan blade-off. As shown in <FIG>, the front section or front cone <NUM> may be split into two sections, a front portion 48A and a rear portion 48B joined together at the joint <NUM>. The front portion 48A extends forwardly of the joint <NUM> towards the forward end <NUM>. The rear portion 48B is joined to (i.e. integral with) the stator <NUM>, e.g. extending a short distance radially inwardly of the stator <NUM>. A gearbox <NUM> may be connected to the rear portion 48B, e.g. via an intermediate member <NUM>, so that gearbox torque can be reacted by the stator <NUM> via the rear portion 48B of the front cone <NUM>. In this way torque imposed on the static "ring" gear (Item <NUM> in <FIG>) is reacted via the rear portion 48B but not the front portion 48A and the joint <NUM> is removed from the torque path. The intermediate member <NUM> may depend from the relevant static part, e.g. ring gear <NUM>, of the gearbox.

In either example of <FIG> and <FIG>, stators <NUM> are disposed circumferentially about the principal rotational axis <NUM> in an annular fashion to form a ring of stators. The stator <NUM> may comprise at least twenty stators. The stator <NUM> may comprise at least thirty or forty stators <NUM>.

In examples described above, the stators <NUM> are structural, load-bearing members. In other examples, stators <NUM> may be axially separated into multiple narrow aerofoils that do not carry structural load and load carrying struts, i.e. which typically do not have an aerodynamic duty. The struts may be fore and/or aft of the aerofoils. Typically there would be at least three struts. The term 'stator(s)' as used herein refers to aerofoil(s)/vane(s) and/or strut(s), depending on which members carry the structural load.

The first section <NUM> comprising the plurality of stators <NUM> may form a unitary piece as shown in <FIG> (i.e. to form a substantially conical wall portion comprising an annulus of stators <NUM>).

The first section <NUM> in this example is configured such that a load path between the fan bearing <NUM> and the stator <NUM> is continuous. In other examples, as shown in <FIG>, the first section <NUM> comprises a discontinuity/joint <NUM> such that a load path between the fan bearing <NUM> and the stator <NUM> is discontinuous, i.e. extending across the interface <NUM> within section <NUM>.

In various examples, the first section <NUM> (e.g. front portion 48A and/or rear portion 48B) may comprise a plurality of sectors. The sectors may comprise conical sectors. Each sector may comprise at least one stator <NUM>. Each sector may comprise a plurality of stators <NUM>. Each of the first section <NUM> (<FIG>) or the rear portion 48B (<FIG>) may be manufactured in single cast or die, or additive layer manufacturing process, to form an integral piece. The plurality sectors are then secured together to form the first section <NUM>. The sectors may be releasably or non-releasably secured.

The following description of the support structure applies to the examples of both <FIG> and <FIG>.

The support structure <NUM> comprises a second section <NUM>. The second section <NUM> is located at the rearward portion <NUM> or rear end <NUM> of the support structure <NUM>. The second section <NUM> is located rearward with respect to the first section <NUM> and/or stator. The second section <NUM> may be located radially inward with respect to at least a portion the first section <NUM>. The second section <NUM> may be located radially inward with respect to at least a portion the stator <NUM>, e.g. an inner end of the stator <NUM>. The second section <NUM> is formed separately (i.e. non integrally) with the first section <NUM> and stator <NUM>.

The second section <NUM> comprises a wall portion <NUM> extending between the radially inner rearward portion <NUM> and a radially outer end <NUM>, e.g. which is connected to the stator <NUM> in use. The radially inner rearward portion <NUM> may comprise a rear end <NUM> of the support structure <NUM> as a whole when assembled.

The wall portion <NUM> is angled forwardly/obliquely with respect to the principal rotational axis <NUM>, e.g. toward the forward end <NUM> of the support structure. The wall portion <NUM> may be angled between <NUM> and <NUM> degrees with respect to the principal rotational axis <NUM> towards the forward end <NUM> of the support structure.

The second section <NUM> may be annular in form, e.g. comprising a conical shape. The second section <NUM> may comprise a truncated conical (frustoconical) shape. The second section <NUM> may be substantially rotationally symmetric about the principal rotational axis <NUM> (i.e. the wall portion <NUM> is substantially the same throughout rotation about the principal rotational axis <NUM>).

The second section <NUM> is releasably connected/secured to the first section <NUM> to provide a housing or enclosure for the internal area <NUM> of the engine. The second section <NUM> may be releasably connected at a radially outer end <NUM> thereof. The second section is releasably connected to the stator <NUM>. The second section may be releasably connected to the radially innermost portion of the stator <NUM>. The second section may be releasably connected to a rearmost portion or edge of the stator <NUM>.

The second section <NUM> and stator <NUM> are joined at an interface <NUM>. A first portion/half of the interface <NUM> may be integrally formed with the stator <NUM>. A second/opposing portion of the interface may be provided by the second section <NUM>.

The second section <NUM> comprises an interface formation <NUM>. The interface formation <NUM> is configured to engage with a corresponding interface portion provided on the first section <NUM>. The first section interface portion may be provided on the stator <NUM>, preferably, on a rearward portion or edge thereof.

The interface formation <NUM> comprises at least one flange. A first flange <NUM> is configured to engage a first face <NUM> provided on the stator <NUM>. The flange may be angled with respect to the wall portion <NUM>. The flange may be angled between <NUM> and <NUM> degrees with respect to the wall portion <NUM>. A fastening mechanism/member <NUM> is configured to extend between the first flange <NUM> and the first face <NUM> to form a releasable connection therebetween.

The fastening mechanism/member <NUM> may comprise a bolt, although other conventional fasteners could be considered.

The interface formation <NUM> may comprise a further flange <NUM>. The further flange <NUM> is configured to engage a corresponding second face provided on the stator. The further flange <NUM> may be angled with respect to the first flange <NUM>. The faces of the opposing stator may be correspondingly angled. The angle may be between <NUM> and <NUM> degrees. The angle may be between <NUM> and <NUM> degrees. The angle may be <NUM> degrees.

Although not shown in the example of <FIG>, a fastening member could be configured to extend between the further flange <NUM> and the second face <NUM> to form a releasable connection therebetween. The fastening mechanism may comprise a bolt, or other conventional fastener. One or more dowels may be configured to extend between the further flange <NUM> and the second face <NUM> to form a releasable connection therebetween and/or provide torque resistance.

The interface and or interface formations of the stator <NUM> and second section <NUM> may comprise a substantially annular shape. The interface formation provided on the first section <NUM> may comprise a substantially annular edge. Where an annular joint <NUM> is also provided on the front cone, the diameter of the annular interface <NUM> is greater than that of interface <NUM>.

A plurality of fastening members <NUM> are disposed around the annular interface to provide a releasable connection between the first section <NUM> (i.e. via the stator <NUM>) and the second section <NUM>.

The interface formation <NUM> provided on the second section <NUM> may comprise a plurality of discrete portions disposed circumferentially around the second section <NUM>. The interface formation <NUM> provided on the first section <NUM> may comprise a plurality of discrete portions disposed circumferentially around the second section <NUM>, corresponding to the plurality of discrete portions provided on the second portion. One or more fastening members/mechanisms <NUM> may be provided between the discrete portions to provide a releasable connection between the first section <NUM> and the second section <NUM>.

The second section <NUM> comprises a support <NUM> for a compressor bearing <NUM>. The compressor may comprise a low pressure, an intermediate pressure compressor or a high pressure compressor. The support <NUM> may be disposed at the rearward portion <NUM> of the second section <NUM>, e.g. at the rear end <NUM>. The support <NUM> is connected to a bearing <NUM> for rotationally supporting the output shaft of a compressor assembly. The support <NUM> may be connected to, and/or integrally formed with, the outer race of the bearing <NUM>.

The second section <NUM> is configured such a load path between the compressor bearing <NUM> and the stator <NUM> is discontinuous, i.e. extending across the interface between the stator <NUM> and second section <NUM>.

<FIG> and <FIG> show a gear box <NUM> for a geared turbofan engine architecture as described above. The gearbox is disposed between, and substantially contained within, the first section <NUM> and the second section <NUM>. Thus the support structure <NUM> forms a housing around, e.g. circumferentially around, the gearbox.

The radial positioning of the interface, i.e. as defined by the radial height of the second section <NUM>, is radially outside the outermost edge of the gearbox <NUM>.

The first section <NUM> may comprise a support <NUM> for bearing <NUM> of the output shaft of the gearbox <NUM>. The support <NUM> is connected to a bearing <NUM> for rotationally supporting the output shaft of the gearbox <NUM>. The support <NUM> may be connected to, or integral with, the outer race of the bearing <NUM>. The support <NUM> may be positioned at an increased radial distance from the principal rotational axis <NUM> than the fan bearing support <NUM>.

The support <NUM> may be supported by a bracket/wall portion <NUM> extending from the wall portion <NUM> of the first section <NUM>. The bracket may extend rearward and/or radially inward from the wall portion <NUM>. The bracket <NUM> may be formed integrally with the wall portion <NUM> or may be formed as a separate component attached to the wall portion <NUM>. The bracket <NUM> may be obliquely/rearwardly angled with respect to the longitudinal axis. The bracket <NUM> may be annular in form. The bracket may comprise a branching wall off the wall portion <NUM>.

The bearing support <NUM> may help support the first section <NUM>, i.e. the front cone.

The second section <NUM> may comprise a support <NUM> for bearing <NUM> of the input shaft of the gearbox <NUM>. The support <NUM> is connected to a bearing <NUM> for rotationally supporting the input shaft of the gearbox <NUM>. The support <NUM> may be connected to, or integral with, the outer race of the bearing <NUM>. The support <NUM> may be positioned at substantially the same radial distance from the principal rotational axis <NUM> as the compressor bearing support <NUM>, or else may be radially offset therefrom.

The support <NUM> may be supported by a bracket/wall portion <NUM> extending from the wall portion <NUM> of the second section <NUM>. The bracket may extend forward and/or radially inward from the wall portion <NUM>. The bracket <NUM> may be formed integrally with the wall portion <NUM> or may be formed as a separate component attached to the wall portion <NUM>. The bracket <NUM> may be obliquely/forwardly angled with respect to the longitudinal axis. The bracket <NUM> may be annular in form. The bracket may comprise a branching wall off the wall portion <NUM>.

The bearing support <NUM> may help support the second section <NUM>, i.e. the rear cone.

The support structure <NUM> may comprise any conventional materials, e.g. metallic, polymer and/or composite materials. The composite material may comprise a fibre reinforced polymer, a metal matrix composite, a ceramic matrix composite or combinations thereof.

In normal use, the first <NUM> and second <NUM> sections are mounted as shown and the bearings provide the interface with the relevant rotating shafts. The first section <NUM> and stator array <NUM> can be mounted as a common piece to the second section <NUM> and bolted to rigidly hold the assembly for use.

During assembly/maintenance/disassembly, the second section <NUM> may be separated from the first portion <NUM> to provide access to an internal area of the gas turbine engine, i.e. radially inside the stator <NUM>. The user removes the fastening members <NUM> connecting the first section <NUM> and the second section <NUM>. One of the first section <NUM> or the second section <NUM> are then removed to expose the internal area provided between the first section <NUM> and the second section <NUM>. This may provide access to inter alia: a gearbox; the shaft system; shaft-system components; the bearings located within the first and/or second sections; or any other components/accessories mounted within the support structure.

Access to the gearbox <NUM> and rear/second section <NUM> can beneficially be achieved by removal of the front section <NUM> with the stator <NUM>. The radial positioning of the interface <NUM>, <NUM> allows a clearance around the gearbox <NUM> when removing the front cone.

Whilst, for ungeared gas turbine engines there is generally no need to access the space within the bearing support structure, as the bearings and other components are typically positioned on the front and rear sides of the structure as well as the inner bore diameter, rather than the enclosed zone within the bearing support structure.

The support structure reduces the risk of the connection between the first section and the second section failing in the case of radially asymmetric loading of the fan assembly. This can occur, for example due to a fan blade-off or compressor blade-off scenario.

The provision of a joint (i.e. bolted interface) behind the engine section stator removes the joint from the fan load path and instead places it in the compressor load path. This reduces the potential loading the joint is required to withstand, thereby permitting a reduction in the size/weight of the joint assembly and associated bolts.

The lack of a bolted joint in the fan load path may reduce the risk of failure or bolt unwinding, e.g. under large fan blade-off loading.

The support structure may remove the joint from the torque reaction path therefore simplifying the design and reducing the chance of failure, e.g. when using an epicyclic gearbox in which the gearbox ring gear is mounted to the fan load path.

The support structure provides convenient access to the internal area of the gas turbine engine.

Claim 1:
A gas turbine engine having a longitudinal axis (<NUM>) comprising :
an engine core (<NUM>) comprising a turbine (<NUM>), a compressor (<NUM>), and a core shaft (<NUM>) connecting the turbine to the compressor;
a fan (<NUM>) located upstream of the engine core (<NUM>), the fan comprising a plurality of fan blades;
a gearbox (<NUM>) that receives an input from the core shaft (<NUM>) and outputs drive to the fan (<NUM>) so as to drive the fan at a lower rotational speed than the core shaft (<NUM>); and
a bearing support structure (<NUM>), the bearing support structure (<NUM>) comprising:
a plurality of stator vanes (<NUM>) disposed circumferentially about the longitudinal axis (<NUM>) in an annular fashion to form a ring of stator vanes;
a first section (<NUM>) depending forwardly from the plurality of stator vanes (<NUM>) relative to the longitudinal axis (<NUM>);
a second section (<NUM>) depending rearwardly from the plurality of stator vanes (<NUM>) relative to the longitudinal axis (<NUM>);
a first bearing assembly (<NUM>) being supported relative to the plurality of stator vanes (<NUM>) by the first section (<NUM>); and
a second bearing assembly (<NUM>) being supported relative to the plurality of stator vanes (<NUM>) by the second section (<NUM>), wherein:
the second section (<NUM>) is detachably mounted to the plurality of stator vanes (<NUM>) at a joint (<NUM>), the radial positioning of the joint (<NUM>) being radially outside the outermost edge of the gearbox (<NUM>);
the first section (<NUM>) and second section (<NUM>) comprise wall sections depending radially inwardly of the plurality of stator vanes (<NUM>) so as to define a housing for an internal volume between the first section, the second section and the longitudinal axis (<NUM>), the housing being formed by the support structure (<NUM>) around the gearbox; and
either:
(i) the stator vanes (<NUM>) are integrally formed with the first section (<NUM>); or
(ii) the first section (<NUM>) is mounted to the stator vanes (<NUM>) at a further joint, the further joint being closer to the longitudinal axis than the joint at which the second section (<NUM>) is detachably mounted to the plurality of stator vanes (<NUM>), with the further joint being at a radial height from the longitudinal axis that is less than the radial height of the gearbox.