Ice crystal protection for a gas turbine engine

Abstract

A gas turbine engine includes a fan mounted to rotate about a main longitudinal axis; an engine core, including a compressor, a combustor, and turbine coupled to the compressor through a shaft, and reduction gearbox; wherein the compressor includes a plurality of stages, each stage including a respective rotor and stator, a first stage of the plurality of stages being arranged at an inlet and including a first rotor with a plurality of blades; each blade extending chordwise from a leading edge to a trailing edge, and from root to tip for a span height, wherein 0% of the span height corresponds to the root and 100% of span height corresponds to tip; wherein a ratio of a leading edge radius of each of the plurality of first rotor blades at 0% span height to a minimum leading edge radius is comprised between 1 and 1.50.

Claims

1. A gas turbine engine comprising: a fan mounted to rotate about a main longitudinal axis; and an engine core comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; wherein the compressor comprises a plurality of stages, each stage comprising a respective rotor and stator, a first stage of said plurality of stages being arranged at an inlet and comprising a first rotor with a plurality of first rotor blades, each of the plurality of first rotor blades extends chordwise from a leading edge to a trailing edge, a ratio of a maximum leading edge radius of each of the plurality of first rotor blades to a fan diameter is comprised between 1.4×10.sup.−4 and 3.6×10.sup.−4, the maximum leading edge radius of each of the plurality of first rotor blades being the maximum radius that is defined by the leading edge in circumferential cross-section, and the fan diameter is greater than 220 cm, or greater than 230 cm, or greater than 240 cm.

2. The gas turbine engine according to claim 1, wherein the ratio of the maximum leading edge radius of each of the plurality of first rotor blades to the fan diameter is equal to or greater than 1.5×10.sup.−4, or equal to or greater than 1.6×10.sup.−4, or equal to or greater than 1.7×10.sup.−4, or equal to or greater than 2.1×10.sup.−4.

3. The gas turbine engine according to claim 1, wherein the fan diameter is less than 390 cm, or less than 370 cm.

4. The gas turbine engine according to claim 1, wherein the fan diameter is in a range from 330 cm to 380 cm.

5. The gas turbine engine according to claim 1, wherein the fan diameter is in a range from 240 cm to 280 cm.

6. The gas turbine engine according to claim 1, wherein the fan diameter is greater than 300 cm, and the ratio of the maximum leading edge radius to the fan diameter is greater than 1.7×10.sup.−4.

7. The gas turbine engine according to claim 1, wherein the maximum leading edge radius is comprised between 0.4 mm and 0.9 mm.

8. The gas turbine engine according to 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.

9. The gas turbine engine according to claim 1, wherein each of the plurality of first rotor blades has a circular or elliptical leading edge cross-section.

10. The gas turbine engine according to claim 1, wherein the fan comprises 16 to 24 fan blades.

11. The gas turbine engine according to claim 1, wherein the compressor comprises 2 to 8 stages.

12. The gas turbine engine according to claim 1, wherein the compressor is an intermediate pressure compressor, the gas turbine engine further comprising a high pressure compressor downstream of the intermediate pressure compressor, the turbine is an intermediate pressure turbine, the gas turbine engine further comprising a high pressure turbine upstream of the intermediate pressure turbine, and the shaft is a first shaft, the gas turbine engine further comprising a second shaft coupling the high pressure turbine to the high pressure compressor.

13. A gas turbine engine comprising: a fan mounted to rotate about a main longitudinal axis; an engine core comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; and 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, wherein the compressor comprises a plurality of stages, each stage comprising a respective rotor and stator, a first stage of said plurality of stages being arranged at an inlet and comprising a first rotor with a plurality of first rotor blades, each of the plurality of first rotor blades extends chordwise from a leading edge to a trailing edge, and from a root to a tip for a span height H, wherein 0% of the span height H corresponds to the root and 100% of the span height H corresponds to the tip, a leading edge radius of each of the plurality of first rotor blades varies between 60% and 100% of the span height H, and the leading edge radius of each of the plurality of first rotor blades between 60% and 100% of the span height H is at least twice as large as a minimum leading edge radius of each of the plurality of first rotor blades, and the leading edge radius is the radius that is defined by the leading edge in circumferential cross-section, and the minimum leading edge radius is the minimum radius that is defined by the leading edge in circumferential cross-section.

14. The gas turbine engine according to claim 13, wherein the compressor comprises three or four stages.

15. The gas turbine engine according to claim 13, wherein the reduction gearbox has a reduction ratio in a range from 3 to 4.2.

16. The gas turbine engine according to claim 13, wherein the leading edge radius is constant between 85% and 100% of the span height H.

17. The gas turbine engine according to claim 13, wherein the leading edge radius is constant between 0% and 50% of the span height H.

18. The gas turbine engine according to claim 13, wherein a maximum leading edge radius is comprised between 0.4 mm and 0.9 mm.

19. The gas turbine engine according to claim 13, wherein a ratio of a maximum leading edge radius of each of the plurality of first rotor blades to a fan diameter is comprised between 1.4×10.sup.−4 and 3.6×10.sup.−4, the maximum leading edge radius of each of the plurality of first rotor blades being the maximum radius that is defined by the leading edge in circumferential cross-section.

20. A gas turbine engine comprising: a fan mounted to rotate about a main longitudinal axis; and an engine core comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; wherein the compressor comprises a plurality of stages, each stage comprising a respective rotor and stator, a first stage of said plurality of stages being arranged at an inlet and comprising a first rotor with a plurality of first rotor blades, each of the plurality of first rotor blades extends chordwise from a leading edge to a trailing edge, a ratio of a maximum leading edge radius of each of the plurality of first rotor blades to a fan diameter is comprised between 1.4×10.sup.−4 and 3.6×10.sup.−4, the maximum leading edge radius of each of the plurality of first rotor blades being the maximum radius that is defined by the leading edge in circumferential cross-section, and the maximum leading edge radius is comprised between 0.4 mm and 0.9 mm.

Description

(1) Embodiments will now be described by way of example only, with reference to the Figures, in which:

(2) FIG. 1 is a sectional side view of a gas turbine engine;

(3) FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

(4) FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

(5) FIG. 4 shows a first rotor blade of a compressor:

(6) FIG. 5 is a partial schematic view, in cross-section, of the first rotor blade of FIG. 4 illustrating the difference between a maximum and a minimum leading edge radius of curvature;

(7) FIG. 6 is a partial schematic view, in cross-section, of the first rotor blade of FIG. 4 illustrating the difference between the maximum leading edge radius of curvature and the leading edge radius of curvature at 0% span height;

(8) FIG. 7 is a partial schematic view, in cross-section, of the first rotor blade of FIG. 4 illustrating the difference between the minimum leading edge radius of curvature and the leading edge radius of curvature at 0% span height.

(9) 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.

(10) The low pressure compressor 14 and the high pressure compressor 15 comprise respective pluralities of compressor stages, each stage comprising a rotor and a stator. FIG. 2 shows a first stage 42 and a second stage 44 of the low pressure compressor 14. The first stage 42 is arranged upstream of the second stage 44. The first stage 42 comprises a first rotor with a row of first rotor blades 50 and, downstream thereof, a first stator with a row of first stator vanes 52. Although the low pressure compressor 14 has been illustrated as comprising two stages, as noted elsewhere herein, the low pressure compressor 14 may comprise a different number of stages, for example two to eight stages.

(11) 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.

(12) 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.

(13) 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.

(14) 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.

(15) 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.

(16) 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.

(17) 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.

(18) 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.

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

(20) 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.

(21) 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.

(22) FIG. 4 illustrates an exemplary first rotor blade 50 comprising a root 54 and an aerofoil portion 56. The root 54 may have any suitable shape adapted to be mounted on a disc (not illustrated). The aerofoil portion 56 comprises a tip 58, opposite to the root 54, and a leading edge 60 and a trailing edge 62. The aerofoil portion 56 further comprises a pressure surface wall 64 extending from the leading edge 60 to the trailing edge 62 and a suction surface wall 66 extending from the leading edge 60 to the trailing edge 62.

(23) The aerofoil portion 56 extends along a spanwise direction S between the root 54 and the tip 58 for a span height H. and along a chordwise direction C between the leading edge 60 and the trailing edge 62.

(24) The first rotor blade 50 may have a span height H comprised between 140 mm and 220 mm and a true chord comprised between 80 mm and 160 mm.

(25) The leading edge 60 features a leading edge radius of curvature variable along the spanwise direction S. In other words, from the root 54, which may be considered at 0% of the span height H, to the tip 58, which may be considered at 100% of the span height H, the leading edge radius of curvature varies as it will now be described in more detail with reference to FIGS. 5-7.

(26) FIG. 5 shows two different transversal sections of the first rotor blade 50 showing the radius of curvature of the leading edge 60 at different span heights. In detail, FIG. 5 shows a first section S1 taken along line L1-L1 of FIG. 4 and containing the maximum leading edge radius of curvature R.sub.MAX and a second section S2 taken along line L2-L2 of FIG. 4 and containing the minimum leading edge radius of curvature R.sub.min.

(27) It is to be noted that the leading edge 60 may not lay on a single radial direction and the leading edge 60 at the first section S1 and at second section S2 may not be aligned along the same radial direction; thus the leading edge 60 at the first section S1 and at second section S2 are illustrated in FIG. 5 as coincident for sake of clarity only. In other words, the blade 50 may have any suitable shape and the leading edge 60 may extend along any suitable direction.

(28) The first section S1 is taken at a span height H.sub.L1 corresponding to 90% of the span height H. In other words, the maximum leading edge radius of curvature R.sub.MAX is arranged at 90% of the span height H. In other non-illustrated embodiments the maximum leading edge radius of curvature R.sub.MAX may be arranged at different span heights, for example in a range between 60% and 100%, or 80% and 100% of the span height H.

(29) The second section S2 is taken at a span height H.sub.L2 corresponding to 30% of the span height H. In other words, the minimum leading edge radius of curvature R.sub.min is arranged at 30% of the span height H. In other non-illustrated embodiments the minimum leading edge radius of curvature R.sub.min may be arranged at different span heights, for example in a range between 20% and 40% of the span height H.

(30) The ratio of the maximum leading edge radius of curvature R.sub.MAX to the minimum leading edge radius of curvature R.sub.min may be equal or greater than 2.2. Moreover, the ratio of the maximum leading edge radius of curvature R.sub.MAX to the minimum leading edge radius of curvature R.sub.min may be equal or less than 3.5. In an embodiment, the ratio of the maximum leading edge radius of curvature R.sub.MAX to the minimum leading edge radius of curvature R.sub.min may be 3.0.

(31) FIG. 6 shows the first section S1 of FIG. 5 containing the maximum leading edge radius of curvature R.sub.MAX and a third section S3 taken along line L3-L3 of FIG. 4 at a span height H.sub.L3 corresponding to 0% of the span height H. The leading edge radius of curvature at the span height H.sub.L3 is R.sub.0%.

(32) As in FIG. 5, the maximum leading edge radius of curvature R.sub.MAX and the leading edge radius of curvature R.sub.0% may not necessarily be aligned along one radial direction and are illustrated as coincident for sake of clarity only.

(33) The ratio of the maximum leading edge radius of curvature R.sub.MAX to the leading edge radius of curvature R.sub.0% may be equal or greater than 1.7. Moreover, the ratio of maximum leading edge radius of curvature R.sub.MAX to the minimum leading edge radius of curvature R.sub.min may be less than 3.2. In an embodiment, the ratio of maximum leading edge radius of curvature R.sub.MAX to the minimum leading edge radius of curvature R.sub.min may be 2.4.

(34) FIG. 7 shows the second section S2 containing the minimum leading edge radius of curvature R.sub.min, and the third section S3 containing the leading edge radius of curvature R.sub.0% at a span height of 0%.

(35) The ratio of the leading edge radius of curvature R.sub.0% to the minimum leading edge radius of curvature R.sub.min may be equal or greater than 1.1. Moreover, the ratio of the leading edge radius of curvature R.sub.0% to the minimum leading edge radius of curvature R.sub.min may be less than 1.4. In an embodiment, the ratio of the leading edge radius of curvature R.sub.0% to the minimum leading edge radius of curvature R.sub.min may be 1.25.

(36) As illustrated, the leading edge radius of curvature varies along the span, decreasing from a maximum value at 80%-100% span height to a minimum value at 20%-40% span height, and then increasing again from the minimum value to the 0% span height value.

(37) In non-illustrated embodiments the leading edge radius of curvature R.sub.0% may be equal to the minimum leading edge radius of curvature R.sub.min, or in other words their ratio may be equal to 1. For example, the leading edge radius of curvature may be constant and equal to the minimum leading edge radius of curvature R.sub.min between 0% and 50% of the span height H.

(38) 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.