Combustor blade and vane spacing for ice crystal protection for a gas turbine engine

Abstract

A gas turbine engine includes a fan, compressor, combustor, and turbine coupled to the compressor through a shaft. The compressor includes first, second, and third stages with respective rotor blades and stator vanes. The rotor blades and stator vanes are arranged such that a sum of mid-span axial gaps between the trailing edge of the first rotor blades and the leading edge of the first stator vanes, the trailing edge of the first stator vanes and the leading edge of the second rotor blades, the trailing edge of the second rotor blades and the leading edge of the second stator vanes, and the trailing edge of the second stator vanes and the leading edge of the third rotor blades is equal to $a\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}}.a75\mathrm{mm}\mathrm{and}\mathrm{the}\mathrm{inlet}\mathrm{flow}\mathrm{function}=\frac{m\sqrt{T}}{P}.$

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, and a reduction gearbox with an input from the shaft and an output to the fan, the fan being adapted to rotate at a lower rotational rate than the shaft, wherein the compressor comprises a first stage with a first rotor and a first stator, a second stage with a second rotor and a second stator, and a third stage with a third rotor and a third stator, each rotor and stator comprising a plurality of blades and vanes, respectively, with a leading edge, a trailing edge and a tip, and adjacent rotors and stators being separated by axial gaps, wherein between the first rotor and the third rotor, the rotors and stators are arranged such that a sum of mid-span axial gaps between the blade trailing edge of the first rotor and the vane leading edge of the first stator, the vane trailing edge of the first stator and the blade leading edge of the second rotor, the blade trailing edge of the second rotor and the vane leading edge of the second stator, and the vane trailing edge of the second stator and the blade leading edge of the third rotor is equal to $a\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein a is equal or greater than 75 mm and less than 115 mm, and the inlet flow function is defined as $\frac{m\sqrt{T}}{P},$ wherein m is a mass airflow at an inlet to the compressor in kg/s, T is the total temperature at the inlet to the compressor in K, and P is the total pressure at the inlet to the compressor in kPa at the compressor aerodynamic design point, and the axial gaps between the rotors and stators in each of the first stage, second stage, and third stage are provided such that the respective blades of each rotor are able to deflect upon ice crystal impact without touching the respective vanes of the adjacent stator.

2. The gas turbine engine according to claim 1, wherein a sum of tip axial gaps between the blade trailing edge of the first rotor and the vane leading edge of the first stator, the vane trailing edge of the first stator and the blade leading edge of the second rotor, the blade trailing edge of the second rotor and the vane leading edge of the second stator, and the vane trailing edge of the second stator and the blade leading edge of the third rotor is equal to $b\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein b is equal or greater than 85 mm and less than 130 mm, and the inlet flow function is defined as in claim 1.

3. The gas turbine engine according to claim 1, wherein a mid-span axial distance between the blade trailing edge of the first rotor and the vane leading edge of the first stator is equal to $c\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein c is equal or greater than 24 mm and less than 40 mm, and the inlet flow function is defined as in claim 1.

4. The gas turbine engine according to claim 1, wherein a tip axial distance between the blade trailing edge of the first rotor and the vane leading edge of the first stator is equal to $d\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein d is equal or greater than 23 mm and less than 40 mm, and the inlet flow function is defined as in claim 1.

5. The gas turbine engine according to claim 1, wherein a is equal or greater than 85 mm and less than 105 mm, and the inlet flow function is between 6 and 20 kg/s.Math.K.sup.0.5/kPa.

6. The gas turbine engine according to claim 1, wherein a tip axial distance between the blade trailing edge of the second rotor and the vane leading edge of the second stator is equal to $e\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein e is equal or greater than 24 mm and less than 38 mm, and the inlet flow function is defined as in claim 1.

7. The gas turbine engine according to claim 1, wherein a mid-span axial distance between the blade trailing edge of the third rotor and the vane leading edge of the third stator is equal to $f\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein f is equal or greater than 15 mm and less than 40 mm, and the inlet flow function is defined as in claim 1.

8. The gas turbine engine according to claim 1, wherein a tip axial distance between the blade trailing edge of the third rotor and the vane leading edge of the third stator is equal to $g\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein g is equal or greater than 17 mm and less than 45 mm, and the inlet flow function is defined as in claim 1.

9. The gas turbine engine according to claim 1, wherein a mid-span axial distance between the vane trailing edge of the first stator and the blade leading edge of the second rotor is equal to $h\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein h is equal or greater than 25 mm and less than 40 mm, and the inlet flow function is defined as in claim 1.

10. The gas turbine engine according to claim 1, wherein a tip axial distance between the vane trailing edge of the first stator and the blade leading edge of the second rotor is equal to $i\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein i is equal or greater than 25 mm and less than 40 mm, and the inlet flow function is defined as in claim 1.

11. The gas turbine engine according to claim 1, wherein a mid-span axial distance between the vane trailing edge of the second stator and the blade leading edge of the third rotor is equal to $l\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein l is equal or greater than 17 mm and less than 30 mm, and the inlet flow function is defined as in claim 1.

12. The gas turbine engine according to claim 1, wherein a tip axial distance between the vane trailing edge of the second stator and the blade leading edge of the third rotor is equal to $m\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein m is equal or greater than 24 mm and less than 30 mm, and the inlet flow function is defined as in claim 1.

13. The gas turbine engine according to claim 1, wherein the compressor comprises 3 to 8 stages.

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

15. The gas turbine engine according to claim 1, wherein the fan has a diameter greater than 230 cm and less than 420 cm, and the inlet flow function is comprised between 4 and 20 kg/s.Math.K.sup.0.5/kPa.

16. 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 with an input from the shaft and an output to the fan, the fan being adapted to rotate at a lower rotational rate than the shaft, wherein the compressor comprises a first stage with a first rotor and a first stator, comprising a plurality of blades and vanes, respectively, with a leading edge, a trailing edge and a tip, the blades and vanes being separated by axial gaps, wherein a mid-span axial distance between the blade trailing edge of the first rotor and the vane leading edge of the first stator is equal to $c\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein c is equal or greater than 24 mm and less than 35 mm, and the inlet flow function is defined as $\frac{m\sqrt{T}}{P},$ wherein m is a mass airflow at an inlet to the compressor in kg/s, T is the total temperature at the inlet to the compressor in K, and P is the total pressure at the inlet to the compressor in kPa at the compressor aerodynamic design point, and the axial gaps between the blades and vanes of the first stage are provided such that the blades of the first rotor are able to deflect upon ice crystal impact without touching the vanes of the adjacent first stator.

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, and a reduction gearbox with an input from the shaft and an output to the fan, the fan being adapted to rotate at a lower rotational rate than the shaft, wherein the compressor comprises a first stage with a first rotor and a first stator, comprising a plurality of blades and vanes, respectively, with a leading edge, a trailing edge and a tip, the blades and vanes being separated by axial gaps, wherein a tip axial distance between the blade trailing edge of the first rotor and the vane leading edge of the first stator is equal to $d\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$ wherein d is equal or greater than 23 mm and less than 40 mm, and the inlet flow function is defined as $\frac{m\sqrt{T}}{P},$ wherein m is a mass airflow at an inlet to the compressor in kg/s, T is the total temperature at the inlet to the compressor in K, and P is the total pressure at the inlet to the compressor in kPa at the compressor aerodynamic design point, and the axial gaps between the blades and vanes of the first stage are provided such that the blades of the first rotor are able to deflect upon ice crystal impact without touching the vanes of the adjacent first stator.

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 is an enlarged view of a part of a compressor of the gas turbine engine.

(6) 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.

(7) 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.

(8) 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.

(9) 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.

(10) 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.

(11) 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.

(12) 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.

(13) 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.

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

(15) 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.

(16) 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.

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

(18) The low pressure compressor 14 comprises a first stage ST1 with a first rotor R1 and a first stator S1, a second stage ST2 with a second rotor R2 and a second stator S2, and a third stage ST3 with a third rotor R3 and third stator S3. The low pressure compressor 14 may comprise other stages, not illustrated.

(19) Each rotor (R1, R2, R3) and stator (S1, S2, S3) comprises a plurality of blades (B1, B2, B3) and vanes (V1, V2, V3), respectively.

(20) In detail, the first rotor R1, the second rotor R2 and the third rotor R3 comprise a row of first blades B1, second blades B2 and third blades B3, respectively; whereas the first stator S1, the second stator S2 and the third stator S3 comprise a row of first vanes V1, second vanes V2, and third vanes V3, respectively.

(21) Each blade B1, B2, B3 and vane V1, V2, V3 comprise an aerofoil portion with a leading edge, a trailing edge and a tip.

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

(23) The second blades B2 may have a leading edge span comprised between 120 mm and 180 mm and a true chord comprised between 50 mm and 90 mm.

(24) The third blades B3 may have a leading edge span comprised between 80 mm and 140 mm and a true chord comprised between 35 mm and 70 mm.

(25) The first vanes V1 may have a leading edge span comprised between 120 mm and 180 mm and a true chord comprised between 30 mm and 60 mm.

(26) The second vanes V2 may have a leading edge span comprised between 90 mm and 150 mm and a true chord comprised between 20 mm and 50 mm.

(27) The third vanes V3 may have a leading edge span comprised between 70 mm and 120 mm and a true chord comprised between 15 mm and 45 mm. Blades and vanes are distanced by respective axial gaps, i.e. gaps along the principal rotational axis 9. The axial gaps may vary along the span of the blades.

(28) A first axial gap G1 is defined between the first blade B1 and the first vane V1, a second axial gap G2 is defined between the first vane V1 and the second blade B2, a third axial gap G3 is defined between the second blade B2 and the second vane V2, a forth axial gap G4 is defined between the second vane V2 and the third blade B3, and a fifth axial gap G5 is defined between the third blade B3 and the third vane V3.

(29) In detail, a first mid-span axial gap AGMS1 and a first tip axial gap AGT1 are defined between the trailing edge 52 of the first blade B1 and the leading edge 56 of the first vane V1, at the first blade B1 mid-span position, i.e. at a mid-span position of the first blade B1 aerofoil portion, and at a first blade B1 trailing edge tip 53, respectively.

(30) The first blade B1 trailing edge tip 53 is where the trailing edge 52 and the tip 54 of the first blade B1 meet.

(31) Corresponding second, third, fourth, and fifth mid-span axial gaps AGMS2, AGMS3, AGMS4, AGMS5 and second, third, fourth, and fifth tip axial gaps AGT2, AGT3, AGT4, AGT5 are defined between the trailing edge 58 of the first vane V1 and the leading edge 62 of the second blade B2, between the trailing edge 64 of the second blade B2 and the leading edge 62 of the second vane V2, between the trailing edge 70 of the third vane V3 and the leading edge 74 of the third blade B3, and between the trailing edge 76 of the third blade B3 and the leading edge 80 of the third vane V3.

(32) The second and third mid-span axial gaps AGMS2, AGMS3 are defined at the second blade B2 mid-span position 61, whereas the fourth and fifth mid-span axial gap AGMS4, AGMS5 is defined at the third blade B3 mid-span position 75.

(33) The second tip axial gap AGT2 is defined at a second blade B2 leading edge tip 63, i.e. where the leading edge 62 and the tip 66 of the second blade B2 meet; the third tip axial gap AGT3 is defined at a second blade B2 trailing edge tip 65, i.e. where the trailing edge 64 and the tip 66 of the second blade B2 meet; the fourth tip axial gap AGT4 is defined at a third blade B3 leading edge tip 73, i.e. where the leading edge 74 and the tip 78 of the third blade B3 meet; and the fifth tip axial gap AGT5 is defined at the third blade B3 trailing edge tip 77, i.e. where the trailing edge 76 and the tip 78 of the third blade B3 meet.

(34) It has to be noted that the distance of the blade mid-span positions and blade leading/trailing edge tips to the principal rotational axis 9 depends on the compressor geometry and may differ from blade to blade, as in the example illustrated in the figures.

(35) According to the disclosure, a sum of the first second, third and fourth mid-span axial gaps AGMS1, AGMS2, AGMS3, AGMS4 may be equal to

(36) $a\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$

(37) wherein a is equal or greater than 75 mm and the inlet flow function is defined as

(38) $\frac{m\sqrt{T}}{P},$

(39) wherein m is a mass airflow at an inlet to the compressor in kg/s, T is the total temperature at the inlet to the compressor in K, and P is the total pressure at the inlet to the compressor in kPa at the compressor aerodynamic design point. For most applications, the inlet flow function may be comprised between 4 and 20 kg/s.Math.K.sup.0.5/kPa, for example between 4 and 18 kg/s.Math.K.sup.0.5/kPa, or between 6 and 18 kg/s.Math.K.sup.0.5/kPa, or between 6 and 16 kg/s.Math.K.sup.0.5/kPa, or between 8 and 20 kg/s.Math.K.sup.0.5/kPa, or between 8 and 18 kg/s.Math.K.sup.0.5/kPa.

(40) For example, a may be equal or greater than 80 mm, or equal or greater than 85 mm, or equal or greater than 90 mm.

(41) a may be equal or less than 115 mm. For example, a may be equal or less than 110 mm, or equal or less than 105 mm, or equal or less than 100 mm, or equal or less than 95 mm. The present inventors have understood that when a is greater than 115 mm, the loss in efficiency exceeds an acceptable value and the benefit in ice crystal impact does not compensate for the loss in efficiency.

(42) Accordingly, a may be comprised between 75 mm and 115 mm, for example between 80 mm and 110 mm, or between 85 mm and 105 mm, or between 90 mm and 100 mm.

(43) A sum of the first, second, third and fourth tip axial gaps AGT1, AGT2, AGT3, AGT4 may be equal to

(44) $b\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$

(45) wherein b may be equal or greater than 85 mm, and the inlet flow function may be as previously defined.

(46) b may be equal or greater than 90 mm, for example equal or greater than 95 mm, or equal or greater than 100 mm, or equal or greater than 105 mm.

(47) b may be equal or less than 130 mm, for example equal or less than 125 mm, or equal or less than 120 mm, or equal or less than 115 mm, or equal or less than 110 mm. As previously b may be comprised between 85 mm and 130 mm, for example between 90 mm and 130 mm, or between 90 mm and 125 mm, or between 90 mm and 120 mm, or between 95 mm and 125 mm, or between 95 mm and 120 mm, or between 95 mm and 115 mm, or between 100 mm and 120 mm, or between 100 mm and 115 mm.

(48) In the illustrated embodiment, the first mid-span axial gap AGMS1 is greater than the second mid-span axial gap AGMS2, the third mid-span axial gap AGMS3, the fourth mid-span axial gap AGMS4, and the fifth mid-span axial gap AGMS5.

(49) In alternative, not illustrated embodiments, the first mid-span axial gap AGMS1 may be less than the second mid-span axial gap AGMS2, which in turn may be greater than the third mid-span axial gap AGMS3.

(50) In an embodiment, the first mid-span axial gap AGMS1 is equal to

(51) 0$c\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein c is equal to 25 mm; the second mid-span axial gap AGM2 is equal to

(52) $h\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein h is equal to 30 mm; the third mid-span axial gap AGMS3 is equal to

(53) $n\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein n is equal to 18 mm; the fourth mid-span axial gap AGMS4 is equal to

(54) $l\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein 1 is equal to 20 mm; and the fifth mid-span axial gap AGMS5 is equal to

(55) $f\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein f is equal to 15 mm. In such embodiment, therefore, a is equal to 93 mm.

(56) In an alternative embodiment, for the first, second, third, fourth, and fifth mid-span axial gap AGMS1, AGMS2, AGMS3, AGMS4, AGSM5 the parameters c, h, n, l, and f are equal to 27 mm, 28 mm, 21 mm, 20 mm, and 29 mm, respectively. In such alternative embodiment, a is equal to 96 mm

(57) In a further alternative embodiment, for the first, second, third, and fourth mid-span axial gap AGMS1, AGMS2, AGMS3, AGMS4 the parameters c, h, n, and l are equal to 20 mm, 20 mm, 18 mm, and 17 mm, respectively. In such further alternative embodiment, a is equal to 75 mm.

(58) Moreover, in the illustrated example, the first tip axial gap AGT1 is greater than the second tip axial gap AGT2, and the third tip axial gap AGT3 is greater than the fourth tip axial gap AGT4. In this embodiment, the blades at the tip have more room to deflect downstream than upstream.

(59) In an alternative, not illustrated embodiment, the second tip axial gap AGT2 is greater than the first and third tip axial gap AGT1, AGT3, the latter in turn being greater than the fourth tip axial gap AGT4.

(60) In an embodiment, the first tip axial gap AGT1 is equal to

(61) $d\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein d is equal to 27 mm; the second tip axial gap AGT2 is equal to

(62) $i\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein i is equal to 35 mm; the third tip axial gap AGT3 is equal to

(63) $e\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein e is equal to 25 mm; the fourth tip axial gap AGT4 is equal to

(64) $m\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein m is equal to 24 mm, and the fifth tip axial gap AGT5 is equal to

(65) $g\sqrt{\frac{\mathrm{inlet}\mathrm{flow}\mathrm{function}}{10}},$
wherein g is equal to 23 mm. In such embodiment, therefore, b is equal to 111 mm.

(66) In an alternative embodiment, for the first, second, third, fourth and fifth tip axial gap AGT1, AGT2, AGT3, AGT4, AGT5 the parameters d, i, e, m, and g are equal to 30 mm, 28 mm, 27 mm, 25 mm, and 41 mm respectively, and b is therefore equal to 110 mm.

(67) In a further alternative embodiment, for the first, second, third, and fourth tip axial gap AGT1, AGT2, AGT3, AGT4 the parameters d, i, e, and m, are equal to 22 mm, 24 mm, 20 mm, and 19 mm, respectively, and b is therefore equal to 85 mm.

(68) Although only the first three stages ST1, ST2, and ST3 have been illustrated in FIG. 4, the low pressure compressor 14 may comprise three or more than three stages, for example up to eight stages.

(69) Equally, the disclosure may be applied to the high pressure compressor 15, in particular to the first three stages of the high pressure compressor 15, which in turn may comprise 3 to 12 stages, for example up to 11 stages, or up to 10 stages.

(70) 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.