Patent Publication Number: US-11391180-B2

Title: Gas turbine engine inner barrel

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is based upon and claims the benefit of priority from GB Patent Application No. GB 1806564.9, filed on 23 Apr. 2018, the entire contents of which are herein incorporated by reference. 
     BACKGROUND 
     Technical Field 
     The present disclosure relates to an inner barrel for a gas turbine engine, the inner barrel bridging an inner ring from which outer guide vanes extend across a bypass duct of the engine and an inner cowl which provides an aerodynamic fairing surrounding the engine core. 
     Description of the Related Art 
     In a turbofan gas turbine engine, a propulsive fan generates two airflows, one which passes through the core engine and one which passes through a surrounding bypass duct. Behind the fan in the bypass duct is a circumferential row of outer guide vanes which straighten out the bypass airflow from the fan. These vanes extend radially outwards from an inner ring which is a rigid structure defining a radially inner surface of the bypass duct. 
     Rearwardly of the plane of the outer guide vanes, the core engine is surrounded by an aerodynamic fairing called an inner cowl. This fairing also defines a radially inner surface of the bypass duct, and typically comprises door sections that can be opened to allow maintenance access to the core engine. 
     An interface structure if therefore needed to bridge the inner ring and the inner cowl. As the inner ring is typically the responsibility of the engine manufacturer, while the configuration of the inner cowl can be the responsibility of the airframer, this structure can take on additional importance. 
     SUMMARY 
     According to a first aspect there is provided a gas turbine engine for an aircraft comprising: 
     an engine core comprising a compressor, a combustor, a turbine, 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, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core; 
     a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct; 
     an inner cowl which provides an aerodynamic fairing surrounding the engine core, the inner cowl being rearwards of and axially spaced from the inner ring, and including one or more door sections which are openable to enable maintenance access to the engine core; and 
     an inner barrel which surrounds the engine core and bridges the inner ring and the inner cowl, the inner barrel having a circumferentially extending rear edge which provides an engagement formation for engagement with the door sections when they are closed. 
     The inner barrel thus not only bridges the inner ring and the inner cowl, but also provides the engagement formation so that when the door sections are fully closed a smooth aeroline can be formed along the radially inner surface of the bypass duct. 
     Optional features of the present disclosure will now be set out. These are applicable singly or in any combination with any aspect of the present disclosure. 
     The inner barrel may be formed as a unitary structure with the inner ring. This can reduce the total weight of the engine and avoid a need for fasteners joining the inner barrel to the inner ring. 
     The inner barrel may have openings for transmission of services therethrough. For example, the services can include air diverted from the bypass airflow for use in turbine case cooling and/or compressed air bled from the compressor for use in aircraft. 
     The gas turbine engine may further comprise an electronic unit such as an electronic core data controller which controls the engine core and/or performs health monitoring of the engine core. Conveniently, the electronic unit can be mounted to the inner barrel. 
     The gas turbine engine may further comprise a surface cooler for cooling engine fluid using the bypass air flow. Conveniently, the surface cooler can be mounted to the inner barrel. For example, the surface cooler can be for cooling an integrated drive generator used to extract electrical power from the engine core. 
     The inner barrel may have one or more drainage apertures for allowing drainage of liquid from inside the engine barrel. 
     The inner barrel may have one or more ventilation holes for ventilating the engine core. 
     Conveniently, the engagement formation for engagement with the door sections when they are closed may be a circumferentially extending V-groove. The door sections can then have corresponding circumferentially extending features which engage with the V-groove when the doors are closed. 
     The rear edge of the inner barrel may be forward of the combustor. 
     An outer surface of the inner barrel typically defines a radially inner surface of the bypass duct. For example, this outer surface can be a surface of a main structural element (e.g. engine fire zone boundary) of the barrel, and/or it can be one or more in-fill panels located in outwardly facing recesses formed in the barrel, and/or it can be one or more fairings supported by a main structural element of the barrel. 
     It is possible for the inner barrel to be formed as a unitary piece. However, another option is for the inner barrel to be formed from plural separate barrel portions. In particular, the inner barrel may be formed as two half barrels located on respective opposite sides of the engine. The half barrels can be spaced apart at the top of the engine by a mounting pylon for mounting the engine to an airframe. For example, the mounting pylon can join to the engine core at a fixture located at top dead centre behind the inner ring. 
     Additionally or alternatively, when the inner barrel is formed as two half barrels on respective opposite sides of the engine, the inner cowl may have two door sections located on the respective opposite sides of the engine, each door section being pivotably openable about a respective pivot line which extends from front to back e.g. along a respective side of the above-mentioned mounting pylon. The two half barrels may then be spaced apart at the bottom of the engine by a keel beam which extends rearwardly from the inner ring at bottom dead centre thereof to provide latching formations rearward of the inner barrel for latching to lower edges of the door sections when they are closed. 
     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. 
     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). 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: 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). 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). 
     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 250 cm to 300 cm (for example 250 cm to 280 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 320 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 1600 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: 0.3, 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). 
     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. 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). The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. 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). 
     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). 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). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 deg 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). 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 16, 18, 20, or 22 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 deg C. 
     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. 
    
    
     
       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  shows schematically a perspective view from the rear of an engine without its nacelle and without its inner cowl; 
         FIG. 5  shows a side view of a half barrel of an inner barrel of the engine of  FIG. 4 ; 
         FIG. 6  shows a cross-section through the inner barrel of  FIG. 5 ; 
         FIG. 7  shows a cross-section through a variant inner barrel; and 
         FIG. 8  shows at top the inner barrel of  FIGS. 4 to 6  attached by fasteners to an inner ring of the engine, and at bottom a variant of the inner barrel formed as a unitary structure with the inner ring. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates a gas turbine engine  10  having a principal rotational axis  9 . The engine is mounted to an airframe, e.g. under a wing, by a mounting pylon  46 . 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 pylon  46  forms an upper bifurcation in the bypass duct where it traverses the duct to join to the engine core  11 . The bypass airflow B flows through the bypass duct  22 , where it is straightened by a row of outer guide vanes  40  before exiting the bypass exhaust nozzle  18 . Rearward of the outer guide vanes  40 , the engine core  10  is surrounded by an inner cowl  41  which provides an aerodynamic fairing defining an inner surface of 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 engine 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  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. 2 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  20 ,  22  meaning that the flow through the bypass duct  22  has its own nozzle that is separate to and radially outside the core engine 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. 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 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. 
       FIG. 4  shows schematically a perspective view from the rear of the engine  10  with its nacelle  21 , inner cowl  41  and pylon  46  removed. A fan case  42  defines an outer surface of the bypass duct  22  and towards the rear of the fan case an inner ring  44  defines an inner surface of the bypass duct  22 . The outer guide vanes  40  extend radially from the inner ring to the fan case, and the engine core  11  projects rearwardly from the plane of the outer guide vanes. A fixture  45  located at top dead centre behind the inner ring provides a connection point for the mounting pylon  46  which mounts the engine to the airframe. 
     The inner cowl  41  can be formed as two door sections, one on either side of the engine  10 , with each door section being pivotably openable about a respective pivot line  54  which extends from front to back along that door section&#39;s side of the pylon  46 . This allows the door sections to be swung upwards and away from the engine core  11  for maintenance access. Conveniently, the top edges of the door sections can form the pivot lines. A keel beam  48  can also be provided, the keel beam extending rearwardly from bottom dead centre of the inner ring  44  to provide latching formations for latching to lower edges of the door sections when they are closed. 
     The inner cowl  41  is axially spaced from the inner ring  44 . An inner barrel  47  bridges this space and provides a gas-washed fireproof structural panel with amalgamated fire sealing, drainage, and mechanical support interface features. The inner barrel  47  is in two halves  47   a ,  47   b  located on opposite sides of the engine. Each half barrel extends circumferentially on its side of the engine from the fixture  45  to the keel beam  48 . In other variants, however, the inner barrel can be formed from more than two components, or can be formed as a single, continuous component (although in that case, the fixture  45  has to be moved or adapted to allow the barrel to continue through top dead centre, and/or the keel beam  48  has to be adapted or removed to allow the barrel to continue through bottom dead centre). Typically, the inner ring extends only a limited length in the axial direction, e.g. such that its rear edge is forward of the combustor of the engine core  11 . This is consistent with providing an overall axially compact engine. 
     The inner barrel can provide some or all of the following functionalities:
         (1) A zone compliant (fire and fluid) boundary between core and bypass zones of the engine.   (2) An interface to the inner cowl  41 , such as circumferentially extending V-groove  49  for engagement, and hence position and load sharing, with the door sections of the inner cowl  41  when they are closed.   (3) An inner aeroline  51  of the bypass duct  22  along its axial station.   (4) Ventilation features for core zone ventilation.   (5) Drainage surfaces and apertures for core zone drainage.   (6) A fire zone compliant interface for turbine case cooling (TCC) air ingestion from the bypass duct  22 .   (7) A fire zone compliant interface(s) for compressed air bleed exit from one of the compressors and optionally for other bleed air systems.   (8) Structural support and positioning features for mounting an electronic unit such as an electronic core data controller which controls the engine core  11  and/or performs health monitoring of the engine core.   (9) Structural support and positioning features for mounting a surface cooler for cooling engine fluid using the bypass air flow. For example, the cooled engine fluid can be used to cool an integrated drive generator for extracting electrical power from the engine core  11 .   (10) Acoustic treatment for noise attenuation.       

       FIG. 5  shows a side view of the half barrel  47   a  and illustrates features of the inner barrel providing some of these functionalities, such as the V-groove  49 .  FIG. 6  shows a cross-section through the inner barrel, with the core and bypass zones of the engine indicated on either side of the zone compliant boundary  50  provided by the barrel. The inner aero line  51  of the bypass duct  22  can then be provided by a fairing supported by the zone compliant boundary. However,  FIG. 7  shows a cross-section through a variant inner barrel in which the zone compliant boundary  50  also forms the inner aeroline of the bypass duct. In this case, any bypass-side components mounted to the barrel can be located in local recesses  56  formed on that side of the barrel, as shown in the inset to  FIG. 7 . 
     The forward edge of the inner barrel can have fastening features, such as a circumferential row of fastener (e.g. screw and/or bolt) holes  52  for fastening the barrel to the inner ring  44 , as shown in the top cross-section of  FIG. 8 . However, another option, shown in the bottom cross-section of  FIG. 8 , is to form the inner barrel as a unitary structure with the inner ring, thereby reducing the total weight of the engine and avoiding a need for fasteners. 
     The zone compliant boundary  50  of the inner barrel can be fabricated from a metal sheet, which can then be forged to a ring providing the V-groove  49  and another ring providing a flange for the bolt holes  52 . However, another option is to form at least the zone compliant boundary the barrel from composite material, e.g. by a laying up process. 
     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.