Super-cooled ice impact protection for a gas turbine engine

Abstract

A gas turbine engine comprises 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; 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 first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

Claims

1. 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, 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, the reduction gearbox having a reduction ratio in a range of from 3 to 4.2, wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, wherein a maximum leading edge radius of the first blades is greater than 0.4 mm, the maximum leading edge radius being the maximum radius that is defined by the leading edge of the first blades in circumferential cross-section, wherein a tip maximum thickness of the first blades is greater than 2.7 mm.

2. The gas turbine engine according to claim 1, wherein the maximum leading edge radius of the first blades is less than 0.9 mm.

3. The gas turbine engine according to claim 1, wherein a minimum leading edge radius of the first blades is greater than 0.15 mm, the minimum leading edge radius being the minimum radius that is defined by the leading edge of the first blades in circumferential cross-section.

4. The gas turbine engine according to claim 3, wherein the minimum leading edge radius of the first blades is located in an area less than 50% of a span height.

5. The gas turbine engine according to claim 3, wherein the minimum leading edge radius of the first blades is less than 0.6 mm.

6. The gas turbine engine according to claim 1, wherein the tip maximum thickness of the first blades is less than 5 mm.

7. The gas turbine engine according to claim 1, wherein a tip maximum thickness of the second blades is between 1.2 mm and 2.25 mm.

8. The gas turbine engine according to claim 1, wherein the fan has a fan diameter that is between 240 cm and 380 cm.

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

10. The gas turbine engine according to claim 1, wherein the compressor comprises 3 to 4 stages.

11. 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.

12. A gas turbine engine comprising: a fan mounted to rotate about a main longitudinal axis at a rotational speed, wherein the gas turbine engine is configured so that the rotational speed is in a range of from 1200 and 2000 rpm at cruise conditions, the fan having a fan diameter in a range of 330 cm to 380 cm, an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft, 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 first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, wherein a maximum leading edge radius of the first blades is greater than 0.4 mm, the maximum leading edge radius being the maximum radius that is defined by the leading edge of the first blades in circumferential cross-section, and wherein a tip maximum thickness of the first blades is greater than 2.7 mm.

13. The gas turbine engine according to claim 12, wherein the reduction gearbox has a reduction ratio in a range of from 3.1 and 3.8.

14. The gas turbine engine according to claim 12, 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.

15. The gas turbine engine according to claim 12, wherein the maximum leading edge radius of the first blades is less than 0.9 mm.

16. The gas turbine engine according to claim 12, wherein the tip maximum thickness of the first blades is less than 5 mm.

17. 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, 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 first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, wherein a maximum leading edge radius of the first blades is greater than 0.4 mm, the maximum leading edge radius being the maximum radius that is defined by the leading edge of the first blades in circumferential cross-section, and wherein a minimum leading edge radius of the first blades is greater than 0.15 mm.

18. The gas turbine engine according to claim 17, wherein the maximum leading edge radius of the first blades is less than 0.9 mm.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(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 is an enlarged schematic view of a part of a compressor of the gas turbine engine;

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

(7) FIG. 6 is a partial schematic view, in cross-section, showing the difference between the maximum leading edge radius of a first rotor blade and a maximum leading edge radius of a second rotor blade.

(8) FIGS. 7a and 7b are top views of tips of a first rotor blade and a second rotor blade, respectively.

DETAILED DESCRIPTION OF THE DISCLOSURE

(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. 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.

(10) 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.

(11) 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.

(12) 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.

(13) 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.

(14) 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.

(15) 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.

(16) 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.

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

(18) 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.

(19) 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.

(20) FIG. 4 illustrates a forward part of the low pressure compressor 14 in further detail.

(21) The low pressure compressor 14 comprises a first stage ST1 with a first rotor R1 and a first stator S1, and a second stage ST2 with a second rotor R2 and a second stator S2. The low pressure compressor 14 may comprise other stages, not illustrated.

(22) Each rotor (R1, R2) and stator (S1, S2) comprises a plurality of blades (B1, B2) and vanes (V1, V2), respectively.

(23) In detail, the first rotor R1 and the second rotor R2 comprise a row of first blades B1 and second blades B2, respectively; whereas the first stator S1 and the second stator S2 comprise a row of first vanes V1 and second vanes V2, respectively.

(24) The first blades B1 may have a span comprised between 140 mm and 220 mm, and a true chord comprised between 80 mm and 160 mm.

(25) The second blades B2 may have a span comprised between 120 mm and 180 mm, and a true chord comprised between 55 mm and 85 mm.

(26) Each blade B1, B2 and vane V1, V2 comprise a root (not illustrated) and an aerofoil portion with a leading edge, a trailing edge and a tip.

(27) The first blade B1 has a leading edge 50, a trailing edge 52, and a tip 54.

(28) The leading edge 50 has a leading edge radius variable along the span between a minimum leading edge radius R1min and a maximum leading edge radius R1max. In FIG. 5, there are schematically illustrated the minimum leading edge radius R1min and the maximum leading edge radius R1max by overlapping corresponding cross-sections 56, 58 of the first blades B1 taken along lines A-A and B-B of FIG. 4, respectively, in such a way that the leading edge 50 at those section coincide.

(29) The section 56 containing the minimum leading edge radius R1min, illustrated in dashed line in FIG. 5, may be at a span height between 20% and 30%, where 0% corresponds to the root and 100% corresponds to the tip.

(30) The section 58 containing the maximum leading edge radius R1max may be at a span height between 70% and 100%.

(31) The minimum leading edge radius R1min may be greater than 0.20 mm, for example equal to 0.25 mm.

(32) The maximum leading edge radius R1max may be greater than 0.4 mm, for example equal to 0.7 mm.

(33) The ratio of the maximum leading edge radius R1max of the first blade B1 to the minimum leading edge radius of the first blade B1 may be greater than 2.2, for example equal to 2.8.

(34) The second blade B2 has a leading edge 70, a trailing edge 72, and a tip 74. Analogously to the first blade B1, the leading edge 70 has a leading edge radius variable along the span between a minimum leading edge radius (not illustrated) and a maximum leading edge radius R2max, which is smaller than the maximum leading edge radius R1max of the first blade B1. The maximum leading edge radius R2max of the second blade B2 may be at a cross section 78, corresponding to a span height between 85% and 100%.

(35) In FIG. 6 there are illustrated the maximum leading edge radius R1max of the first blade B1 and the maximum leading edge radius R2max of the second blade B2 by superimposing section 58 and section 78 taken along lines C-C of FIG. 4, respectively, in such a way that the leading edge 50 of the first blade B1 and the leading edge 70 of the second blade B2 coincide. Section 78 containing the maximum leading edge radius R2max of the second blade B2 is illustrated in dashed line in FIG. 6.

(36) The maximum leading edge radius R2max of the second blade B2 may be comprised between 0.1 mm and 0.2 mm, for example equal to 0.16 mm.

(37) According to the disclosure, the ratio of the maximum leading edge radius R1max of the first blade B1 to the maximum leading edge radius R2max of the second blade B2 may be greater than 2.8. In an example, the maximum leading edge radius R1max of the first blade B1 may be equal to 0.7 mm and the maximum leading edge radius R2max of the second blade B2 may be equal to 0.16, such that the ratio of the maximum leading edge radius R1max of the first blade B1 to the maximum leading edge radius R2max of the second blade B2 may be equal to about 4.4.

(38) In FIGS. 7a and 7b there are illustrated top views of the tips 54, 74 of first blade B1 and the second blade B2, respectively.

(39) The tip 54 of the first blade B1 features a maximum thickness T1max that may be greater than 2.7 mm, for example equal to 4.3 mm. The maximum thickness T1max may be arranged at a chordwise position between 48% and 54%, for example between 50% and 52%, or about 51%, where 0% corresponds to the leading edge 50 and 100% corresponds to the trailing edge 52.

(40) The tip 74 of the second blade B2 features a maximum thickness T2max that may be greater than 1.2 mm and less than 2.25 mm, for example equal to 1.7 mm. The maximum thickness T2max may be arranged at a chordwise position between 42% and 62%, for example between 48% and 54%, or between 50% and 52%, or about 51%, where 0% correspond to the leading edge 70 and 100% corresponds to the trailing edge 72.

(41) In an example, the tip maximum thickness T2max of the second blade B2 is equal to 1.7 mm, and the tip maximum thickness T1max of the first blade B1 is equal to 4.3 mm, such that their ratio is equal to about 0.40.

(42) In another example, the tip maximum thickness T2max of the second blade B2 is equal to 1.3 mm, and the tip maximum thickness T1max of the first blade B1 is equal to 3.0 mm, such that their ratio is equal to about 0.43.

(43) 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.