LOUVRE SYSTEM

Abstract

A louvre assembly (400) for a gas turbine engine bleed system comprises: an air discharge opening (401); a first louvre (402) comprising a first plurality of slats (404) each pivotably mounted to rotate about a first direction (405); and a second louvre (406) comprising a second plurality of slats (408) each pivotably mounted to rotate about a second direction (409), wherein the first and second directions (405, 409) are angled relative to each other such that bleed air exiting through the air discharge opening (401) diverges away from a central axis (410) of the louvre assembly (400).

Claims

1. 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 air 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 air 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.

2. The louvre assembly of claim 1 wherein an angle between the first direction and second direction is between 10 and 30 degrees.

3. The louvre assembly of claim 1 wherein an angle between the first direction and the central axis is between 75 and 85 degrees.

4. The louvre assembly of claim 1 wherein an angle between the second direction and the central axis is between 75 and 85 degrees.

5. The louvre assembly of claim 1 wherein the first and second louvres are symmetrically arranged about the central axis.

6. The louvre assembly of claim 1 comprising an actuation mechanism connecting the first and second louvres for synchronous rotation of the first and second plurality of slats between the open and closed positions.

7. The louvre assembly of claim 6, further comprising an actuator connected to an actuator rod configured to cause rotation of the first and second pluralities of slats between the open and closed positions.

8. The louvre assembly of claim 1 wherein the first and second pluralities of slats are 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.

9. The louvre assembly of claim 8 wherein the first and second pluralities of slats are 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.

10. 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 comprising: an air discharge opening; a first louvre extending across a first portion of the air 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 air 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 according to claim 1, 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.

11. The gas turbine engine of claim 10, further comprising a plurality of fan outlet guide vanes; and a pylon, arranged in the bypass duct downstream of the fan outlet guide vanes, wherein the air discharge opening is arranged downstream of the fan outlet guide vanes and upstream of the pylon.

12. The gas turbine engine of claim 11, wherein the central axis of the louvre assembly is aligned with a central axis of the pylon.

13. The gas turbine engine of claim 10, comprising 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

14. The gas turbine engine according to claim 13, 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.

15. 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 comprising: an air discharge opening; a first louvre extending across a first portion of the air 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 air 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, 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.

16. The gas turbine engine of claim 15 wherein the first and second pluralities of slats are arranged to be aligned in the open position at an angle of between 10 and 35 degrees relative to a plane extending across the air discharge opening.

17. The gas turbine engine of claim 15, wherein the conduit provides cooling air for the turbine.

18. The gas turbine engine of claim 15, comprising 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.

19. The gas turbine engine according to claim 18, 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.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0067] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0068] FIG. 1 is a sectional side view of a gas turbine engine;

[0069] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0070] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0071] FIG. 4 is a schematic plan view of an example louvre assembly;

[0072] 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;

[0073] FIG. 6 is a schematic sectional view of an example louvre assembly in position for providing a bleed air flow into a bypass duct;

[0074] FIG. 7 is a schematic sectional view of an alternative example louvre assembly in position for extracting air flow from a bypass duct;

[0075] FIG. 8 is a schematic plan view of an example louvre assembly with an actuation mechanism;

[0076] FIG. 9 is a schematic sectional view of an example actuation mechanism for a louvre assembly;

[0077] FIG. 10 is a schematic sectional view of an alternative example actuation mechanism for a louvre assembly; and

[0078] 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

[0079] 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.

[0080] 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.

[0081] 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.

[0082] 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.

[0083] 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.

[0084] 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.

[0085] 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.

[0086] 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.

[0087] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0088] 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.

[0089] 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.

[0090] 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.

[0091] 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 1, 2 (FIG. 5) of between 5 and 15 degrees, optionally between 6 and 12 degrees, to a plane orthogonal to the central axis 410.

[0092] 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.

[0093] 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 1, 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 1, 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.

[0094] 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 1, 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.

[0095] 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.

[0096] 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 relative to a plane 601 of the louvre assembly 400. The angle 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.

[0097] 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 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.

[0098] 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.

[0099] 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.

[0100] 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.

[0101] 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.

[0102] 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.