Core duct assembly

Abstract

A core duct assembly for a gas turbine engine includes a core duct including an outer and an inner wall, the outer wall having an interior surface; a gas flow path member extending across the gas flow path at least partly between the inner and outer walls, the rotor blade having a radial span extending from a blade platform to a blade tip, wherein an upstream wall axis is defined as an axis tangential to a point on a first portion of the interior surface of the outer wall of the core duct extending downstream from the gas flow path member, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing the rotational axis of the engine, and wherein the upstream wall axis intersects the rotor blade at a point spaced radially inward from the blade tip of the rotor blade.

Claims

1. A core duct assembly for a gas turbine engine, the core duct assembly comprising: a core duct comprising an outer wall and an inner wall, the outer wall having a straight portion and a convex portion to form an interior surface of the outer wall, the core duct defining a gas flow path; a gas flow path member extending across the gas flow path at least partly between the inner wall and outer wall, the gas flow path member having a trailing edge, wherein the straight portion of the outer wall extends from the trailing edge of the gas flow path member to the convex portion of the outer wall; and at least one rotor blade located immediately downstream of the gas flow path member within the gas flow path and comprising a leading edge, the at least one rotor blade having a radial span extending from a blade platform of the at least one rotor blade to a blade tip of the leading edge of the at least one rotor blade, wherein an upstream wall axis is defined as an axis tangential to the interior surface of the outer wall at a point on the straight portion of the outer wall, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing a rotational axis of the gas turbine engine, a downstream wall axis is defined as an axis tangential to the interior surface of the outer wall at a point on the outer wall at an axial location of the blade tip of the leading edge of the at least one rotor blade, the downstream wall axis lying in the longitudinal plane of the gas turbine engine, the upstream wall axis intersects the at least one rotor blade at an intersection point spaced radially inward from the blade tip of the leading edge of the at least one rotor blade, and intersects the downstream wall axis at a location inside the gas flow path, an intersection distance is defined as a radial distance with respect to the rotational axis of the gas turbine engine between the intersection point and the blade tip of the leading edge of the at least one rotor blade, the intersection distance is in a range between 10% and 50% of the radial span of the at least one rotor blade.

2. The core duct assembly according to claim 1, wherein the intersection distance is in a range between 20 mm and 80 mm.

3. The core duct assembly according to claim 1, wherein the convex portion of the outer wall extends along the core duct between an upstream boundary and a downstream boundary.

4. The core duct assembly according to claim 3, wherein the straight portion of the outer wall extends from the trailing edge of the gas flow path member at an intersection of the trailing edge of the gas flow path member and the interior surface of the outer wall to the upstream boundary of the convex portion of the outer wall.

5. The core duct assembly according to claim 4, wherein an acceleration distance is defined as a distance along the interior surface of the outer wall between the intersection of the trailing edge of the gas flow path member and the interior surface and a point on the upstream boundary of the convex portion of the outer wall; and the convex portion of the outer wall has a centre point midway between the upstream boundary of the convex portion of the outer wall and the downstream boundary of the convex portion of the outer wall, and a separation distance is defined as an axial distance between the centre point of the convex portion of the outer wall and the point on the outer wall at an axial location of the blade tip of the leading edge of the at least one rotor blade.

6. The core duct assembly according to claim 5, wherein a ratio defined as: $\frac{\mathrm{the}\mathrm{acceleration}\mathrm{distance}}{\mathrm{the}\mathrm{separation}\mathrm{distance}}$ is in a range between 0.13 and 2.

7. The core duct assembly according to claim 1, wherein a trajectory angle is defined as an angle extending between the upstream wall axis and the downstream wall axis.

8. The core duct assembly according to claim 5, wherein a trajectory angle is defined as an angle extending between the upstream wall axis and the downstream wall axis, and a ratio defined as: $\frac{\mathrm{the}\mathrm{acceleration}\mathrm{distance}}{\mathrm{the}\mathrm{trajectory}\mathrm{angle}}$ is in a range between 0.25 mm/degree and 3.33 mm/degree.

9. The core duct assembly according to claim 5, wherein a trajectory angle is defined as an angle extending between the upstream wall axis and the downstream wall axis, and a ratio defined as: $\frac{\mathrm{the}\mathrm{separation}\mathrm{distance}}{\mathrm{the}\mathrm{trajectory}\mathrm{angle}}$ is in a range between 0.63 mm/degree and 5 mm/degree.

10. The core duct assembly according to claim 5, wherein a trajectory angle is defined as an angle extending between the upstream wall axis and the downstream wall axis, and a ratio defined as: $\frac{\left(\mathrm{acceleration}\mathrm{distance}/\mathrm{the}\mathrm{separation}\mathrm{distance}\right)}{\mathrm{the}\mathrm{trajectory}\mathrm{angle}}$ is in a range between 0.0033 degree.sup.−1 and 0.13 degree.sup.−1.

11. The core duct assembly according to claim 5, wherein the acceleration distance is at least 10 mm.

12. The core duct assembly according to claim 5, wherein the separation distance is at least 25 mm.

13. The core duct assembly according to claim 7, wherein the trajectory angle is in a range between 15 degrees and 40 degrees.

14. 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; wherein the engine core comprises the core duct assembly of claim 1, wherein the at least one rotor blade is at least one compressor rotor blade.

15. The gas turbine engine according to claim 14, further 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.

16. The gas turbine engine according to claim 14, wherein the turbine is a low pressure turbine, the compressor is a low pressure compressor, and the core shaft connects the low pressure turbine to the low pressure compressor, and the at least one rotor blade is provided in an upstream most stage of the low pressure compressor.

17. A core duct assembly for a gas turbine engine, the core duct assembly comprising: a core duct comprising an outer wall and an inner wall, the outer wall having a straight portion and a convex portion to form an interior surface of the outer wall, the core duct defining a gas flow path; a static vane extending across the gas flow path at least partly between the inner wall and outer wall, the static vane having a trailing edge, wherein the straight portion of the outer wall extends from the trailing edge of the static vane to the convex portion of the outer wall; and at least one rotor blade located downstream of the static vane within the gas flow path and comprising a leading edge, the at least one rotor blade having a radial span extending from a blade platform of the at least one rotor blade to a blade tip of the leading edge of the at least one rotor blade, wherein an upstream wall axis is defined as an axis tangential to the interior surface of the outer wall at a point on the straight portion of the outer wall, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing a rotational axis of the gas turbine engine, the upstream wall axis intersects the at least one rotor blade at an intersection point spaced radially inward from the blade tip of the leading edge of the at least one rotor blade, an intersection distance is defined as a radial distance with respect to the rotational axis of the gas turbine engine between the intersection point and the blade tip of the leading edge of the at least one rotor blade, the intersection distance is in a range between 10% and 50% of the radial span of the at least one rotor blade, and the at least one rotor blade is provided in an upstream most stage of a low pressure compressor of the gas turbine engine.

18. The core duct assembly according to claim 17, wherein the at least one rotor blade is disposed immediately downstream of the static vane.

19. A core duct assembly for a gas turbine engine, the core duct assembly comprising: a core duct comprising an outer wall and an inner wall, the outer wall having a straight portion and a convex portion to form an interior surface of the outer wall, the core duct defining a gas flow path; a static vane extending across the gas flow path at least partly between the inner wall and outer wall, the static vane having a trailing edge, wherein the straight portion of the outer wall extends from the trailing edge of the static vane to the convex portion of the outer wall; and at least one rotor blade is disposed immediately downstream of the static vane within the gas flow path and comprising a leading edge, the at least one rotor blade having a radial span extending from a blade platform of the at least one rotor blade to a blade tip of the leading edge of the at least one rotor blade, wherein an upstream wall axis is defined as an axis tangential to the interior surface of the outer wall at a point on the straight portion of the outer wall, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing a rotational axis of the gas turbine engine, the upstream wall axis intersects the at least one rotor blade at an intersection point spaced radially inward from the blade tip of the leading edge of the at least one rotor blade, an intersection distance is defined as a radial distance with respect to the rotational axis of the gas turbine engine between the intersection point and the blade tip of the leading edge of the at least one rotor blade, the intersection distance is in a range between 10% and 50% of the radial span of the at least one rotor blade.

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 sectional side view of a core duct assembly of the gas turbine engine according to an embodiment; and

(6) FIG. 5 shows a sectional side view of a core duct assembly according to another embodiment.

(7) 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.

(8) 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.

(9) 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.

(10) 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.

(11) 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.

(12) 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.

(13) 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.

(14) 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.

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

(16) 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.

(17) 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.

(18) FIG. 4 illustrates a core duct assembly 43 that forms part of the engine core 11 of the gas turbine engine 10. The core duct assembly 43 comprises a core duct 44 that defines a gas flow path through the engine core 11. The gas flow path runs through the compressors, combustion equipment and turbines described above and so carries the core airflow A through the gas turbine engine.

(19) The core duct 44 comprises an inner wall 46 and an outer wall 48. Between the inner and outer walls 46, 48 a generally annular gas flow path is defined. The inner and outer walls 46, 48 have respective interior surfaces 50, 52 arranged to confine gas within the gas flow path.

(20) The core duct assembly 43 further comprises a gas flow path member extending within the gas flow path. In the present embodiment, the gas flow path member is a vane 54 (e.g. a static or variable vane). The gas flow path member may however be any structure extending at least partly across the gas flow path, including any type of vane, support strut or similar structure. The gas flow path member extends across the gas flow path at least partly between the interior surfaces 50, 52 of the core duct 44. In the described embodiment, the vane 54 extends all of the way across the gas flow path. In other embodiments, it may extend only part of the way across the core duct 44. In that case, it may extend from the interior surface 52 of the outer wall 48. The vane 54 has a leading edge 54a and a trailing edge 54b relative to the direction of gas flow through the core duct 44.

(21) The core duct assembly 43 further comprises at least one rotor blade 56. The rotor blade may be one of a plurality of rotor blades provided in the low pressure compressor 14 or high-pressure compressor 15 described above. In the present embodiment the rotor blade 56 is provided in the low pressure compressor 14 (i.e. the lowest pressure compressor, which may be termed the intermediate pressure compressor) and located downstream of the vane 54. The rotor blade 56 is one of an array of rotor blades driven by the shaft 24 provided in the engine core as described above. The rotor blade 56 extends within the gas flow path so as to provide compression of the core airflow A. In other embodiments, the rotor blade may be a rotor blade of any of the other compressors provided in the engine. The rotor blade may, for example, be provided as part of the high pressure compressor 15.

(22) The low pressure compressor 14 comprises any number of stages, for example multiple stages. Each stage comprises a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes are axially offset from each other. In the presently described embodiment, the rotor blade 56 is provided in the first stage, i.e. in the most upstream stage, of the low pressure compressor 14, and the vane 54 is an inlet guide vane, or a strut, provided upstream of the rotor blade 56 in the core duct 44. The vane 54 may be a variable or static inlet guide vane.

(23) In yet other embodiments, the rotor blade 56 may be provided in the high pressure compressor, rather than the lower pressure compressor. The high pressure compressor 15 comprises any number of stages, for example multiple stages. Each stage comprises a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes are axially offset from each other. The rotor blade 56 may be provided in the first stage, i.e. in the most upstream stage, of the high pressure compressor 15, and the vane 54 is a vane, or a strut, provided upstream of the rotor blade 56 in the core duct 44. The vane 54 may be a variable or static inlet guide vane, for example a vane of the last stage, i.e. the most downstream stage, of the low pressure compressor 14, or an additional vane provided downstream of the low pressure compressor 14 in the core duct 44.

(24) The features described in combination with the low pressure compressor may therefore apply equally to the high pressure compressor, or any other compressor provided in the engine.

(25) The rotor blade has a leading edge 56a and a trailing edge 56b relative to the direction of gas flow through the core duct 44. The rotor blade 56 extends in a radial direction between a blade root or platform 56c and a radially outer blade tip 56d. In the present embodiment, the rotor blade 56 is an unshrouded blade and the blade tip 56d is spaced from the interior surface 52 of the core duct outer wall 48. In other embodiments, the rotor blade may be shrouded. In such an embodiment, the blade tip is defined as the radially outer extent of the aerofoil or gas-washed surface formed by the rotor blade (i.e. the point wherein the leading edge 56a intersects the shroud).

(26) Referring again to FIG. 4, an upstream wall axis 102 is defined as an axis tangential to a point on a first portion 62 of the interior surface 52 of the outer wall 48 of the core duct 44. The first portion is a portion extending downstream from the vane 54, and is upstream of the rotor blade 56. The upstream wall axis 102 lies in a longitudinal plane of the gas turbine engine 10 containing the rotational axis 9 of the engine (i.e. the longitudinal plane is the cross section plane of FIG. 2). The upstream wall axis 102 intersects the rotor blade 56 at a point spaced radially inward from the blade tip of the rotor blade (e.g. distance ‘c’ labelled in FIG. 4).

(27) During use of the gas turbine engine ice may be accreted on the vane 54 (or any other structure forming the gas flow path member) at a position at or near to the core duct outer wall 48. When this ice is shed, the gas flow through the core duct 44 will accelerate the ice rearwards along the core airflow path A. Where the shape of the interior surface 52 of the outer wall is straight (e.g. conical) or concave, the ice will be accelerated along the interior surface whist remaining close to or in contact with it. By shaping the interior surface of the bypass duct so that the upstream wall axis intersects the rotor blade away from its tip a turning point (e.g. a convex profile) in the shape of the interior surface may be formed. This may create a separation point at which the ice that has been shed at the vane 54 becomes separated from the interior surface 52 as it travels down the core duct 44.

(28) Once the ice has been separated, despite the gas stream vector remaining parallel to the direction of the core duct 44, the momentum gained by the piece of ice carries it away from the interior surface 52 of the outer wall 48 in an almost straight trajectory. After the piece of ice has been separated from the interior surface, the change in acceleration vector will cause some curvature to the flight path of the ice, as shown by the dotted line labelled 104 in FIG. 4, which shows an approximate path of the ice. By following the trajectory shown, the ice will impact the compressor rotor blade 56 at a point closer to the blade platform 56c (in other words, further from the blade tip 56d). At this point the rotor blade 56 is inherently more robust and so damage caused by the impact may be reduced.

(29) An intersection distance, labelled as ‘c’ in FIG. 4, may be defined as the radial distance between: i) a point radially level with the intersection of the upstream wall axis 102 and the leading edge 56a of the compressor rotor blade 56; and ii) the radial tip of the leading edge 56a of the compressor rotor blade 56.

(30) In order to provide a desired level of shielding of the compressor blade tip 56d the intersection distance (c) may be at least 10% of the radial span of the compressor rotor blade (and less than 100%). The radial span is defined as the radial distance between the radial tip of the leading edge 56a and the blade platform 56c. A minimum value of 10% may direct ice away from the structurally weaker part of the rotor blade. More specifically, the intersection distance may be in a range between 10% and 50% of the radial span of the compressor rotor blade. This range may provide a suitable level of damage mitigation without significantly affecting the geometry of the engine and/or the compressor rotor blade. For example, above a 50% span position the rotor blade is more likely to have sufficient structural strength to withstand ice impact, whereas achieving a value above 50% is likely to require significant changes in engine and/or compressor rotor blade geometry and may result in reduced performance. In other embodiments, the intersection distance may be in a range between 10% and 40% of the radial span of the compressor blade.

(31) In one embodiment, the value of the intersection distance may be in an inclusive range between 20 mm and 200 mm. In this embodiment, the total span of the rotor blade may be about 200 mm. More specifically, the intersection distance may be in an inclusive range between 20 mm and 80 mm.

(32) In other various embodiments, the intersection distance may be a proportion of the rotor blade span of any of the following: 10%, 15%, 20%, 25%, 30%, 35%, 40%, 45%, 50%, 55%, 60%, 65%, 70%, 75% or 80%. The intersection distance may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(33) In other various embodiments, intersection distance may be any of the following: 20 mm, 25 mm, 30 mm, 35 mm, 40 mm, 45 mm, 50 mm, 55 mm, 60 mm, 65 mm, 70 mm, 75 mm, or 80 mm. The intersection distance may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(34) The interior surface 52 of the core duct outer wall 48 extending between the vane 54 and the compressor rotor blade 56 further comprises a second portion 60. The second portion may be a downstream portion that is located downstream of the first, or upstream portion. The second portion 60 is formed by a region of the interior surface 52 which has a convex profile, and so may be termed a convex portion. In the present embodiment, the convex profile is a curvature that is convex in the direction of gas flow through the core duct 44 shown in FIG. 4 (e.g. a direction having a component parallel to the rotational axis of the engine). In other words, a line segment connecting two points on the convex profile lies outside the core duct 44.

(35) The convex portion 60 extends along the gas flow path from an upstream boundary 60a to a downstream boundary 60b. The convex portion 60 further has a centre point midway between the upstream boundary 60a and the downstream boundary 60b.

(36) The first portion of the core duct outer wall 48 forms a non-convex portion 62 extending between an upstream boundary 62a and a downstream boundary 62b. The non-convex portion extends between the intersection of the trailing edge 54b of the vane 54 and the interior surface 52; and the upstream boundary 60a of the convex portion 60. The upstream boundary of the convex portion 60 and the downstream boundary of the non-convex portion 62 therefore coincide. The non-convex portion may extend in a generally straight path angled toward the engine centreline in downstream direction along the core duct 44 as shown in FIG. 4. It may therefore form a conical shape. In other embodiments, the non-convex portion may have a generally concave curved shape in a direction along the core duct 44, i.e. a line segment connecting two points on the non-convex profile lies inside the core duct 44. In other words, the non-convex portion may be straight or have a concave curved shape in the plane of FIG. 4.

(37) The upstream wall axis 102 may be defined as an axis tangential to the interior surface 52 of the core duct outer wall 48 at a point on the upstream boundary 60a of the convex portion 60. The upstream wall axis may however be defined at any other point on the first portion 62 of the interior surface 52.

(38) The trajectory of the ice 104 once it leaves the interior surface 52 of the core duct outer wall 48 may depend on one or more of the following parameters:

(39) i) an acceleration distance (a) along which ice that has been shed can accelerate before separating from the interior surface of the bypass duct;

(40) ii) a separation distance (b) between the centre of the second portion 60 and the leading edge 56a of the rotor blade 56; and

(41) iii) a trajectory angle (ϕ) between the upstream wall axis 102 and the interior surface 52 of the core duct at the compressor rotor blade.

(42) As will be understood by the skilled person, it is desirable to increase the values of the above parameters in order to provide a sufficient intersection distance to mitigate damage to the compressor rotor blade. There are however constraints on the values of the parameters that would otherwise have a negative impact on the engine geometry (e.g. would impact other engine components and the desire for compact core geometry).

(43) Referring again to FIG. 4, the acceleration distance (a) is defined as: i) the distance along the interior surface 52 of the core outer wall 48 between the intersection of the trailing edge 54b of the vane 54 and the interior surface 52; and ii) a point on the upstream boundary 60a of the second portion 60 of the interior wall surface. The acceleration distance is measured in a longitudinal plane of the engine that contains the engine centreline (e.g. the plane of the cross section shown in FIG. 4).

(44) By increasing the acceleration distance, the amount of distance over which ice shed from the vane 54 can be accelerated by gas flow within the core duct is increased. By increasing the speed of the ice as it separates from the interior surface of the core duct the momentum of the ice may be increased, and the curvature of the flight path in a direction away from the centreline of the engine may be reduced (e.g. reducing the curve of trajectory 104 so that it is closer to the path of the upstream axis 102).

(45) In the present embodiment, the acceleration distance (a) is at least 10 mm. This has been found to provide a suitable level of acceleration of ice to reduce the risk of damage to the rotor blade 56 caused by ice impact.

(46) More specifically, the acceleration distance (a) may be in an inclusive range between 10 mm and 50 mm. This may provide a suitable level of ice acceleration, without requiring excessively large values of the acceleration distance, separation distance and the trajectory angle and so still allow compact engine geometry.

(47) Yet more specifically, the acceleration distance (a) may be in an inclusive range between 20 mm and 40 mm. This may provide a particularly advantageous balance of ice acceleration, without impacting the engine geometry.

(48) In various other embodiments, the acceleration distance may be any of the following: 10 mm, 15 mm, 20 mm, 25 mm, 30 mm, 35 mm, 40 mm, 45 mm or 50 mm. The acceleration distance may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(49) The separation distance (b) is defined as: i) the axial distance between the centre point of the second (i.e. convex) portion 60 and an axial position level with the radial tip of the leading edge 56a of the compressor rotor blade 56. In some embodiments, the region of the core duct outer wall 48 downstream of the convex portion 60 may not be parallel (or approximately parallel) with the engine centre line 9 as illustrated in FIG. 4. In other embodiments, the core duct outer wall 48 may extend towards or away from the engine centre line 9 in a downstream direction following the convex portion 60. Hence the radial distance from the engine centre line 9 to the blade tip 56d may be less or greater than the radial distance from the engine centre line to the centre point of the convex portion 60 from which the separation distance is measured. The separation distance is however measured in an axial direction parallel to the engine centreline in all geometries.

(50) As can be seen in FIG. 4, increasing the separation distance allows the ice to travel further along its trajectory away from the interior surface 52 before it impacts the compressor rotor blade 56. By increasing the separation distance the trajectory of the ice may result in an impact suitably remote from the rotor blade tip 56d. This may therefore help reduce damage in the case of an ice impact.

(51) In the present embodiment, the separation distance (b) is at least 25 mm. This has been found to provide a suitably long trajectory away from the interior surface of the core duct to mitigate rotor blade damage.

(52) More specifically, the separation distance (b) may be in an inclusive range between 25 mm and 75 mm. This may provide suitable ice travel away from the duct wall, without requiring excessively large values of the acceleration distance, separation distance and the trajectory angle and so allowing compact engine geometry.

(53) Yet more specifically, the separation distance (b) may be in an inclusive range between 35 mm and 55 mm. This may provide a particularly advantageous balance of ice flight path without impacting the engine geometry.

(54) In various other embodiments, the separation distance may be any of the following: 25 mm, 30 mm, 35, mm, 40 mm, 45 mm, 50 mm, 55 mm, 60 mm, 65 mm, 70 mm, or 75 mm. The separation distance may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(55) The trajectory of the ice may be dependent on the relative size of the acceleration distance and the separation distance. For example, a reduction in the speed of the ice may be compensated for by increasing the distance over which it travels before reaching the rotor blade. A ratio may be defined as: the acceleration distance (a)/separation distance (b). In the presently described embodiment, this ratio is in an inclusive range between 0.13 and 2 (e.g. 2.00).

(56) In various other embodiments, the ratio of acceleration distance divided by separation distance (a/b) may be any of the following: 0.13, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1.0, 1.1, 1.2, 1.3, 1.4, 1.5, 1.6, 1.7, 1.8, 1.9 or 2.0. The ratio of acceleration distance divided by separation distance (a/b) may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(57) Referring again to FIG. 4, a downstream wall axis 106 is defined as an axis tangential to the interior surface 52 of the core duct outer wall 48 at a point on the interior surface 52 level with leading edge tip of the rotor blade 56. The downstream wall axis 106 lies in a longitudinal plane of the gas turbine engine containing the rotational axis 9.

(58) The trajectory angle (ϕ) is defined as the angle extending between the upstream wall axis 102 and the downstream axis 104 as shown in FIG. 4 (e.g. the smallest angle measured between them). By increasing the trajectory angle, the direction of travel away from the interior surface of the core duct may be increased. This may help to increase the size of the intersection distance, without requiring excessively large values of the acceleration distance and the separation distance.

(59) In the presently described embodiment, the trajectory angle is in an inclusive range between 15 degrees and 40 degrees.

(60) More specifically, the trajectory angle may be in an inclusive range between 20 degrees and 30 degrees.

(61) In various other embodiments, the trajectory angle may be any of the following: 15 degrees, 20 degrees, 25 degrees, 30 degrees, 35 degrees, or 40 degrees. The acceleration distance may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(62) The size of the intersection distance may depend on the relative size of the acceleration distance and the trajectory angle. In the presently described embodiment, a ratio defined as the acceleration distance (a)/the trajectory angle (ϕ) is in an inclusive range between 0.25 mm/degree and 3.33 mm/degree.

(63) The size of the intersection distance may similarly depend on the relative size of the separation distance and the trajectory angle. In the presently described embodiment, a ratio of the separation distance (b)/the trajectory angle (ϕ) is in an inclusive range between 0.63 mm/degree and 5 mm/degree (e.g. 5.00 mm/degree).

(64) By choosing a core duct interior wall geometry within these ranges a suitable level of damage mitigation may be achieved, without impacting on other engine performance factors.

(65) In various other embodiments, the ratio of acceleration distance divided by trajectory angle (a/ϕ) may be any of the following: 0.25 mm/degree, 0.50 mm/degree, 0.75 mm/degree, 1.00 mm/degree, 1.25 mm/degree, 1.50 mm/degree, 1.75 mm/degree, 2.00 mm/degree, 2.25 mm/degree, 2.50 mm/degree, 2.75 mm/degree, 3.00 mm/degree, 3.25 mm/degree or 3.33 mm/degree. The ratio of acceleration distance divided by trajectory angle (a/ϕ) may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(66) In various other embodiments, the ratio of separation distance divided by trajectory angle (b/ϕ) may be any of the following: 0.63 mm/degree, 0.75 mm/degree, 2.00 mm/degree, 2.25 mm/degree, 2.50 mm/degree, 2.75 mm/degree, 3.00 mm/degree, 3.25 mm/degree, 3.50 mm/degree, 3.75 mm/degree, 4.00 mm/degree, 4.25 mm/degree, 4.50 mm/degree 4.75 mm/degree or 5.00 mm/degree. The ratio of separation distance divided by trajectory angle (b/ϕ) may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(67) As will be understood by the skilled person, the impact point of ice on the rotor blade may be controlled by the choice of the ratio of the acceleration distance, separation distance and trajectory angle. A ratio (referred to as the combined ratio) of these parameters may be defined as:

(68) $\frac{\mathrm{the}\mathrm{acceleration}\mathrm{distance}\left(a\right)/\mathrm{the}\mathrm{separation}\mathrm{distance}\left(b\right)}{\mathrm{trajectory}\mathrm{angle}\left(\varphi \right)}$

(69) In the presently described embodiment, the combined ratio is in an inclusive range between 0.0033 degree.sup.−1 and 0.13 degree.sup.−1.

(70) In various other embodiments, the combined ratio may be any of the following: 0.0033 degree.sup.−1, 0.010 degree.sup.−1, 0.020 degree.sup.−1, 0.030 degree.sup.−1, 0.040 degree.sup.−1, 0.050 degree.sup.−1, 0.060 degree.sup.−1, 0.070 degree.sup.−1, 0.080 degree.sup.−1, 0.090 degree.sup.−1, 0.10 degree.sup.−1 or 0.13 degree.sup.−1. The combined ratio may, for example, be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

(71) In one embodiment, the acceleration distance (a) may be 10 mm, the separation distance (b) may be 43 mm, the trajectory angle (ϕ) may be 25 degrees and the intersection distance (c) may be 20 mm. This may allow the acceleration distance to be minimised, while still providing a suitable value of the intersection distance.

(72) In another embodiment, the acceleration distance (a) may be 40 mm, the separation distance (b) may be 24 mm, the trajectory angle (ϕ) may be 40 degrees and the intersection distance (c) may be 20 mm. This may allow the separation distance to be minimised, while still providing a suitable value of the intersection distance.

(73) In yet another embodiment, the acceleration distance (a) may be 30 mm, the separation distance (b) may be 75 mm, the trajectory angle (ϕ) may be 15 degrees and the intersection distance (c) may be 20 mm. This may allow the trajectory angle to be minimised, while still providing a suitable value of the intersection distance.

(74) In the embodiment shown in FIG. 5, the first (i.e. non-convex) portion 62 of the duct outer wall 48 comprises an upstream portion 64 and a downstream portion 66. In this embodiment, the upstream wall axis 102 is a first upstream wall axis defined at a point on the downstream portion 66 of the outer wall. A second upstream wall axis 108 is defined as an axis tangential to a point on the upstream portion 64. The second upstream wall axis 108 lies in a longitudinal plane of the gas turbine engine 10 containing the rotational axis 9 of the engine similarly to the first upstream wall axis. As can be seen in FIG. 5, the first upstream wall axis 102 extends at an angle 102a relative to an axis 110 parallel to the rotational axis 9 of the gas turbine engine. The second upstream wall axis 108 similarly extends at an angle 108a relative to the axis 110 parallel to the rotational axis 9. The first upstream wall axis 102 extends at a greater angle relative to an axis parallel to the rotational axis 9 of the gas turbine engine compared to that of the second upstream axis 108.

(75) The downstream portion 66 therefore forms a ramp portion of the interior surface 52 of the core duct 44 to provide a greater deflection of ice away from the interior surface of the wall. This may help to move the impact point on the rotor blade further from the blade tip and so reduce this risk of damage cause by ice impact.

(76) The shape of the interior surface of the outer wall shown in FIGS. 4 and 5 are to be understood as examples only. Other shapes may be provided that also allow the upstream wall axis 102 to intersect the rotor blade at a point spaced radially inward from the blade tip of the rotor blade. In some embodiment, for example, the convex portion may have an angular profile, rather than being curved in the direction of gas flow through the core duct. In the presently described embodiment, a single combination of a convex and non-convex portion is shown between a respective vane 54 and rotor blade 56. In other embodiments, additional arrangement of convex and non-convex portion may be provided between other respective pairs of a vane and a relatively downstream rotor blade. These additional convex and non-convex portions may be provided in the same compressor, or a different compressor (e.g. a convex portion may be provided in both the low and high pressure compressors).

(77) 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.