GAS TURBINE ENGINE FOR AN AIRCRAFT COMPRISING AN AIR INTAKE

Abstract

A gas turbine engine for an aircraft includes an engine core, fan, air intake, nacelle, and gearbox. The core includes a turbine, compressor, and core shaft connecting the turbine and compressor. The fan is upstream of the core and includes a plurality of fan blades, and has a diameter greater than 2.0 m. The air intake is upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60 and defines highlight, throat and diffuser areas. The nacelle at least partially surrounds the core and fan. The gearbox receives input from the core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft. The engine has a bypass ratio greater than 10. The nacelle has a length and the ratio of the length of the nacelle to the fan diameter is 0.4 to 2.5.

Claims

1. A gas turbine engine for an aircraft, the gas turbine engine 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, the fan comprising a plurality of fan blades and having a fan diameter greater than 2.0 m; an air intake located upstream of the fan, the air intake having a ratio of intake length to fan diameter of from 0.20 to 0.60 and defining a highlight area, a throat area and a diffuser area; a nacelle that at least partially surrounds the engine core and the fan; 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; and the gas turbine engine has a bypass ratio greater than 10; wherein the nacelle has a length and the ratio of the length of the nacelle to the fan diameter is from 0.4 to 2.5.

2. The gas turbine engine of claim 1, wherein the ratio of the length of the nacelle to the fan diameter is from 1.2 to 2.0.

3. The gas turbine engine of claim 1, wherein the ratio of the length of the nacelle to the fan diameter is from 0.9 to 2.0.

4. The gas turbine engine of claim 1, wherein the ratio of the length of the nacelle to the fan diameter is from 0.9 to 1.8.

5. The gas turbine engine of claim 1, wherein the ratio of the length of the nacelle to the fan diameter is from 0.5 to 1.2.

6. The gas turbine engine of claim 1, wherein the ratio of the highlight area to the throat area is from 1.15 to 1.35 and the ratio of the diffuser area to the throat area is from 0.85 to 1.15.

7. The gas turbine engine of claim 1, wherein the ratio of the throat area to fan face area of the fan is from 0.94 to 1.05.

8. The gas turbine engine of claim 1, wherein the fan diameter is greater than 2.2 m.

9. The gas turbine engine of claim 8, wherein the fan diameter is from 2.5 m to 4.5 m.

10. The gas turbine engine of claim 1, wherein the contraction ratio of the gas turbine engine is from 1.10 to 1.35.

11. The gas turbine engine of claim 1, wherein the local contraction ratio at bottom dead centre of the gas turbine engine is from 1.20 to 1.35.

12. The gas turbine engine of claim 1, wherein the local contraction ratio at the top dead centre of the gas turbine engine is from 1.15 to 1.35.

13. The gas turbine engine of claim 1, wherein the local contraction ratio at one or both lateral sides of the gas turbine engine is from 1.15 to 1.35.

14. The gas turbine engine of claim 1, wherein the ratio of the intake length to fan diameter is from 0.20 to 0.60.

15. The gas turbine engine of claim 14, wherein the ratio of the intake length to fan diameter is from 0.25 to 0.45.

16. The gas turbine engine of claim 1, wherein at cruise, the quasi-non-dimensional mass flow rate Q for the gas turbine engine is from 0.029 to 0.036 kgs.sup.1N.sup.1K.sup.1/2.

17. The gas turbine engine of claim 1, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0090] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0091] FIG. 1 is a sectional side view of a gas turbine engine;

[0092] FIG. 2 is a close-up sectional side view of an upstream portion of a gas turbine engine;

[0093] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0094] FIG. 4 is a schematic cross section view of an air intake for a gas turbine engine with the spinner omitted;

[0095] FIG. 5 is a schematic cross section view of an air intake for a gas turbine engine with the spinner included; and

[0096] FIG. 6 is a schematic cross section view of an air intake for a gas turbine engine with the spinner omitted.

[0097] The following table lists the reference numerals used in the drawings with the features to which they refer:

TABLE-US-00001 No. Feature FIG. A Core airflow 1 B Bypass airflow 1 9 Principal rotational axis 1, 2, 4, 5, 6 10 Gas turbine engine 1 11 Engine core 1 12 Air Intake 1 14 Low pressure compressor 1 15 High pressure compressor 1 16 Combustion equipment 1 17 High pressure turbine 1 18 Bypass exhaust nozzle 1 19 Low pressure turbine 1 20 Core exhaust nozzle 1 21 Nacelle 1 22 Bypass duct 1 23 Fan 1, 2 24 Stationary support structure 2 26 Shaft 1, 2 27 Shaft 1 28 Sun gear 2, 3 30 Epicyclic gearbox 1, 2, 3 32 Planet gear 2, 3 34 Planet carrier 2, 3 36 Linkage 2 38 Ring gear 2, 3 40 Linkage 2 100 Air intake 4, 5, 6 102 Inner wall of nacelle 4, 5, 6 104 Throat 4, 5, 6 106 Lip (or highlight) 4, 5, 6 107 Highlight area 4, 5, 6 108 Highlight radius 4 110 Throat radius 4 114 Throat area 4 116 Diffuser region 4 120 Scarf angle 4 122 Intake centreline axis 4, 6 124 Intake droop 4 126 Fan face 4, 5, 6 202 Downstream end of nacelle 1 204 Spinner 5 206 Mid-point of spinner 5 208 Spinner area at mid-point of spinner 5 210 Length of spinner 5 212 Diffuser area 5 220 Spinner area at the fan face 5 224 Intake length i.e. linear distance between 6 points C1 and C2 230 Length of nacelle 1 302 Local diffuser angle 6 304 Line extending between the apex of the 6 throat and the point where the fan face intersects the intake inner wall 402 Lip length 6 404 Lip height 6 408 Line perpendicular to the intake centreline 6 axis passing through the throat

DETAILED DESCRIPTION OF THE DISCLOSURE

[0098] Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

[0099] 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. For the purposes of this disclosure, the gas turbine comprises a nacelle. The nacelle 21 surrounds or at least partially 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.

[0100] 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 core exhaust 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.

[0101] 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.

[0102] 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.

[0103] 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.

[0104] 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.

[0105] 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.

[0106] 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.

[0107] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0108] 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 exhaust 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.

[0109] 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.

[0110] Referring to FIGS. 1, 4, 5 and 6 schematic cross sections of an air intake 100 of a gas turbine engine are shown. The cross sections extend through the engine centreline, the top dead centre and the bottom dead centre of the air intake. The air intake 100 is the part of the nacelle 21 upstream from the fan. The air intake 100 is shown bounded by the fan (roughly illustrated as line 126, which is the fan face of the fan 23) on the downstream side and by the inner walls 102 of the nacelle 21, which extend from the fan face 126 to the lip 106 of the air intake. The upstream end of the air intake 100 is open to allow airflow to enter in to engine via the air intake 100. The lip 106 is the upstream most portion of the air intake, also referred to as the highlight or leading edge. The lip 106 extends annularly around the air intake and defines the opening through which the airflow enters the air intake. The area of the opening defined by the annular lip is referred to as the highlight area 107. The highlight area 107 is the area through which air passes to enter the air intake, airflow outside of the highlight area passes around the exterior of the nacelle.

[0111] The air intake 100 has an intake centreline axis 122 which defines the centre of the air intake. The intake centreline axis may be coincident with the engine axis 9 or may be non-parallel with the engine centreline axis 9. The perpendicular distance from the intake centreline axis 122 to the lip defines the highlight radius 108. The highlight radius 108 is equivalent to % of the distance between diametrically opposed points on the lip 106. The highlight radius may not be constant around the circumference.

[0112] Moving in the downstream direction from the lip, the air intake narrows from the lip 106 to a minima, the minima defining the throat 104. The throat may be a circular or ellipsoid annulus around the interior of the air intake and defines a 2D area called the throat area 114. In some embodiments, the throat may not be purely circular or elliptical in shape and instead may vary irregularly in axial position around the circumference. In these embodiments, the throat area may be considered as the minimum area of the surface circumscribed by the throat. The distance from the centre of the throat area to the throat 104 is referred to as the throat radius 110. Where the throat 104 is not a circular annulus, the throat radius 110 is the shortest distance from the centroid of the throat area to the throat and may be expressed as an average value for measurements of throat radius around the circumference of the air intake 100.

[0113] Moving downstream from the throat 104, the air intake 100 widens towards the fan face 126 of the fan 23. This region is called the diffuser region (shown roughly as 116 in FIG. 4). As stated above, airflow enters the air intake 100 via the highlight area 107. The airflow speeds up as it flows through the constriction of the throat 104 before slowing down again as it passes into the diffuser region 116 before entering the fan. The fan may comprise a spinner 204 located within the diffuser region 116.

[0114] The contraction ratio is the ratio of the highlight area 107 to the throat area 114. The air intake may have a contraction ratio of from 1.10 to 1.35, or from 1.15 to 1.25, or from 1.23 to 1.35, or a range of any combination of the aforesaid end points.

[0115] The size of the throat 104 relative to the lip 106 may vary at different points around the circumference of the air intake 100, for example, the throat 104 may have a greater prominence from the lip 106 at the bottom dead centre of the air intake 100 than at the top dead centre or sides of the air intake. Differences in throat size around the air intake may be characterised by the local contraction ratio, which is the ratio of the highlight radius 108 to the throat radius 110 at an individual circumferential point on the air intake. For example, at the bottom dead centre, the local contraction ratio may be from 1.20 to 1.35, or from 1.20 to 1.25, or from 1.25 to 1.35 or a range of any combination of the aforesaid end points. The local contraction ratio at the top dead centre may be from 1.15 to 1.35, or from 1.15 to 1.25, or from 1.23 to 1.35 or a range of any combination of the aforesaid end points. The local contraction ratio at one or both sides of the air intake may be from 1.15 to 1.35, or from 1.15 to 1.25, or from 1.23 to 1.35, or any range formed from any combination of the aforesaid end points.

[0116] Wth reference to FIG. 4, the diffuser region 116 is the region between the throat 104 and the fan face 126. The air intake generally widens from the upstream to downstream in the diffuser region, and the spinner 204 (FIG. 5) typically occupies a part of the diffuser region around the downstream part of the engine centreline axis 9. The diffuser area 212 is a measure of the cross-sectional area of the air intake 100 in the diffuser region 116 where the air flows through. The diffuser area 212 is equivalent to the area perpendicular to the engine centreline axis 9, bounded by the inner walls 102 of the nacelle 21, measured coincident with the mid-point 206 of the length 210 of the spinner 204, minus the spinner area 208 at the mid-point 206. The length of the spinner 204 is measured parallel to the engine centreline axis 9, from the tip of the spinner to the base of the spinner.

[0117] The fan face area of an engine may be defined as the total area of a fan (i.e. the fan area) with the area of the spinner at the fan face (220) subtracted. E.g. Fan Face Area=/4 (D.sub.fancase.sup.2-D.sub.spinner.sup.2), where D.sub.fancase is the fan outer casing diameter at the same axial location as the fan blade tips leading edges, D.sub.spinner is the diameter of the spinner at the same axial location . In some embodiments, the fan face area may be from 2.8 m.sup.2 to 12 m.sup.2; or from 4.5 m.sup.2 to 10 m.sup.2; or from 6 m.sup.2 to 8 m.sup.2, or a range of any combination of the aforesaid end points.

[0118] In some embodiments, the ratio of the throat area to fan face area may be from 0.94 to 1.05, or from 1 to 1.05, or from 1.02 to 1.04, or between a range of any combination of the preceding endpoints.

[0119] An air intake 100 may also be partially characterised by the ratio of the intake length 224 to the fan diameter (L/D). The intake length, where the air intake has a non-zero scarf and non-zero droop is equivalent to the distance along the intake centreline axis from the plane defined by the highlight to the axial plane defined by the leading edges of the fan. Alternatively, where the air intake has no droop and zero scarf, the intake length is equivalent to the distance from the centre point of the highlight area to the centre of the fan face area, which may be measured parallel to the intake centreline axis. In embodiments, the ratio of intake length to fan diameter may be from 0.20 to 0.60, or from 0.20 to 0.50, or from 0.25 to 0.45, or from 0.30 to 0.40, or a range of any combination of the aforesaid end points.

[0120] The gas turbine engine may comprise a nacelle 21 and the air intake 100 may be comprised as part of the nacelle 21. The nacelle 21 of the gas turbine 10 may have a length 230 of from 1.0 m to 5.0 m; or from 1.7 m to 3.5 m; or from 1.9 m to 3.0 m, or from 1.1 to 2.5 m, or a range of any combination of the aforesaid end points. The length 230 of the nacelle 21 may be measured from highlight 106 to the downstream end 202 of the nacelle 21 at the bypass nozzle 18 as shown in FIG. 1. The length 230 of the nacelle 21 is measured along the engine centreline axis 9 from the intersection of the engine centreline axis with the highlight, to the intersection of the engine centreline axis with a plane defined by the bypass nozzle exit 18.

[0121] The ratio of length of the nacelle 230 to fan diameter is from 0.4 to 2.5, or from 1.2 to 2.0, or from 0.9 to 2.0, or from 0.9 to 1.8, or from 0.5 to 1.2, or a range of any combination of the aforesaid end points.

[0122] The ratio of the intake length 224 to the length of the nacelle 230 may be from 0.1 to 0.75, or from 0.15 to 0.5, or from 0.25 to 0.45, or a range of any combination of the aforesaid end points.

[0123] The nacelle 21 may have a ratio of the length of the nacelle 230 to the nacelle maximum diameter of 1 to 1.5, or 1.1 to 1.35 or 1.2 to 1.3 or a range of any combination of the aforesaid end points.

[0124] The air intake may comprise a non-zero droop or a droop of zero degrees. The intake droop 124 is the angle the intake centreline axis 122 is inclined at relative to the engine centreline axis 9. The air intake may have a droop angle from 0 to 6 degrees, or from 0 to 3 degrees.

[0125] The air intake 100 may comprise a non-zero scarf angle or a scarf angle of zero degrees.

[0126] The scarf angle 120 is the angle between a line from the top dead centre lip to the bottom dead centre lip relative to a line perpendicular and vertically upwards from the intake centreline axis as shown in FIG. 4. The air intake may have a scarf angle from 15 degrees to +10 degrees, or from 5 degrees to +5 degrees. A negative scarf angle suggests the lip at bottom dead centre is further upstream from the fan than the lip at top dead centre. The scarf angle 120 shown in FIG. 4 is a positive scarf angle.

[0127] Wth reference to FIG. 6, the lip 106 of the air intake 100 may have an aspect ratio which is defined by the ratio of the lip length 402 to lip height 404. Lip length 402 is defined by the distance from the lip 106, measured parallel to the intake centreline axis 122, to a line 408 perpendicular to the intake centreline axis passing through the throat 104. The lip height 404 is defined by the shortest distance from the minima of the throat 104 to a line from the lip 106 extending parallel to the intake centreline axis 122. The lip aspect ratio may be measured at individual circumferential points, for example, the lip at top dead centre may have an aspect ratio from 1.8 to 2.8, or from 1.8 to 2.5, or from 2.0 to 2.4, or from any combination of these end points. The lip at bottom dead centre may have an aspect ratio from 1.8 to 3.5, or from 1.8 to 3.2, or from 2.0 to 2.5, or from any combination of these end points. Lip aspect ratio may also be expressed as an average of points around the circumference of the air intake.

[0128] A quasi-non-dimensional mass flow rate Q for the gas turbine engine is defined as:

[00001]$Q=W.Math.\frac{\sqrt{T.Math.0}}{P.Math.0.A_{f.Math.a.Math.n}}.$

[0129] where:

[0130] W is mass flow rate through the fan in kg/s;

[0131] T0 is average stagnation temperature of the air at the fan face in Kelvin;

[0132] P0 is average stagnation pressure of the air at the fan face in Pa;

[0133] A.sub.fan is the area of the fan face in m.sup.2.

[0134] At engine cruise conditions the quasi-non-dimensional mass flow rate Q may be in the range of from 0.029 kgs.sup.1N.sup.1K.sup.1/2 to 0.036 kgs.sup.1N.sup.1K.sup.1/2.

[0135] At cruise conditions, the value of Q may be in the range of from: 0.0295 to 0.0335; 0.03 to 0.033; 0.0305 to 0.0325; 0.031 to 0.032 or on the order of 0.031 or 0.032. Thus, it will be appreciated that the value of Q may be in a range having a lower bound of 0.029, 0.0295, 0.03, 0.0305, 0.031, 0.0315 or 0.032 and/or an upper bound of 0.031, 0.0315, 0.032, 0.0325, 0.033, 0.0335, 0.034, 0.0345, 0.035, 0.0355 or 0.036 (all values in this paragraph being in SI units, i.e. kgs.sup.1N.sup.1K.sup.1/2).

[0136] Cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of decent.

[0137] The local diffuser angle 302 is the angle relative to the engine centreline axis 9 of a line 304 extending between the apex of the throat 104 and the point where the fan face 126 of the fan 23 intersects the inner wall 102 of the nacelle 21, measured at particular circumferential point. Referring to FIG. 6, the local diffuser angle is illustrated for an individual circumferential point at top dead centre. In some embodiments, the local diffuser angle may vary around the circumference of the air intake. In some embodiments, the local diffuser angle at top dead centre may be from 0 to 18 degrees; or from 5 to 18 degrees. In some embodiments, the local diffuser angle at bottom dead centre may be from 0 to 18 degrees; or from 5 to 18 degrees. In some embodiments, the local diffuser angle at one or both lateral sides may be from 0 to 18 degrees; or from 5 to 18 degrees. At other circumferential points, the local diffuser angle can also be from 0 to 18 degrees, or from 5 to 18 degrees.

[0138] The bulk diffuser angle is the mean of the local diffuser angles 302 around the inner circumference of the air intake 100. In some embodiments, the bulk diffuser angle may be from 0 to 15 degrees; or from 3 to 15 degrees.

[0139] The local peak diffuser angle is the angle between the engine centreline axis 9 and a line tangent to the local surface in the diffuser region, such that it gives the largest angle between the diffuser and the engine centreline. In some embodiments, local peak diffuser angle may be from 0 to 22 degrees; or from 6 to 22 degrees.

[0140] It will be further understood by those within the art that if a specific number of an introduced claim recitation is intended, such an intent will be explicitly recited in the claim, and in the absence of such recitation no such intent is present. For example, as an aid to understanding, the following appended claims may contain usage of the introductory phrases at least one and one or more to introduce claim recitations. However, the use of such phrases should not be construed to imply that the introduction of a claim recitation by the indefinite articles a or an limits any particular claim containing such introduced claim recitation to embodiments containing only one such recitation, even when the same claim includes the introductory phrases one or more or at least one and indefinite articles such as a or an (e.g., a and/or an should typically be interpreted to mean at least one or one or more); the same holds true for the use of definite articles used to introduce claim recitations. In addition, even if a specific number of an introduced claim recitation is explicitly recited, those skilled in the art will recognize that such recitation should typically be interpreted to mean at least the recited number (e.g., the bare recitation of two recitations, without other modifiers, typically means at least two recitations, or two or more recitations).

[0141] Furthermore, in those instances where a convention analogous to at least one of A, B, and C, etc. is used, in general such a construction is intended in the sense one having skill in the art would understand the convention (e.g., a system having at least one of A, B, and C would include but not be limited to systems that have A alone, B alone, C alone, A and B together, A and C together, B and C together, and/or A, B, and C together, etc.). In those instances where a convention analogous to at least one of A, B, or C, etc. is used, in general such a construction is intended in the sense one having skill in the art would understand the convention (e.g., a system having at least one of A, B, or C would include but not be limited to systems that have A alone, B alone, C alone, A and B together, A and C together, B and C together, and/or A, B, and C together, etc.). It will be further understood by those within the art that virtually any disjunctive word and/or phrase presenting two or more alternative terms, whether in the description, claims, or drawings, should be understood to contemplate the possibilities of including one of the terms, either of the terms, or both terms. For example, the phrase A or B will be understood to include the possibilities of A or B or A and B.

[0142] All numbers expressing quantities of ingredients, reaction conditions, and so forth used in the specification are to be understood as being modified in all instances by the term about. Accordingly, unless indicated to the contrary, the numerical parameters set forth herein are approximations that may vary depending upon the desired properties sought to be obtained. At the very least, and not as an attempt to limit the application of the doctrine of equivalents to the scope of any claims in any application claiming priority to the present application, each numerical parameter should be construed in light of the number of significant digits and ordinary rounding approaches.

[0143] 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.