Patent Publication Number: US-2020291868-A1

Title: Louvre system

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This specification is based upon and claims the benefit of priority from UK Patent Application Number 1903465.1 filed on Mar. 14, 2019, the entire contents of which are incorporated herein by reference. 
     BACKGROUND 
     1. Field of the Disclosure 
     The present disclosure relates to a louvre system for a gas turbine engine. In particular, but not exclusively, the louvre system may be used to extract from a compressor, and introduce into a bypass duct, bleed air; or to extract air from the bypass duct and introduce it into a conduit to provide cooling air for a turbine or other systems of the gas turbine engine. 
     2. Description of the Related Art 
     Bleed valves, as fitted on the external casing of intermediate and high pressure compressors of a gas turbine engine, are used to force excess compressor air into the bypass duct of the engine. Bleed valves may be located downstream of the outlet guide vanes of the engine, with their axial position determined by the engine architecture. In typical current engine architectures, the axial location of a bleed valve may be close to the trailing edge of an outlet guide vane, and may be directly in front of a bifurcation in airflow caused by a structural pylon component in the bypass duct. 
     Bleed valves may be scheduled to operate over a range of operating conditions, and may in some circumstances be open during high speed, high thrust flight conditions. In such conditions, high temperature air is blown into the bypass duct, which produces significant flow distortions in the bypass flow. These flow distortions can propagate upstream and lead to the buffeting of the outlet guide vanes and can increase fan forcing to an unacceptable level. 
     SUMMARY 
     According to a first aspect there is provided a louvre assembly for a gas turbine engine, the louvre assembly comprising: 
     an air discharge opening; 
     a first louvre extending across a first portion of the discharge opening and comprising a first plurality of slats each pivotably mounted to rotate about a first direction between a closed position to obstruct air flowing through the first portion of the discharge opening and an open position to allow air to flow through the first portion of the air discharge opening; and 
     a second louvre extending across a second portion of the air discharge opening and comprising a second plurality of slats each pivotably mounted to rotate about a second direction between a closed position to obstruct air flowing through the second portion of the air discharge opening and an open position to allow air to flow through the second portion of the air discharge opening, 
     wherein the first and second directions are angled relative to each other such that air exiting through the air discharge opening with the first and second plurality of slats in the open position diverges away from a central axis of the louvre assembly extending between the first and second portions of the air discharge opening. 
     For example, the louvre assembly of the first aspect may be used in a compressor bleed system. Configuring the louvre assembly such that bleed air exits in diverging directions allows bleed air to be introduced into, or extracted from, the bypass duct of an engine with minimal disruption to air flow. 
     The first plurality of slats may be each pivotably mounted to rotate about a respective first axis. The first axes may be parallel to the first direction. 
     The second plurality of slats may be each pivotably mounted to rotate about a respective second axis. The second axes may be parallel to the second direction. 
     An angle between the first direction and the second direction may for example be between around 5 degrees and around 30 degrees from parallel. 
     An angle between the first direction and the central axis may be between around 75 degrees and 85 degrees. 
     An angle between the second direction and the central axis may be between around 75 degrees and 85 degrees. 
     The first and second louvres may be symmetrically arranged about the central axis. 
     The louvre assembly may further comprise an actuation mechanism connecting the first and second louvres for synchronous rotation between the open and closed positions. The actuation mechanism may comprise an actuator connected to an actuation rod configured to cause rotation of the first and second pluralities of slats between the open and closed positions. 
     The first and second pluralities of slats may be arranged to be aligned in the open position at an angle of between 10 and 45 degrees relative to a plane extending across the air discharge opening. For example, the first and second pluralities of slats may be arranged to be aligned in the open position at an angle of between 30 and 45 degrees relative to a plane extending across the air discharge opening. In an embodiment, for example for introducing bleed air into a bypass duct, the angle may be between 30 and 45 degrees. In a further embodiment, for example for extracting air from a bypass duct, the angle may be between 10 and 35 degrees. 
     In accordance with a second aspect there is provided a gas turbine engine for an aircraft 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 bypass duct downstream of the fan; 
     a bleed conduit arranged to receive bleed air from the compressor; and 
     a louvre assembly according to the first aspect, 
     wherein the discharge opening is arranged to direct bleed air from the bleed conduit into the bypass duct with the first and second pluralities of slats in the open position. 
     The gas turbine engine may further comprise a plurality of fan outlet guide vanes and a pylon, arranged in the bypass duct downstream of the fan outlet guide vanes. The air discharge opening may be arranged downstream of the fan outlet guide vanes and upstream of the pylon. 
     The central axis of the louvre assembly may be aligned with a central axis of the pylon. 
     In accordance with a third aspect there is provided a gas turbine engine for an aircraft 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 bypass duct downstream of the fan; 
     a conduit arranged to receive air from the bypass duct; and 
     a louvre assembly according to the first aspect, 
     wherein the discharge opening is arranged to extract air from the bypass duct with the first and second pluralities of slats in the open position and direct the extract air into the conduit. 
     The conduit may provide cooling air for the turbine. 
     The gas turbine engine may further 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. 
     With the turbine a first turbine, the compressor a first compressor, and the core shaft 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 being 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 is a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a “planetary” or “star” gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, 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. A higher gear ratio may be more suited to “planetary” style gearbox. 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). 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, 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). 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 −l 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 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 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. 
    
    
     
       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 schematic plan view of an example louvre assembly; 
         FIG. 5  is a schematic sectional view through a bypass duct of a gas turbine engine, with a louvre assembly positioned behind a series of outlet guide vanes; 
         FIG. 6  is a schematic sectional view of an example louvre assembly in position for providing a bleed air flow into a bypass duct; 
         FIG. 7  is a schematic sectional view of an alternative example louvre assembly in position for extracting air flow from a bypass duct; 
         FIG. 8  is a schematic plan view of an example louvre assembly with an actuation mechanism; 
         FIG. 9  is a schematic sectional view of an example actuation mechanism for a louvre assembly; 
         FIG. 10  is a schematic sectional view of an alternative example actuation mechanism for a louvre assembly; and 
         FIG. 11  is a schematic sectional view of an example actuation mechanism for a louvre assembly for extracting air flow from a bypass duct. 
     
    
    
     DETAILED DESCRIPTION OF THE DISCLOSURE 
       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 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  18 ,  20  meaning that the flow through the bypass duct  22  has its own nozzle  18  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. 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 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  is a schematic plan view of an example louvre assembly  400  for a gas turbine engine, for example for a gas turbine engine bleed system. The louvre assembly  400  comprises an air discharge opening  401 , which surrounds first and second louvres  402 ,  406 . The first louvre  402  extends across a first portion  403  of the air discharge opening  401 , while the second louvre  406  extends across a second portion  407  of the air discharge opening  401 . Each louvre  402 ,  406  comprises a respective plurality of slats  404 ,  408 , which are pivotably mounted to rotate about respective first and second rotational axes between a closed position in which air is obstructed from flowing through the first and second portions  403 ,  407  of the air discharge opening  401  and an open position in which air is allowed to flow through the first and second portions  403 ,  407  of the air discharge opening  401 . The first rotational axes are mutually parallel to a first direction  405 . The second rotational axes are mutually parallel to a second direction  409 . 
     The first direction  405  is angled relative to the second direction  409  such that air, for example bleed air, exiting through the air discharge opening  401  with the first and second plurality of slats  404 ,  408  in the open position diverges away from a central axis  410  of the louvre assembly  400 , the central axis  410  extending between the first and second portions  403 ,  407  of the air discharge opening  401 . An angle  411  between the first and second directions  405 ,  409  may for example be between 10 and 30 degrees, and in particular embodiments may be between 12 and 24 degrees. Each direction  405 ,  409  may for example be aligned at an angle  412 ,  413  of between 75 and 85 degrees, optionally between 78 and 84 degrees, relative to the central axis  410 , i.e. at an angle α 1 , α 2  ( FIG. 5 ) of between 5 and 15 degrees, optionally between 6 and 12 degrees, to a plane orthogonal to the central axis  410 . 
     In the example shown in  FIG. 4 , the louvre assembly  400  is symmetrically arranged about the central axis  410 , i.e. the angles of the first and second directions  405 ,  409  to the central axis  410  are the same. This feature is advantageous when the louvre assembly  400  is positioned ahead of a symmetrically placed obstruction such as an engine core support pylon, as shown in further detail below. In some examples the angles may be selected to be different, depending on the prevailing direction of airflow in the bypass duct. 
       FIG. 5  illustrates the louvre assembly  400  in position within a bypass duct of a gas turbine engine between a series of outlet guide vanes  501  and a centrally positioned pylon  502  for supporting the engine core of the gas turbine engine. Angles α 1 , α 2  are chosen to closely match the direction of air flow from the outlet guide vanes  501  and so that air, for example bleed air, exiting the louvre assembly  400  is aligned to diverge either side of the pylon  502 , thereby reducing disturbance to the air flow in the bypass duct when bleed air is introduced. In this case, angles α 1 , α 2  are equal in magnitude and opposite in direction relative to the central axis  410  of the louvre assembly  400 , which coincides with the central axis of the pylon  502 . 
     The outlet guide vanes (OGVs) are designed to divert air flow away from pylons to reduce the pressure disturbance generated by pylons reaching to the fan blades upstream of the OGVs. Therefore, the fan OGVs tend to have a cyclic stagger and camber pattern, in which each OGV has a different exit flow angle depending on its position in the annulus and proximity to the pylon. The OGVs leave some residual swirl into the bypass duct and hence generate the flow exit angles α 1 , α 2  as shown in  FIG. 5 . To minimise disruption of air flow when introducing bleed air, these same angles can be used for the orientation of the slats in the louvre assembly. The angles will only be zero when the structural components such as pylons or struts are not present in the flow field. In practice, the exit flow angles are always present behind the stator vanes of any spool of the axial flow machine of all gas turbines. 
     A louvre assembly of the type described herein may also be used for air offtake locations between the intermediate compressor OGVs and intercase struts in the intermediate pressure compressor (IPC), air offtakes for IP bleeds and air offtakes on the pylon walls, used for other functions such as pre-coolers and heat-exchangers. 
     A section through the portion of the louvre assembly  400  marked A-A in  FIG. 5  is shown in  FIG. 6 , with the second louvre  406  in an open position to allow air to flow through the air discharge opening  401 . Each slat in the second louvre is oriented such that a plane of each slat is aligned at an angle β relative to a plane  601  of the louvre assembly  400 . The angle β may for example be within the range of 30 to 45 degrees relative to the plane  601 , the angle being a balance between maximising air flow through the opening  401  and minimising disruption of the air flow in the bypass duct  22 . Optimising this angle avoids flow blockage and pressure disturbances in the bypass duct. 
     In an alternative arrangement, illustrated in  FIG. 7 , a louvre assembly  700  may be positioned and arranged to extract air from, rather than introduce air into, a bypass duct of a gas turbine engine. The slats of the louvre assembly  700  are in this arrangement aligned against the direction of air flow  701  in the bypass duct, and may be aligned at an angle β of between 10 and 35 degrees to the plane of the louvre assembly  700 . The plane of the louvre assembly  700  may be parallel to the central axis  9  of the gas turbine engine (see for example  FIG. 1 ), or may be aligned at an angle to match an angle of a portion of the bypass duct in which the louvre assembly  700  is mounted. In the example of  FIG. 7 , the louvre assembly  700  is mounted downstream of the outlet guide vanes  501 , and may be connected to extract air from the bypass duct to a conduit  702  to provide cooling air for the turbine or other systems. 
       FIG. 8  illustrates schematically the louvre assembly  400  described above, comprising an actuation mechanism  801  that connects the first and second louvres  402 ,  406  for synchronous rotation of the first and second plurality of slats  404 ,  408  between the open and closed positions. The actuation mechanism comprises an actuator  802 , which may for example be a linear actuator, connected to an actuator rod  803  connecting each of the slats  404 ,  408  to the actuator  802 . When the actuator is operated to move the rod  803  in the direction  804  shown, each of the slats  404 ,  408  rotate about their respective axes (either along the first or second direction  405 ,  409 ) to move the slats between the open and closed positions. 
       FIGS. 9 and 10  illustrate more detailed schematic arrangements for actuating slats of a louvre assembly  900 ,  1000  with a linear actuator  802  and an actuator rod  803 , the louvre assembly  900 ,  1000  in each case being mounted in line with an inner annular surface  910  of a bypass duct. The actuator rod  803  in each case is attached to each of the slats  904  of the louvre assembly  900 ,  1000  with a pin joint  905 , and each slat or vane  904  is mounted to rotate about a pivot  906 . In the example shown in  FIG. 9 , actuation of the rod  803  causes each of the vanes  904  to actuate simultaneously by the same amount. In the example shown in  FIG. 10 , actuation of the rod  803  causes the slats  904  to actuate by differing amounts due to a gradual change in distance between the pin joint  905  and pivot  906  for each slat  904 , causing the slats nearer the actuator  802  to actuate by a larger amount than the slats further away from the actuator  802 . 
       FIG. 11  illustrates schematically an alternative arrangement for a louvre assembly  1100  mounted in line with an annular surface  1110  of a bypass duct, the louvre assembly  1100  in this case being arranged to extract air from the bypass duct. The constructional details of the louvre assembly  1100  are similar to that shown in  FIG. 9 . Air flow  1101  from the bypass duct is directed through the louvre assembly  1100  and into a conduit  1102  for providing an air flow to another system on the engine or elsewhere. The annular surface  1110  to which the louvre assembly  1100  is mounted may be parallel to the engine axis, as in  FIG. 11 , or may be aligned at an angle to the engine axis, corresponding to a conical surface annular surface of the bypass duct. 
     The specific arrangement of louvres shown herein may result in the louvres being placed exactly in front of the lower bifurcation, i.e. at the bottom dead centre of the engine. Conventionally, air bleeds are generally not placed in proximity to pylons because this can result in air flow disruption. However, the design of louvre assembly described herein allows the assembly to be placed in close proximity to pylons, and has the added advantage of allowing any water collected in the core compressor stages to be diverted into the bypass duct. 
     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.