GAS TURBINE ENGINE BLEED DUCT

Abstract

A single-piece gas turbine engine bleed duct for a gas turbine engine including: a main airflow conduit configured to transmit a bleed flow to a location outside the gas turbine engine; a pressure regulating valve for regulating airflow through the main airflow conduit; a first inlet duct for directing airflow to the main airflow conduit and toward the pressure regulating valve, the first inlet duct including a non-return valve; and a second inlet duct for directing airflow to the main airflow conduit and toward the pressure regulating valve, the second inlet duct including a control valve; wherein each of the first and second inlet ducts are directly connectable to an engine casing of a gas turbine engine.

Claims

1. 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.

2. The single-piece gas turbine engine bleed duct of claim 1, wherein at least one of the main airflow conduit, the first inlet duct and the second inlet duct comprises a valve for regulating airflow therethrough.

3. The single-piece gas turbine engine bleed duct of claim 2, wherein the main airflow conduit comprises a pressure regulating valve for regulating airflow through the main airflow conduit.

4. The single-piece gas turbine engine bleed duct of claim 2, wherein the first inlet duct comprises a non-return valve.

5. The single-piece gas turbine engine bleed duct of claim 2, wherein the second inlet duct comprises a control valve.

6. The single-piece gas turbine engine bleed duct of claim 5, wherein at least one of the control valve and the pressure-regulating valve is fueldraulically or hydraulically actuated.

7. The single-piece gas turbine engine bleed duct of claim 5, wherein at least one of the control valve and the pressure-regulating valve is electrically actuated via a motor.

8. 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 of claim 1.

9. The gas turbine engine of claim 8, wherein: 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 comprises 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.

10. The gas turbine engine of claim 8, further comprising a nacelle and a bypass duct defined by the nacelle and surrounding the engine core, and wherein at least part of the main airflow conduit of the single-piece gas turbine engine bleed duct is located within the bypass duct.

11. The gas turbine engine of claim 8, wherein the engine core comprises 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 being lower than the second pressure, and wherein the first inlet duct is connected to the first air plenum chamber and the second inlet duct is connected to the second air plenum chamber.

12. The gas turbine engine of claim 8, further comprising a first actuator to actuate the pressure regulating valve and a second actuator to actuate the control valve.

13. The gas turbine engine of claim 12 wherein at least one of the first and second actuators are controllable from a location outside the gas turbine engine.

14. The gas turbine engine of claim 12, wherein the first actuator comprises a first servo motor connected to a first linkage, wherein the first linkage is to actuate the pressure regulating valve and wherein the second actuator comprises a second servo motor connected to a second linkage, wherein the second linkage is to actuate the control valve.

15. The gas turbine engine of claim 12, wherein each actuator is connected to a pressure source and is pneumatically controlled.

16. The gas turbine engine of claim 12, further comprising an actuator plate connected to the single-piece gas turbine engine bleed duct and wherein each actuator is connected to the actuator plate.

17. The gas turbine engine of claim 8, further comprising a casing for the engine core, wherein the first and second inlet ducts are located within the casing for the engine core.

18. The gas turbine engine of claim 17, wherein at least a portion of the main air conduit is located within the casing for the engine core.

19. The gas turbine engine of claim 8, wherein a single duct transmits bleed flow through a bypass duct, bypassing the engine core.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0047] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0048] FIG. 1 is a sectional side view of a gas turbine engine;

[0049] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0050] FIG. 3 is a sectional side view of a single-piece gas turbine engine bleed duct;

[0051] FIG. 4 is a sectional side view of a gas turbine engine comprising a single-piece gas turbine engine bleed duct; and

[0052] FIG. 5 is a sectional side view of an example actuating mechanism for a single-piece gas turbine engine bleed duct.

DETAILED DESCRIPTION

[0053] 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.

[0054] 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.

[0055] 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.

[0056] 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.

[0057] 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.

[0058] 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.

[0059] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0060] 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.

[0061] 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.

[0062] 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.

[0063] 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.

[0064] 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.

[0065] 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.

[0066] 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.

[0067] 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.

[0068] 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.

[0069] 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.

[0070] 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.

[0071] 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.

[0072] 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.

[0073] 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.

[0074] The first and second linkages 75, 76 are each depicted as a bellcrank.

[0075] 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.

[0076] 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.

[0077] 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.