GAS TURBINE ENGINE WITH IMPROVED VIGV SHIELDING

Abstract

A gas turbine engine includes: a fan rotating about an engine main axis; a core duct; an engine core; an Engine Section Stator (ESS) including a plurality of ESS vanes and arranged in the core duct downstream of the fan; and a plurality of variable inlet guide vanes (VIGV) adapted to rotate about a pivot axis and arranged in the core duct downstream of the ESS. The VIGV vanes are arranged angularly rotated with respect to the ESS vanes such that the VIGVs are shielded by the ESS, thereby protecting the VIGVs from icing and from ice shedding from the ESS vanes.

Claims

1. A gas turbine engine comprising: a fan rotating about an engine main axis and generating a core airflow and a bypass airflow; a core duct, across which the core airflow flows; an engine core comprising: a compressor for compressing the core airflow and comprising a plurality of stages, each stage comprising a row of rotor blades and a row of stator vanes, a first stage of said plurality of stages being arranged at an inlet of the compressor; combustion equipment; and a turbine connected to the compressor through a shaft; an Engine Section Stator (ESS) comprising a plurality of ESS vanes and arranged in the core duct downstream of the fan, each ESS vane comprising an ESS leading edge and an ESS trailing edge; and a plurality of variable inlet guide vanes (VIGVs) adapted to rotate about a pivot axis and arranged in the core duct downstream of the ESS and upstream of the compressor, each variable inlet guide vanes (VIGVs) comprising a VIGV leading edge and a VIGV trailing edge, and wherein the plurality of VIGVs are arranged angularly rotated with respect to the plurality of ESS vanes such that first longitudinal planes passing through respective 70% span ESS leading edge points are angularly rotated with respect to corresponding second longitudinal planes passing through respective 70% span VIGV pivot axis points by a rotation angle α comprised between 0.1° and 6°.

2. The gas turbine engine of claim 1, further comprising a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft.

3. The gas turbine engine of claim 2, wherein the gearbox has a reduction ratio comprised between 3 and 4.2

4. The gas turbine engine of claim 1, wherein the pivot axis extends along a radial direction.

5. The gas turbine engine of claim 1, wherein the fan has a diameter in a range from 220 cm to 380 cm

6. The gas turbine engine of claim 1, wherein the fan has a diameter in a range from 220 cm to 300 cm and is configured to have a rotational speed at cruise conditions in a range from 1700 rpm to 2500 rpm.

7. The gas turbine engine of claim 1, wherein the fan has a diameter in a range from 330 cm to 380 cm and is configured to have a rotational speed at cruise conditions in a range from 1200 rpm to 2000 rpm.

8. The gas turbine engine of claim 1, wherein a bypass ratio defined as a ratio of mass flow rate of the bypass airflow to mass flow rate of the core airflow at cruise conditions is in a range from 10 to 12.

9. The gas turbine engine of claim 1, wherein a bypass ratio defined as a ratio of mass flow rate of the bypass airflow to mass flow rate of the core airflow at cruise conditions is in a range from 13 to 15.

10. The gas turbine engine of claim 1, wherein the compressor is a first compressor, the turbine is a first turbine, and the shaft is a first shaft, the engine core further comprising a second compressor downstream of the first compressor, a second turbine upstream of the first turbine, and a second shaft connecting the second turbine with the second compressor.

11. A gas turbine engine comprising: a fan rotating about an engine main axis and generating a core airflow and a bypass airflow; a core duct, across which the core airflow flows; an engine core comprising: a compressor for compressing the core airflow and comprising a plurality of stages, each stage comprising a row of rotor blades and a row of stator vanes, a first stage of said plurality of stages being arranged at an inlet of the compressor; combustion equipment; and a turbine connected to the compressor through a shaft; an Engine Section Stator (ESS) comprising a plurality of ESS vanes and arranged in the core duct downstream of the fan, each ESS vane comprising an ESS leading edge and an ESS trailing edge; and a plurality of variable inlet guide vanes (VIGVs) adapted to rotate about a pivot axis and arranged in the core duct downstream of the ESS and upstream of the compressor, each variable inlet guide vanes (VIGVs) comprising a VIGV leading edge and a VIGV trailing edge, and wherein the plurality of VIGVs are arranged angularly rotated with respect to the plurality of ESS vanes such that first longitudinal planes passing through respective 70% span ESS leading edge points are angularly rotated with respect to corresponding second longitudinal planes passing through respective 70% span VIGV pivot axis points by a rotation angle α, and wherein the plurality of VIGVs are arranged angularly rotated with respect to the plurality of ESS vanes such that longitudinal planes passing through respective 90% span ESS leading edge points are angularly rotated with respect to corresponding longitudinal planes passing through respective 90% span VIGV pivot axis points by a rotation angle α1 less than the rotation angle α at 70% span and comprised between 0.05° and 5°.

12. The gas turbine engine of claim 11, wherein a mid-span ESS leading edge point is at an axial distance L from a mid-span VIGV leading edge point, the distance L being comprised between 80 mm and 650 mm.

13. The gas turbine engine of claim 11, further comprising a strut arranged in the core duct between the ESS and the plurality of VIGVs.

14. The gas turbine engine of claim 11, wherein the ESS comprises 40 to 80 ESS vanes

15. A gas turbine engine comprising: a fan rotating about an engine main axis and generating a core airflow and a bypass airflow; a core duct, across which the core airflow flows; an engine core comprising: a compressor for compressing the core airflow and comprising a plurality of stages, each stage comprising a row of rotor blades and a row of stator vanes, a first stage of said plurality of stages being arranged at an inlet of the compressor; combustion equipment; and a turbine connected to the compressor through a shaft; an Engine Section Stator (ESS) comprising a plurality of ESS vanes and arranged in the core duct downstream of the fan, each ESS vane comprising an ESS leading edge and an ESS trailing edge; and a plurality of variable inlet guide vanes (VIGVs) adapted to rotate about a pivot axis and arranged in the core duct downstream of the ESS and upstream of the compressor, each variable inlet guide vanes (VIGVs) comprising a VIGV leading edge and a VIGV trailing edge, and wherein the plurality of VIGVs are arranged angularly rotated with respect to the plurality of ESS vanes such that first longitudinal planes passing through respective 70% span ESS leading edge points are angularly rotated with respect to corresponding second longitudinal planes passing through respective 70% span VIGV pivot axis points by a rotation angle α, and wherein the plurality of VIGVs are arranged angularly rotated with respect to the plurality of ESS vanes such that longitudinal planes passing through respective 90% span ESS leading edge points are angularly rotated with respect to corresponding longitudinal planes passing through respective 90% span VIGV pivot axis points by a rotation angle α1 greater than the rotation angle α at 70% span and comprised between 0.02° and 6°.

16. The gas turbine engine of claim 15, further comprising a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft.

17. The gas turbine engine of claim 16, wherein the reduction gearbox has a gear ratio in a range from 3.1 to 3.8.

18. The gas turbine engine of claim 15, wherein the engine is configured to have a Turbine Entry Temperature (TET) at cruise conditions in a range from 1400 K and 1650 K.

19. The gas turbine engine of claim 15, wherein the engine is configured to have a Turbine Entry Temperature (TET) at maximum take-off condition in a range from 1800 K and 1950 K.

20. The gas turbine engine of claim 15, wherein a mid-span ESS leading edge point is arranged at a first radius (R1) from the engine main axis and a mid-span VIGV leading edge point is arranged at a second radius (R2) from the engine main axis, wherein a difference ΔR between the first radius and the second radius is between 60 mm and 280 mm.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0068] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0069] FIG. 1 is a sectional side view of a gas turbine engine;

[0070] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0071] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0072] FIG. 4 is a longitudinal section of a core duct;

[0073] FIG. 5 is a schematic front view of the core duct taken along arrows 6a-6a and 6b-6b of FIG. 4 to show ESS vanes and VIGVs with respective first and second longitudinal planes and corresponding rotation angle; and

[0074] FIG. 6 is a schematic isometric view of the core duct with an ESS vane and a corresponding VIGV, showing a first longitudinal plane passing through a 70% span VIGV pivot axis point and a corresponding second longitudinal plane passing through a respective 70% span ESS leading edge point.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0075] FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis, or engine main 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 some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

[0076] 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.

[0077] 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.

[0078] 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.

[0079] 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.

[0080] 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.

[0081] 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.

[0082] 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.

[0083] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0084] 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.

[0085] 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.

[0086] FIG. 4 illustrates an inlet duct 50 through which the core flow A enters the core 11 of the gas turbine engine 10. The inlet duct 50 is defined between an inner wall 64 and an outer wall 66 provided radially outwardly of the inner wall 64. An annular spacing between a radially outer surface 65 of the inner wall 64 and a radially inner surface 67 of the outer wall 66 defines the core duct which contains the core airflow A.

[0087] Arranged in the inlet duct 50, there are in flow series an Engine Section Stator (ESS) 52 with a plurality of vanes 54, a row of Variable Inlet Guide Vanes (VIGVs) 58, and a first rotor 60 of the low pressure compressor 14 with a plurality of first blades 62. Optionally, a strut 56 may be arranged in the inlet duct 50 between the ESS and the first rotor 60 of the low pressure compressor 14, for example between the ESS and the VIGVs 58. The strut 56 may be omitted in case the ESS is structural, i.e. if the ESS 52 is provided to support load between the inner wall 64 and the outer wall 66.

[0088] In general, the inner wall 64 and the outer wall 66 feature a curved profile and extend inwardly from the ESS 52 to the VIGVs 58 towards the engine main axis 9.

[0089] The ESS vanes 54 and the VIGVs direct the air entering the inlet duct 50 appropriately towards the compressor inlet, for example to improve the engine performance and avoid flow separation at the first blades 62.

[0090] The ESS vanes 54 are uniformly spaced circumferentially around the engine main axis 9 and have an aerofoil profile with an ESS leading edge 72 and an ESS trailing edge 74. The aerofoil profile extends in a chordwise direction from the leading edge to the trailing edge and in a spanwise direction from a first, radially inward end to a second, radially outward end. In this light, the ESS leading edge 72 extends from a first, radially inward, point 71 corresponding to 0% span to a second, radially outward, point 73 corresponding to 100% span. Point 71 may be referred to as 0% span ESS leading edge point 71 and point 73 as 100% span ESS leading edge point 73.

[0091] A mid-span ESS leading edge point PLE1 is defined on the ESS leading edge 72 at 50% span, midway between the inner wall 64 and the outer wall 66, or in other words midway between the 0% span ESS leading edge point 71 and the 100% span ESS leading edge point 73. The mid-span ESS leading edge point PLE1 is arranged at a first radius R1 from the engine main axis 9.

[0092] The VIGVs 58 are uniformly arranged circumferentially about the engine main axis 9 immediately upstream of the first rotor 60 of the low pressure compressor 14. The VIGVs 58 are rotatable about respective radial, or nearly radial, pivot axes 80, for example by means of a rotation mechanism per se known and therefore not illustrated in detail. In a typical arrangement each individual vane in a VIGV row is typically supported in two journal bearings at the radially inner and outer ends of the vane aerofoil section. The journal bearings permit the vane aerofoil to rotate or pivot about its spanwise axis. This axis is typically radial, or nearly radial, relative to the compressor or engine main axis 9.

[0093] The VIGVs 58 have an VIGV leading edge 76 and a VIGV trailing edge 78 and extend in a chordwise direction from the VIGV leading edge 76 to the VIGV trailing edge 78 and in a spanwise direction from a first, radially inward, end to a second, radially outward, end. In this light, the VIGV leading edge 76 extends from a first, radially inward point 75 corresponding to 0% span to a second, radially outward point 77 corresponding to 100% span. Point 75 may be referred to as 0% span VIGV leading edge point 75 and point 77 as 100% span VIGV leading edge point 77.

[0094] A mid-span VIGV leading edge point PLE2 is defined on the VIGV leading edge 76 at 50% span, midway between the inner wall 64 and the outer wall 66, or in other words midway between the 0% span VIGV leading edge point 75 and 100% span VIGV leading edge point 77. The mid-span VIGV leading edge point PLE2 is arranged at a second radius R2 from the engine main axis 9. The first radius R1 is generally greater than the second radius R2 and the difference ΔR between R1 and R2 is comprised between 60 mm and 280 mm, preferably between 150 mm and 260 mm. In an embodiment, the difference ΔR between R1 and R2 is for example 200 mm.

[0095] The mid-span ESS leading edge point PLE1 is axially distanced from the mid-span VIGV leading edge point PLE2 by an axial distance L. For example, the axial distance L is comprised between 300 mm and 650 mm, preferably between 450 mm and 650 mm. In an embodiment, the axial distance L is for example 500 mm. The ratio ΔR/L of the difference ΔR between the first radius R1 and the second radius R2 to the axial distance L is comprised between 0.23 and 0.70, preferably between 0.40 and 0.70. In an embodiment, the ratio ΔR/L is for example equal to 0.45.

[0096] The VIGVs 58 are arranged angularly rotated about the engine main axes 9 by a rotation angle α with respect to the ESS vanes 54, so that the VIGVs 58 are positioned in a shielded position with respect to the ESS vanes 54 to reduce water droplets contacting the VIGVs and therefore ice buildup. In particular, the VIGVs 58 are positioned such that a radially outer part of the VIGVs 58 is shielded. The radially outer part may be defined as the part of the VIGVs 58 between 40 and 100% span height.

[0097] The mutual arrangement of the ESS vanes 54 and VIGVs 58 will be further described with reference to FIGS. 5 and 6.

[0098] FIG. 5 is a schematic, simplified front view of the inlet duct 50 taken along arrows 6a-6a and 6b-6b of FIG. 4 to show ESS vanes 54 and corresponding VIGVs 58. In substance, FIG. 5 shows a first cross-section 6a-6a taken at a first transversal plane passing through points 90 defined at 70% span on the ESS leading edge 72 and a second cross-section 6b-6b taken at a second transversal plane passing through points 92 defined at 70% span on the pivot axes 80. Points 90 may be referred to as 70% span ESS leading edge points 90 and points 92 as 70% span VIGV pivot axis points 92.

[0099] For sake of simplicity, not all of the ESS vanes 54 and VIGVs 58 of the core duct 50 have been illustrated.

[0100] FIG. 5 further illustrates first longitudinal planes LP1 passing through respective 70% span ESS leading edge points 90 and second longitudinal planes LP2 passing through respective 70% span VIGV pivot axis points 92.

[0101] The first longitudinal plane LP1 and second longitudinal plane LP2 pass through the engine main axis 9.

[0102] Each first longitudinal plane LP1 is angularly rotated with respect to a corresponding second longitudinal plane LP2 by the rotation angle α.

[0103] For each first longitudinal plane LP1 the corresponding second longitudinal plane LP2 is defined as the adjacent second longitudinal plane LP2 in a clockwise direction as seen from the front of the engine, namely as shown in FIG. 5. Similarly, for each ESS vane 54 a corresponding VIGV 58 is defined as the adjacent VIGV 58 in a clockwise direction as seen from the front of the engine.

[0104] The ESS vanes 54 and VIGVs 58 are in a same number comprised between 40 and 80, for example 48, or 54, or 60, and, as both the ESS vanes 54 and VIGVs 58 are uniformly angularly arranged about the engine main axis 9, for each pair of ESS vane 58 and corresponding VIGV 58 the rotation angle α is the same.

[0105] It has to be noted that the VIGVs 58 are rotatable about the respective pivot axes 80 and therefore the VIGV leading edge 76 moves with respect to the ESS leading edge 72, whereas the pivot axis 80 does not. Consequently, the mutual position of the ESS leading edge 72 and the pivot axis 80, for any span height, does not change, even if the VIGV rotates. In other words, the rotation angle α between the first longitudinal plane LP1 and the second longitudinal plane LP2 does not depend on the VIGV rotation and does not vary with the VIGV rotation about the pivot axis 80.

[0106] Furthermore, it has to be noted that the rotation angle between corresponding longitudinal planes passing through points at different span heights (i.e. at span heights different from 70%) on the ESS leading edge 72 and on the VIGV pivot axis 80 may vary with span height because of the generally 3D shape of the ESS vanes, and/or the non-perfectly radially orientation of the pivot axis 80, but not with VIGV rotation.

[0107] FIG. 6 shows an isometric view of a detail of an embodiment of the core duct 50 with an ESS vane 54 and a corresponding VIGV 58, arranged between the inner wall 64 and the outer wall 66. The embodiment of FIG. 6 does not comprise the strut 56.

[0108] FIG. 6 further shows the first longitudinal plane LP1 passing through the 70% span ESS leading edge point 90 of the ESS vane 54 and the second longitudinal plane LP2 passing through the 70% span VIGV pivot axis point 92 of the VIGV 58. The second longitudinal plane LP2 is rotated by the rotation angle α from the first longitudinal plane LP1 about the engine main axis 9 in a clockwise direction when seen from the front of the engine. In embodiments the rotation angle α is comprised between 0.1° and 6°, for example 2°, or 2.5°, or 3°.

[0109] The first longitudinal plane LP1 intersects the inner wall 64 along a first line 94 and the outer wall 66 along a second line 96. Both first and second lines 94, 96 are shown as dotted lines in FIG. 6.

[0110] The second longitudinal plane LP2 intersects the inner wall 64 along a third line 98 and the outer wall 66 along a fourth line 100. Both the third and fourth lines 98, 100 are shown as dotted lines in FIG. 6.

[0111] 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 scope of the invention as described in the appended claims. 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.