Patent Publication Number: US-2020291867-A1

Title: Fuel manifold cooling

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This specification is based upon and claims the benefit of priority from United Kingdom patent application number GB1903328.1 filed on 12 Mar. 2019, the entire contents of which are incorporated herein by reference. 
     BACKGROUND 
     Technical Field 
     The present disclosure relates to the handling of fuel in fuel manifolds in gas turbine engines, and provides a gas turbine engine as set out in the appended claims. 
     Description of the Related Art 
     During soakback and some operational conditions, the temperature in fuel manifolds can become high. This may lead to the creation of fuel breakdown products (sometimes referred to as “coking”) within the manifolds that can cause failures in fuel system components. This is typically not a problem during cruise conditions, during which fuel flows through the fuel manifold at a sufficient rate to provide cooling, and in which the fuel flow rate is such that the time taken for fuel to pass through the fuel manifold is relatively short. However, particular problems may be caused at the end of cruise, when fuel flows are reduced and the fuel manifold ceases to be fuel-cooled in a meaningful way, while at the same time heat from the surrounding metal and air, still hot as a result of normal engine operation, soaks into the fuel manifold. 
     Previous approaches to this problem have included: designing the fuel system components to survive the worst case temperatures and coking products; providing recirculation manifolds and valves in order to manage the amount of time fuel spends in the hot zone; providing heat transfer systems or chillers to create or cool a low-temperature working fluid; and/or draining fuel from the fuel manifold post-cruise. However, each of these approaches has disadvantages, for example a cost and/or weight penalty due to the inclusion of additional components. In the case of a system recirculating fuel, there can be an undesirable failure mode if the cooling flow is erroneously directed into the combustor, leading to a hot streak. Systems in which the fuel is drained from the fuel manifold may also lengthen start times, which is undesirable for users, due to the need to prime the system before engine restart. 
     SUMMARY 
     According to a first aspect there is provided a gas turbine engine having an engine core, the engine core comprising: a compressor and a compressor bleed valve in fluid communication with the compressor, the compressor bleed valve configured to release bleed air from the compressor; a combustor comprising a fuel manifold configured to provide fuel to the combustor; and a cooling conduit that is in thermal contact with the fuel manifold, wherein the engine core further comprises a fluid conduit to supply bleed air from the compressor bleed valve to the cooling conduit. 
     In an arrangement, the engine core may comprise first and second compressors, configured such that the second compressor operates at a higher pressure than the first compressor; and the compressor bleed valve is in fluid communication with the first compressor. 
     In an arrangement, the gas turbine engine may comprise a bypass duct configured to carry a bypass airflow; and the gas turbine engine may be configured such that air provided to the cooling conduit of the fuel manifold is exhausted to the bypass duct. 
     In an arrangement, the gas turbine engine may comprise a bypass duct configured to carry a bypass airflow; and the gas turbine engine may further comprise an ancillary compressor configured to provide air to the cooling conduit of the fuel manifold and a conduit configured to supply air from the bypass duct to the ancillary compressor. 
     In an arrangement, the gas turbine engine may further comprise an ancillary compressor configured to increase the pressure of the bleed air supplied to the cooling conduit of the fuel manifold. 
     In an arrangement, the gas turbine engine may comprise a bypass duct configured to carry a bypass airflow; and the gas turbine engine may further comprise a conduit, configured to supply air from the bypass duct to the ancillary compressor. 
     In an arrangement, the gas turbine engine may further comprise a control valve configured to control whether air is supplied to the ancillary compressor from the compressor bleed valve and/or from the bypass duct. 
     In an arrangement, the gas turbine engine may comprise a bypass duct configured to carry a bypass airflow; and the gas turbine engine may further comprise a second ancillary compressor configured to provide air to the cooling conduit of the fuel manifold and a conduit configured to supply air from the bypass duct to the second ancillary compressor. 
     In an arrangement, the engine core may comprise 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; and 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. 
     In an arrangement, the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprising a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft. 
     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, 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 220 cm to 300 cm (for example 240 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 330 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 1800 rpm. 
     In use of the gas turbine engine, the fan (with associated fan blades) rotates about a principal 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, 0.30, 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 engine 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, 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 form 12 to 16, 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 −1 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 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 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 bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. 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; and 
         FIGS. 4, 5, 6, 7 and 8  depict alternative arrangements of part of a gas turbine engine according to the present disclosure. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates a gas turbine engine  10  having a principal rotational axis  9 . The gas turbine 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 an engine 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 (hereafter referred to as a combustor)  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 core 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 combustor  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 core exhaust nozzle  20  to provide some propulsive thrust. The high pressure turbine  17  drives the high pressure compressor  15  by a second core 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 core shaft  26 , which is coupled to a sun wheel, or sun gear,  28  of the epicyclic gearbox  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 first 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 second linkages  40 , to a fixed 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 core 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 disclosure. 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 first 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 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 epicyclic gearbox  30  may be a differential epicyclic 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 gas turbine engine  10  and/or for connecting the gearbox  30  to the gas turbine 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 gas turbine engine  10  (such as the core 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 bypass exit nozzle  18  that is separate to and radially outside the core exhaust 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 engine 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 , or it may comprise a gearbox of a non-epicyclic variety. 
     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. 
     Gas turbine engines may have compressor bleed valves to release pressure from compressor stages within the gas turbine engine core. A compressor bleed valve may be configured to release bleed air from the compressor during some operating states in order to optimise operation of the engine core. 
       FIG. 4  schematically depicts an arrangement of a part of a gas turbine engine, in which bleed air from a compressor bleed valve  50  is utilised rather than exhausted, as is common in gas turbine engines. In particular, as shown, the compressor bleed valve  50  provided to a compressor  14  is connected via a fluid conduit  55  to a fuel manifold  60  that provides fuel to the combustor  16 . The bleed air may be supplied to a cooling conduit  61  that is in thermal contact with the fuel manifold. Cooling air that has passed through the cooling conduit  61  may be exhausted to the bypass duct  22 , for example via an exit conduit  63 . 
     In a gas turbine engine having two compressors  14 ,  15  one of the compressors  14  may operate at a lower pressure than the other compressor  15 . In such an arrangement, the bleed air used to cool the fuel manifold  60  may be bleed air from the compressor  14  that operates at a lower pressure, as shown in  FIG. 4 . This may be beneficial because the compressor bleed valve  50  may be opened at the end of cruise on a gas turbine engine used on an aircraft, when the power generated by the gas turbine engine is reduced, in order to maintain the operating condition of the gas turbine engine. This coincides with a situation in which it may be beneficial to provide cooling to the fuel manifold  60 . 
       FIG. 5  depicts a variation of the arrangement depicted in  FIG. 4 . As shown, in this arrangement, in addition to supplying bleed air to the cooling conduit  61  of the fuel manifold  60 , cooling air is also provided from the bypass duct  22 . As shown, a first ancillary compressor  70  may be provided to supply air through a conduit  71  from the bypass duct  22  to the cooling conduit  61  of the fuel manifold  60 . 
     Such an arrangement may be used to supplement the cooling effect provided by the bleed air. Alternatively or additionally, the cooling air from the bypass duct  22  may be used when cooling is desirable in the fuel manifold  60  but it is not desirable to open the compressor bleed valve  50  in order to maintain the optimum operating state of the gas turbine engine. Alternatively or additionally, it may be desirable to provide cooling air from the bypass duct  22  to the fuel manifold  60  after shutdown of a gas turbine engine, when no bleed air can be provided from the compressor bleed valve  50  because the compressor  14  is not in operation. 
     A controller may be provided in order to determine when it is appropriate to utilise air from the bypass duct  22  and control the first ancillary compressor  70  accordingly. This controller may be used in conjunction with a controller controlling compressor bleed valve  50  or a single controller may control both systems. 
       FIG. 6  depicts a further variation of the arrangement depicted in  FIG. 4 . As shown, in this arrangement, a first ancillary compressor  70  may be provided to increase the pressure of the bleed air from the compressor bleed valve  50  provided to the cooling conduit  61  of the fuel manifold  60 . This may increase the flow of bleed air, and therefore the effect of the cooling. This may be particularly beneficial when the gas turbine engine is operating at low idle powers. 
     As with the arrangement discussed above in relation to  FIG. 5 , a control system may be provided in order to determine when it is appropriate to utilise the first ancillary compressor  70  and/or control the compressor bleed valve  50 . 
       FIG. 7  depicts a further variation in which a first ancillary compressor  70  may be used to provide air from the bypass duct  22  through a conduit  71  to the cooling conduit  61  of the fuel manifold and/or to increase the pressure of bleed air provided from the compressor bleed valve  50  to the cooling conduit  61  of the fuel manifold  60 . In such an arrangement, a control valve  75  may connect the conduit  71  providing air from the bypass duct  22  and the fluid conduit  55  providing bleed air from the compressor bleed valve  50  to the first ancillary compressor  70 . 
     The control valve  75  may be used to control whether air from one or both sources is provided to the cooling conduit  61  of the fuel manifold  60 . The control valve  75  may be configured such that it can operate in any two or more of the following states: providing air from the bypass duct  22 , providing air from the compressor bleed valve  50 , providing air from both sources and closed, providing air from neither source. It will be appreciated that a controller used to control the operation of the first ancillary compressor  70  and/or the compressor bleed valve  50  may also control the operation of the control valve  75 . 
       FIG. 8  depicts an alternative arrangement to  FIG. 7 , in which a first ancillary compressor  70  is provided to increase the pressure of the air provided from the compressor bleed valve  50  to the cooling conduit  61  of the fuel manifold  60  and a second ancillary compressor  80  provides cooling air from the bypass duct  22  to the cooling conduit  61  of the fuel manifold  60 . It should be appreciated that appropriate control of the compressor bleed valve  50  and the first and second ancillary compressors  70 ,  80 , for example using a controller, may enable the provision of the same operating states discussed above in relation to the arrangement depicted in  FIG. 7 . 
     Advantageously, the above described embodiments do not require the inclusion of any additional heat-reducing hardware systems to provide a cooled working fluid to the cooling conduit, Rather, it has been realised that the temperature of the bleed air from the compressor is sufficiently low in comparison to the fuel manifolds that it can provide a useful cooling effect on the fuel manifolds without the need for any additional heat-exchange or cooling apparatus to further lower the temperature of the bleed air. Thus, such a system is simpler, lighter and more reliable than those of the prior art. 
     It will be understood that the disclosure 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.