Patent Publication Number: US-2020300118-A1

Title: Gas turbine engine bearing support structure

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This specification is based upon and claims the benefit of priority from United Kingdom patent application number GB 1903703.5 filed on Mar. 19, 2019, the entire contents of which is incorporated herein by reference. 
     BACKGROUND 
     Technical Field 
     The present disclosure relates to a support structure, particularly to a bearing support structure for a gas turbine engine. 
     Description of the Related Art 
     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&#39;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  38  in  FIG. 3 ). 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. 
     SUMMARY 
     According to a first aspect there is provided a bearing support structure for a gas turbine engine having a longitudinal axis, the bearing support structure comprising: a plurality of stators; a first section depending forwardly from the plurality of stators relative to the longitudinal axis; a second section depending rearwardly from the plurality of stators relative to the longitudinal axis; a first bearing assembly being supported relative to the plurality of stators by the first section; and a second bearing assembly being supported relative to the plurality of stators by the second section; wherein the second section is detachably mounted to the plurality of stators. 
     The first section may be a forward section, e.g. a front cone. The second section may be a rear section, e.g. being rearward of the forward section, such as a rear cone. 
     At least a portion of the first section may be integral with the plurality of stators such that said portion of the first section is not detachable therefrom. 
     The plurality of stators may comprise an integral interface portion and the second section may comprise an opposing interface portion, the second section may be detachably mounted by a plurality of fasteners releasably holding said first and opposing interface portions together. 
     The opposing interface portions may be are annular in form and the plurality of fasteners may be circumferentially spaced. 
     Each fastener may be provided adjacent a stator of the plurality of stators. 
     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 second section may comprise wall sections depending radially inwardly of the plurality of stators so as to define a housing for an internal volume between the first section, the second section and the longitudinal axis. 
     The first section and/or the second section may be substantially conical in form. 
     A gearbox may be mounted radially inside an inner end of the plurality of stators 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 stators that may be angularly spaced about the longitudinal axis. 
     The second section may be detachably mounted to the plurality of stators at an interface adjacent and/or beneath a radially inner end of the plurality of stators. 
     The interface may be annular in form. It may comprises first and second interface portions when viewed in section, said first and second interface portions may be angularly spaced. 
     The second section may be detachably mounted to the stator at a joint (e.g. a second joint). The first section (or a portion thereof) may be mounted to the stator at a first/further joint, e.g. located forward of the second joint. The first joint may be closer to the longitudinal axis than the second joint. 
     The first joint may be at a radial height from the axis that is less than the radial height of a gearbox. The gearbox may be mounted between the first and second bearing assemblies. The gearbox may be mounted to a portion of the first section that is integral with the stator. 
     The first bearing assembly may be provided at a radially inner end of the first section and/or the second bearing assembly may be provided at a radially inner end of the second section. 
     A front cone may depend forwardly of an engine section stator and a portion of the front cone may be integrally formed with the engine section stator and a rear cone may be removably attached to the engine section stator at an interface using fasteners. The front cone may support a first bearing and the rear cone may support a second bearing. 
     According to a second aspect there is provided a gas turbine engine comprising a bearing support structure according to the first aspect. 
     According to a third aspect there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; 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; and a bearing support structure according to the first aspect. 
     Any of the optional features of the claims may be applied to the first and/or further aspects. 
     As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core. 
     Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise 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 input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed). 
     The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft. 
     In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor. 
     The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above. 
     The gearbox may be 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 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. 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 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges. 
     In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s). 
     The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other. 
     The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other. 
     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 0% span position, to a tip at a 100% 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: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. 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 0.28 to 0.32. 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: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, or 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 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 240 cm to 280 cm or 330 cm to 380 cm. 
     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 2500 rpm, for example less than 2300 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 2250 cm to 300 cm (for example 2450 cm to 280 cm or 250 cm to 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 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 3320 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 18600 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 U tip . The work done by the fan blades  13  on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/U tip   2 , where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and U tip  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: 0.28, 0.29, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being) Jkg −1 K −1 /(ms −1 ) 2 ). 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 0.28 to 0.31 or 0.29 to 0.3. 
     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: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, or 17, 17.5, 18, 18.5, 19, 19.5 or 20. 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 13 to 16, or 13 to 15, or 13 to 14. 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: 35, 40, 45, 50, 55, 60, 65, 70, 75. 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 50 to 70. 
     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: 110 Nkg −1 s, 105 Nkg −1 s, 100 Nkg −1 s, 95 Nkg −1 s, 90 Nkg −1 s, 85 Nkg −1 s or 80 Nkg −1 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 80 Nkg −1 s to 100 Nkg −s , or 85 Nkg −1 s to 95 Nkg −1 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: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. 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 330 kN to 420 kN, for example 350 kN to 400 kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C. (ambient pressure 101.3 kPa, temperature 30 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: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. 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: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. 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 1800K to 1950K. 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 14, 16, 18, 20, 22, 24 or 26 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 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. 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 0.7 or above Mach 0.9. 
     Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions at an altitude that is in the range of from 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. 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 0.8; a pressure of 23000 Pa; and a temperature of −55 degrees C. Purely by way of further example, the cruise conditions may correspond to: a forward Mach number of 0.85; a pressure of 24000 Pa; and a temperature of −54 degrees C. (which may be standard atmospheric conditions at 35000 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 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust. 
     The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Embodiments will now be described by way of example only, with reference to the Figures, in which: 
         FIG. 1  is a sectional side view of a gas turbine engine; 
         FIG. 2  is a close up sectional side view of an upstream portion of a gas turbine engine; 
         FIG. 3  is a partially cut-away view of a gearbox for a gas turbine engine; 
         FIG. 4  is a close up sectional side view of a bearing support structure of the present disclosure for a gas turbine engine; and 
         FIG. 5  is a close up sectional side view of a further example of a bearing support structure of the present disclosure for a gas turbine engine. 
     
    
    
     DETAILED DESCRIPTION 
     Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art. 
       FIG. 1  illustrates a gas turbine engine  10  having a principal rotational axis  9 . The engine  10  comprises an air intake  12  and a propulsive fan  23  that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine  10  comprises a core  11  that receives the core airflow A. The engine core  11  comprises, in axial flow series, a low pressure compressor  14 , a high-pressure compressor  15 , combustion equipment  16 , a high-pressure turbine  17 , a low pressure turbine  19  and a core exhaust nozzle  20 . A nacelle  21  surrounds the gas turbine engine  10  and defines a bypass duct  22  and a bypass exhaust nozzle  18 . The bypass airflow B flows through the bypass duct  22 . The fan  23  is attached to and driven by the low pressure turbine  19  via a shaft  26  and an epicyclic gearbox  30 . 
     In use, the core airflow A is accelerated and compressed by the low pressure compressor  14  and directed into the high pressure compressor  15  where further compression takes place. The compressed air exhausted from the high pressure compressor  15  is directed into the combustion equipment  16  where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines  17 ,  19  before being exhausted through the nozzle  20  to provide some propulsive thrust. The high pressure turbine  17  drives the high pressure compressor  15  by a suitable interconnecting shaft  27 . The fan  23  generally provides the majority of the propulsive thrust. The epicyclic gearbox  30  is a reduction gearbox. 
     An exemplary arrangement for a geared fan gas turbine engine  10  is shown in  FIG. 2 . The low pressure turbine  19  (see  FIG. 1 ) drives the shaft  26 , which is coupled to a sun wheel, or sun gear,  28  of the epicyclic gear arrangement  30 . Radially outwardly of the sun gear  28  and intermeshing therewith is a plurality of planet gears  32  that are coupled together by a planet carrier  34 . The planet carrier  34  constrains the planet gears  32  to precess around the sun gear  28  in synchronicity whilst enabling each planet gear  32  to rotate about its own axis. The planet carrier  34  is coupled via linkages  36  to the fan  23  in order to drive its rotation about the principal rotational axis  9 . Radially outwardly of the planet gears  32  and intermeshing therewith is an annulus or ring gear  38  that is coupled, via linkages  40 , to a stationary supporting structure or stator  24 . 
     Note that the terms “low pressure turbine” and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan  23 ) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft  26  with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan  23 ). In some literature, the “low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan  23  may be referred to as a first, or lowest pressure, compression stage. 
     The epicyclic gearbox  30  is shown by way of example in greater detail in  FIG. 3 . Each of the sun gear  28 , planet gears  32  and ring gear  38  comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in  FIG. 3 . There are four planet gears  32  illustrated, although it will be apparent to the skilled reader that more or fewer planet gears  32  may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox  30  generally comprise at least three planet gears  32 . 
     The epicyclic gearbox  30  illustrated by way of example in  FIGS. 2 and 3  is of the planetary type, in that the planet carrier  34  is coupled to an output shaft via linkages  36 , with the ring gear  38  fixed. However, any other suitable type of epicyclic gearbox  30  may be used. By way of further example, the epicyclic gearbox  30  may be a star arrangement, in which the planet carrier  34  is held fixed, with the ring (or annulus) gear  38  allowed to rotate. In such an arrangement the fan  23  is driven by the ring gear  38 . By way of further alternative example, the gearbox  30  may be a differential gearbox in which the ring gear  38  and the planet carrier  34  are both allowed to rotate. 
     It will be appreciated that the arrangement shown in FIGS. and  3  is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox  30  in the engine  10  and/or for connecting the gearbox  30  to the engine  10 . By way of further example, the connections (such as the linkages  36 ,  40  in the  FIG. 2  example) between the gearbox  30  and other parts of the engine  10  (such as the input shaft  26 , the output shaft and the fixed structure  24 ) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of  FIG. 2 . For example, where the gearbox  30  has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in 
       FIG. 2 . 
     Accordingly, the present disclosure 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 and/or alternative 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. 1  has a split flow nozzle  18 ,  20  meaning that the flow through the bypass duct  22  has its own nozzle  18  that is separate to and radially outside the engine core nozzle  20 . 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  22  and the flow through the core  11  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. In some arrangements, the gas turbine engine  10  may not comprise a gearbox  30 . 
     The geometry of the gas turbine engine  10 , and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the principal rotational axis  9 ), a radial direction (in the bottom-to-top direction in  FIG. 1 ), and a circumferential direction (perpendicular to the page in the  FIG. 1  view). The axial, radial and circumferential directions are mutually perpendicular. 
     The following disclosure concerns a support structure (indicated generally as structure  42  in  FIG. 2  and  FIG. 4 ) located behind the fan  23 , i.e. axially between the fan and the low pressure compressor  14 . The support structure  42  supports bearings, to be described in further detail below, on a stator array  24 . 
     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  14 . 
     The stator array  24  is conventionally referred to as the engine section stator. Whilst the stator array is referred to herein as comprising a plurality of stator vanes, e.g. of aerofoil cross section, it may also be referred to in the singular, i.e. as a singular stator structure. The stator  24  may comprises an aerofoil and may extend into the core airflow A flow path upstream of the low pressure compressor  14 . 
     A support structure  42  is shown in  FIG. 4 . The support structure  42  is located in the compressor region of the gas turbine engine  10 . The support structure is located radially inside of the annular intake A of the engine core  11 . The bearing support structure is located generally radially inward of the stator  24 , i.e. between the radially inner end of the stator  24  and the principal rotational axis  9 . 
     The bearing support structure  42  is annular in form, disposed about the principal rotational axis  9 . 
     The bearing support structure is generally located axially between, and supporting, a bearing assembly of the fan  23  and a bearing assembly of a compressor. 
     The support structure  42  provides a housing for an internal area  43  or compartment of the engine  10 . 
     The support structure  42  comprises a first section  48 . The first section  48  is located forward, or upstream, of the support structure  42 . The first section  48  comprises a wall portion  50  extending between a radially inner forward end  44  and the stator  24 . The wall portion  50  is angled obliquely with respect to the principal rotational axis  9 , i.e. is rearwardly slanted or leaning, towards the stator  24 . The wall portion  50  may be angled between 20 and 70 degrees with respect to the principal rotational axis  9 . 
     The first section  48  joins with the stator  24  at a rear/outer end  54  thereof. 
     The wall portion  50  may comprise a curved portion. The wall portion  50  may be curved at a forward portion  52  thereof. The wall portion  50  may comprise a linear portion. The wall portion  50  may comprise a substantially linear/straight portion, e.g. at a central and/or rear portion  54  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  48  may be annular in form, e.g. so as to comprise a conical shape. The first section  48  may comprise a truncated conical (frustoconical) shape. The first section  48  may be substantially rotationally symmetric about the principal rotational axis  9  (i.e. the wall portion  50  is substantially the same throughout rotation about the principal rotational axis  9 ). 
     The first section  48  comprises a support  56  for a fan bearing  58 . The support  56  may be disposed at the forward portion  52  of the support structure and/or at the forward end  44  thereof. The support  56  is connected to a bearing  58  for rotationally supporting the fan  23 . The support  56  may be connected to or integral with the outer race of the bearing  58 . 
     The stator  24  is disposed radially outward from the first section  48 . The stator  24  is disposed axially rearward from the first section. The stator  24  is connected to the rear and/or radially-inner portion  54  of the first section  48  and carries the first section  48 . 
     The stator  24  is integrally formed with the first section  48  in this example. The stator  24  and the first section  48  are manufactured as an integral, single piece, unitary or monolithic component. The stator  24  and the first section  48  may be manufactured as a single casting to form a single integral piece, i.e. an annular/ring piece. The stator  24  and the first section  48  may be manufactured using an additive layer manufacturing technique to form a single integral piece. The stator  24  and the first section  48  may be manufactured using one or more pre-pregs and cured to form an integral piece. 
     In other examples, the stator  24  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. 5 , the stator  24  and the first section  48  may not be integral but have bolted joint  100 , e.g. to make the relatively-large assembly easier to manufacture. The bolted joint  100  would be part way along the wall of the first section  48 , i.e. the front cone. The joint  100  is a low diameter joint relative to the joint/interface  70  on the rear cone, to be described below. In such examples, the interface/joint in the front section  48  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  70 . 
     This bolted joint  100  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. 5 , the front section or front cone  48  may be split into two sections, a front portion  48 A and a rear portion  48 B joined together at the joint  100 . The front portion  48 A extends forwardly of the joint  100  towards the forward end  44 . The rear portion  48 B is joined to (i.e. integral with) the stator  24 , e.g. extending a short distance radially inwardly of the stator  24 . A gearbox  30  may be connected to the rear portion  48 B, e.g. via an intermediate member  102 , so that gearbox torque can be reacted by the stator  24  via the rear portion  48 B of the front cone  48 . In this way torque imposed on the static “ring” gear (Item  38  in  FIG. 3 ) is reacted via the rear portion  48 B but not the front portion  48 A and the joint  100  is removed from the torque path. The intermediate member  102  may depend from the relevant static part, e.g. ring gear  38 , of the gearbox. 
     In either example of  FIGS. 4 and 5 , stators  24  may be disposed circumferentially about the principal rotational axis  9  in an annular fashion to form a ring of stators. The stator  24  may comprise at least twenty stators. The stator  24  may comprise at least thirty or forty stators  24 . 
     In examples described above, the stators  24  are structural, load-bearing members. In other examples, stators  24  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  48  comprising the plurality of stators  24  may form a unitary piece as shown in  FIG. 4  (i.e. to form a substantially conical wall portion comprising an annulus of stators  24 ). 
     The first section  48  in this example is configured such that a load path between the fan bearing  58  and the stator  24  is continuous. In other examples, as shown in  FIG. 5 , the first section  48  comprises a discontinuity/joint  100  such that a load path between the fan bearing  58  and the stator  24  is discontinuous, i.e. extending across the interface  100  within section  48 . 
     In various examples, the first section  48  (e.g. front portion  48 A and/or rear portion  48 B) may comprise a plurality of sectors. The sectors may comprise conical sectors. Each sector may comprise at least one stator  24 . Each sector may comprise a plurality of stators  24 . Each of the first section  48  ( FIG. 4 ) or the rear portion  48 B ( FIG. 5 ) 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  48 . The sectors may be releasably or non-releasably secured. 
     The following description of the support structure applies to the examples of both  FIGS. 4 and 5 . 
     The support structure  42  comprises a second section  60 . The second section  60  is located at the rearward portion  66  or rear end  46  of the support structure  42 . The second section  60  is located rearward with respect to the first section  48  and/or stator. The second section  60  may be located radially inward with respect to at least a portion the first section  48 . The second section  60  may be located radially inward with respect to at least a portion the stator  24 , e.g. an inner end of the stator  24 . The second section  60  is formed separately (i.e. non integrally) with the first section  48  and stator  24 . 
     The second section  60  comprises a wall portion  62  extending between the radially inner rearward portion  66  and a radially outer end  64 , e.g. which is connected to the stator  24  in use. The radially inner rearward portion  66  may comprise a rear end  46  of the support structure  42  as a whole when assembled. 
     The wall portion  62  is angled forwardly/obliquely with respect to the principal rotational axis  9 , e.g. toward the forward end  44  of the support structure. The wall portion  62  may be angled between 20 and 70 degrees with respect to the principal rotational axis  9  towards the forward end  44  of the support structure. 
     The second section  60  may be annular in form, e.g. comprising a conical shape. The second section  60  may comprise a truncated conical (frustoconical) shape. The second section  60  may be substantially rotationally symmetric about the principal rotational axis  9  (i.e. the wall portion  62  is substantially the same throughout rotation about the principal rotational axis  9 ). 
     The second section  60  is releasably connected/secured to the first section  48  to provide a housing or enclosure for the internal area  43  of the engine. The second section  60  may be releasably connected at a radially outer end  64  thereof. The second section may be releasably connected to the stator  24 . The second section may be releasably connected to the radially innermost portion of the stator  24 . The second section may be releasably connected to a rearmost portion or edge of the stator  24 . 
     The second section  60  and stator  24  may be joined at an interface  70 . A first portion/half of the interface  70  may be integrally formed with the stator  24 . A second/opposing portion of the interface may be provided by the second section  60 . 
     The second section  60  comprises an interface formation  68 . The interface formation  68  is configured to engage with a corresponding interface portion provided on the first section  48 . The first section interface portion may be provided on the stator  24 , preferably, on a rearward portion or edge thereof. 
     The interface formation  68  comprises at least one flange. A first flange  72  is configured to engage a first face  76  provided on the stator  24 . The flange may be angled with respect to the wall portion  62 . The flange may be angled between 90 and 180 degrees with respect to the wall portion  62 . A fastening mechanism/member  80  is configured to extend between the first flange  72  and the first face  76  to form a releasable connection therebetween. 
     The fastening mechanism/member  80  may comprise a bolt, although other conventional fasteners could be considered. 
     The interface formation  68  may comprise a further flange  74 . The further flange  74  is configured to engage a corresponding second face provided on the stator. The further flange  74  may be angled with respect to the first flange  72 . The faces of the opposing stator may be correspondingly angled. The angle may be between 45 and 180 degrees. The angle may be between 90 and 180 degrees. The angle may be 135 degrees. 
     Although not shown in the example of  FIG. 4 , a fastening member could be configured to extend between the further flange  74  and the second face  78  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  74  and the second face  78  to form a releasable connection therebetween and/or provide torque resistance. 
     The interface and or interface formations of the stator  24  and second section  60  may comprise a substantially annular shape. The interface formation provided on the first section  48  may comprise a substantially annular edge. Where an annular joint  100  is also provided on the front cone, the diameter of the annular interface  70  is greater than that of interface  100 . 
     A plurality of fastening members  80  are disposed around the annular interface to provide a releasable connection between the first section  48  (i.e. via the stator  24 ) and the second section  60 . 
     The interface formation  68  provided on the second section  60  may comprise a plurality of discrete portions disposed circumferentially around the second section  60 . The interface formation  70  provided on the first section  48  may comprise a plurality of discrete portions disposed circumferentially around the second section  60 , corresponding to the plurality of discrete portions provided on the second portion. One or more fastening members/mechanisms  80  may be provided between the discrete portions to provide a releasable connection between the first section  48  and the second section  60 . 
     The second section  60  comprises a support  84  for a compressor bearing  86 . The compressor may comprise a low pressure, an intermediate pressure compressor or a high pressure compressor. The support  84  may be disposed at the rearward portion  66  of the second section  60 , e.g. at the rear end  46 . The support  84  is connected to a bearing  86  for rotationally supporting the output shaft of a compressor assembly. The support  84  may be connected to, and/or integrally formed with, the outer race of the bearing  86 . 
     The second section  60  is configured such a load path between the compressor bearing  86  and the stator  24  is discontinuous, i.e. extending across the interface between the stator  24  and second section  60 . 
       FIGS. 4 and 5  show a gear box  30  for a geared turbofan engine architecture as described above, although the gearbox may be omitted in other examples. The gearbox is disposed between, and substantially contained within, the first section  48  and the second section  60 . Thus the support structure  42  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  60 , is radially outside the outermost edge of the gearbox  30 . 
     The first section  48  may comprise a support  90  for bearing  92  of the output shaft of the gearbox  30 . The support  90  is connected to a bearing  92  for rotationally supporting the output shaft of the gearbox  30 . The support  90  may be connected to, or integral with, the outer race of the bearing  92 . The support  90  may be positioned at an increased radial distance from the principal rotational axis  9  than the fan bearing support  56 . 
     The support  90  may be supported by a bracket/wall portion  88  extending from the wall portion  50  of the first section  48 . The bracket may extend rearward and/or radially inward from the wall portion  50 . The bracket  88  may be formed integrally with the wall portion  50  or may be formed as a separate component attached to the wall portion  50 . The bracket  88  may be obliquely/rearwardly angled with respect to the longitudinal axis. The bracket  88  may be annular in form. The bracket may comprise a branching wall off the wall portion  50 . 
     The bearing support  90  may help support the first section  48 , i.e. the front cone. 
     The second section  60  may comprise a support  96  for bearing  96  of the input shaft of the gearbox  30 . The support  96  is connected to a bearing  98  for rotationally supporting the input shaft of the gearbox  30 . The support  96  may be connected to, or integral with, the outer race of the bearing  98 . The support  96  may be positioned at substantially the same radial distance from the principal rotational axis  9  as the compressor bearing support  84 , or else may be radially offset therefrom. 
     The support  96  may be supported by a bracket/wall portion  94  extending from the wall portion  62  of the second section  60 . The bracket may extend forward and/or radially inward from the wall portion  62 . The bracket  94  may be formed integrally with the wall portion  62  or may be formed as a separate component attached to the wall portion  62 . The bracket  94  may be obliquely/forwardly angled with respect to the longitudinal axis. The bracket  94  may be annular in form. The bracket may comprise a branching wall off the wall portion  62 . 
     The bearing support  96  may help support the second section  60 , i.e. the rear cone. 
     The support structure  42  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  48  and second  60  sections are mounted as shown and the bearings provide the interface with the relevant rotating shafts. The first section  48  and stator array  24  can be mounted as a common piece to the second section  60  and bolted to rigidly hold the assembly for use. 
     During assembly/maintenance/disassembly, the second section  60  may be separated from the first portion  48  to provide access to an internal area of the gas turbine engine, i.e. radially inside the stator  24 . The user removes the fastening members  80  connecting the first section  48  and the second section  60 . One of the first section  48  or the second section  60  are then removed to expose the internal area provided between the first section  48  and the second section  60 . 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  30  and rear/second section  60  can beneficially be achieved by removal of the front section  48  with the stator  24 . The radial positioning of the interface  68 ,  70  allows a clearance around the gearbox  30  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. 
     Advantages of 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. 
     It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.