Patent Publication Number: US-2020290741-A1

Title: Gas turbine engine bleed duct

Description:
The present disclosure relates to a gas turbine engine bleed duct, and more particularly to a single-piece gas turbine engine bleed duct for a gas turbine engine. 
     BACKGROUND 
     Typically, a portion of compressed air in a gas turbine engine for an aircraft does not enter the combustion equipment of the gas turbine engine, but is instead redirected as “bleed air” to other areas of the aircraft. Because the ambient air at high altitudes is too thin to contain enough oxygen to be fit for human inhalation, bleed air is sometimes cooled and directed to the aircraft&#39;s environmental control system where it is used to air-condition and pressurise the cabin. Due to its relatively high temperature, bleed air can also be used to de-ice parts of the aircraft. 
     To provide bleed air from the gas turbine engine, ducts may be used to redirect air from the compressor stage of the gas turbine engine to other areas of the aircraft. 
     STATEMENTS OF INVENTION 
     According to a first aspect there is provided a single-piece gas turbine engine bleed duct for a gas turbine engine comprising: a main airflow conduit configured to transmit a bleed flow to a location outside the gas turbine engine; a first inlet duct for directing airflow to the main airflow conduit; and a second inlet duct for directing airflow to the main airflow conduit; wherein each of the first and second inlet ducts are directly connectable to an engine casing of a gas turbine engine. 
     The single-piece gas turbine engine bleed duct may be for a gas turbine engine for an aircraft. The single-piece gas turbine engine bleed duct may provide bleed air to the aircraft. 
     At least one of the main airflow conduit, the first inlet duct and the second inlet duct may comprise a valve for regulating airflow therethrough. For example, the main airflow conduit may comprise a pressure regulating valve for regulating airflow through the main airflow conduit. Alternatively, or additionally, the first inlet duct may comprise a non-return valve. Alternatively, or additionally, the second inlet duct may comprise a control valve. 
     The pressure regulating valve may be a pressure regulating shut-off valve. The control valve may be a shut-off valve, e.g. a high pressure shut-off valve. The non-return valve may be a flapper valve. 
     At least one of the control valve and the pressure-regulating valve may be fueldraulically or hydraulically actuated. 
     At least one of the control valve and the pressure-regulating valve may be electrically actuated via a motor. 
     According to the invention there may be 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; and the single-piece gas turbine engine bleed duct as described above. 
     The turbine may be a first turbine, the compressor 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; and the second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft. 
     The first inlet duct may be connected to the first compressor and the second inlet duct may be connected to the second compressor. In an alternative example, each of the first and second inlet ducts may be connected to the first or second compressor, the first inlet duct being positioned upstream of the second inlet duct. In this way, the first inlet duct may receive air at a lower pressure than the second inlet duct. 
     The gas turbine engine may further comprise a nacelle and a bypass duct defined by the nacelle and surrounding the engine core, and the single-piece gas turbine engine bleed duct may be located within the bypass duct. At least a part of the main airflow conduit of the single-piece gas turbine engine bleed duct may be located within the bypass duct. At least one of the valves of the single-piece gas turbine engine bleed duct may be located within the bypass duct. In one example, the pressure regulating valve, the control valve, and the non-return valve are located within the bypass duct. 
     The engine core may comprise a first air plenum chamber containing air at a first pressure and a second air plenum chamber containing air at a second pressure, the first pressure may be lower than the second pressure, and the first inlet duct may be connected to the first air plenum chamber and the second inlet duct may be connected to the second air plenum chamber. The first air plenum chamber may be connected to, or may be part of, the first compressor and the second air plenum chamber may be connected to, or maybe part of, the second compressor. 
     The gas turbine engine may comprise a first actuator to actuate the pressure regulating valve and a second actuator to actuate the control valve. At least one of the first and second actuators may be controllable from a location outside the gas turbine engine. The first actuator may comprise a first servo motor connected to a first linkage, wherein the first linkage may be to actuate the pressure regulating valve and the second actuator may comprise a second servo motor connected to a second linkage, wherein the second linkage may be to actuate the control valve. In one example, at least one of the first and second linkages may comprise a bellcrank. 
     Each actuator may be connected to a pressure source and may be pneumatically controlled. 
     The gas turbine engine may further comprise an actuator plate connected to the single-piece gas turbine engine bleed duct and each actuator may be connected to the actuator plate. 
     The gas turbine engine may further comprise a casing for the engine core, and the first and second inlet ducts may be located within the casing for the engine core. 
     At least a portion of the main air conduit may be located within the casing for the engine core. 
     A single duct may transmit bleed flow through a bypass duct, bypassing the engine core. 
     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. 
     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. 
     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. 
    
    
     
       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 sectional side view of a single-piece gas turbine engine bleed duct; 
         FIG. 4  is a sectional side view of a gas turbine engine comprising a single-piece gas turbine engine bleed duct; and 
         FIG. 5  is a sectional side view of an example actuating mechanism for a single-piece gas turbine engine bleed duct. 
     
    
    
     DETAILED DESCRIPTION 
       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 optional 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. 
     It will be appreciated that the arrangement shown in  FIG. 2  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. 
     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. 
     Referring now to  FIG. 3  there is shown a single-piece gas turbine engine bleed duct  50  for the gas turbine engine  10 . The single-piece gas turbine engine bleed duct  50  comprises a main airflow conduit  52  configured to transmit a bleed air flow to a location outside the gas turbine engine  10 , and a pressure regulating valve  54  for regulating airflow through the main airflow conduit  52 . The single-piece gas turbine engine bleed duct  50  also comprises a first inlet duct  56  for directing airflow to the main airflow conduit  52  and toward the pressure regulating valve  54 , and a second inlet duct  58  for directing airflow to the main airflow conduit  52  and toward the pressure regulating valve  54 . The first inlet duct  56  comprises a non-return valve  57 . The second inlet duct  58  comprises a control valve  59 . Each of the first and second inlet ducts  56 ,  58  are directly connectable to an engine casing of the gas turbine engine  10 . 
     Each of the valves may be actuatable between a closed position, preventing air flow through their respective ducts, and an open position, permitting air flow through their respective ducts. As will be described below the pressure regulating valve and control valve may be actuated by an actuator between open and closed positions. The non-return valve may be a passively-actuated valve. 
     The pressure regulating valve  54  may be a pressure regulating shut-off valve to throttle air flow through the main airflow conduit  52  of the single-piece gas turbine engine bleed duct  50 . The control valve  59 , in one example, may be a high-pressure shut-off valve and the second inlet duct  58  may be connected to a high-pressure air chamber. The control valve  59  may be to meet the demands of the aircraft at low power. The non-return valve  57  may be to reduce the risk of re-ingestion of air back into part of the engine casing, or the engine core  11 , of the gas turbine engine  10 . Accordingly the control valve  59  may be an air-modulating high pressure control valve and the non-return valve  57  may be a passive non-return valve. The non-return valve  57  may be connected to a low-pressure air chamber. In one example the non-return valve  57  may be a flapper valve, for example a spring-balanced flap. In one example the pressure regulating valve  54  and the control valve  59  may be controlled by a part of an airframe using at least one of a pneumatic force or an electrical current. 
     As shown in  FIG. 3 , the first inlet duct  56  is connected to a first air plenum chamber  61  which contains air at a first pressure, and the second inlet duct  58  is connected to a second air plenum chamber  62  which contains air at a second pressure. The first pressure is lower than the second pressure and so the first air plenum chamber  61  may be regarded as a low-pressure plenum chamber and the second air plenum chamber  62  may be regarded as a high-pressure plenum chamber. 
     Referring now to  FIG. 4 , the gas turbine engine  10  comprises the single-piece gas turbine engine bleed valve  50 . As shown in  FIG. 4 , the first inlet duct  56  is connected to the low pressure compressor  14  and the second inlet duct  58  is connected to the high pressure compressor  15 . In other examples, the first and second inlet ducts  56 ,  58  may be connected to the same compressor but at different relative positions upstream/downstream of the airflow through the compressor. For example the first inlet duct  56  may be connected to a compressor upstream of the second inlet duct  58  so as to receive air at a lower pressure than the second inlet duct  58 . 
     The first air plenum chamber  61  may be part of the low pressure compressor  14  of the gas turbine engine  10  and the second air plenum chamber  62  may be part of the high pressure compressor  15  of the gas turbine engine  10 . 
     The single-piece gas turbine engine bleed valve  50  may be utilised to take air from the low pressure compressor  14 , via the first inlet duct  56 , and from the high pressure compressor  15 , via the second inlet duct  58 . Bleed air from the low pressure compressor  14  may be used when the engine is operating at high power and bleed air from the high pressure compressor  15  may be used when the engine is operating at low power. 
     Accordingly, during a “normal” or high power operation, the control valve  59  may be in a closed configuration and airflow may flow from a low-pressure air plenum chamber  61  and through the first inlet duct  56  where it is directed to flow through the main airflow conduit  52  and to another part of the aircraft. Flow through the main airflow conduit  52  is regulated by the pressure regulating valve  54 , During a low power operation, the control valve  59  may be in an open configuration to permit airflow from a high-pressure air plenum chamber  62  and through the second inlet duct  58  where it is directed to flow through the main airflow conduit  52  and to another part of the aircraft. Air flow back through the first inlet duct  57  is prevented by the non-return valve  57 . 
     As is also shown in  FIG. 4 , the single-piece gas turbine engine bleed  50  is located within the bypass duct  22  of the gas turbine engine. At least one of the pressure regulating valve, the control valve, and the non-return valve is therefore located in the bypass duct  22 . In another example the single-piece gas turbine engine bleed duct  50  is located within a casing  51  for the engine core  11 . At least one of the pressure regulating valve, the control valve, and the non-return valve may be within the casing  51 . In another example, at least a portion of the main air flow conduit  52  is located in the casing  51 . 
     Referring now to  FIG. 5  there is shown an example actuation mechanism for the single-piece gas turbine engine bleed duct  50 . Specifically there is shown a mechanism for actuating the pressure regulating valve (not shown) and the non-return valve (not shown). A first actuator  71  is provided to actuate the pressure regulating valve and a second actuator  72  is provided to actuate the control valve. The actuators may be located outside the gas turbine engine  10 . 
     The first actuator  71  comprises a first servo motor  73  connected to a first linkage  75 . The first linkage  75  is to actuate the pressure regulating valve. The first linkage  75  may be a multi-positional linkage, moveable between a range of positions, and movement of the linkage between two positions may cause the pressure regulating valve to transition between an open and a closed position, permitting or prevent airflow through the main inlet conduit  52 , respectively. 
     The second actuator  72  comprises a second servo motor  74  connected to a second linkage  76 . The second linkage  76  is to actuate the control valve. The second linkage  76  may be a multi-positional linkage, moveable between a range of positions, and movement of the linkage between two positions may cause the control valve to transition between an open and a closed position, permitting or prevent airflow through the second inlet duct  58 , respectively. 
     The first and second linkages  75 ,  76  are each depicted as a bellcrank. 
     Each of the first and second actuators  71 ,  72  is hydraulically, fueldraulically or pneumatically actuatable. The first actuator  71  is connected to a first pressure source line  77  and the second actuator is connected to a second pressure source line  78 . The pressure source lines  77 ,  78  may be configured to house and transmit liquid (e.g. fuel) or gas for actuating the first and second bellcranks  75 ,  76 . 
     The gas turbine engine  10  comprises an actuator plate  80 . Each of the first and second actuators  71 ,  72  are connected to the actuator plate  80 . For example they may be mounted on the actuator plate. The actuator plate may be connected to the single-piece gas turbine engine bleed duct  50 . 
     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.