NACELLE INTAKE

Abstract

A nacelle defining an intake is provided for channelling and conditioning freestream airflow to a fan of a ducted fan gas turbine engine. The nacelle intake comprises in flow series an intake lip, and a diffuser. The intake lip has at its forwardmost end a highlight which is a closed loop defining a boundary between inner and outer surfaces of the nacelle, and the diffuser terminates at its rearwardmost end at a front face of the fan. The intake has a circumferentially extending shock control bump or a series of circumferentially spaced shock control bumps formed thereon. On longitudinal cross sections through the intake containing the bump or series of bumps: the bump or each bump of the series of bumps has a profile which forms in flow series an up-ramp, a maximum and a down-ramp, at least one of the up-ramp and the down-ramp forming an inflection point or curvature discontinuity point on the profile, and the inflection point or curvature discontinuity point being axially located between the highlight and a position which is axially rearward therefrom by a distance of no more than 0.6 L, where L is the axial distance between the highlight and the front fan face.

Claims

1. A nacelle defining an intake for channelling and conditioning freestream airflow to a fan of a ducted fan gas turbine engine (1), the intake comprising in flow series an intake lip, and a diffuser, the intake lip having at its forwardmost end a highlight (H) which is a closed loop defining a boundary between inner and outer surfaces of the nacelle, and the diffuser terminating at its rearwardmost end at a front face of the fan; wherein the intake has a circumferentially extending shock control bump or a series of circumferentially spaced shock control bumps formed thereon; and wherein, on longitudinal cross sections through the intake containing the bump or series of bumps: the bump or each bump of the series of bumps has a profile which forms in flow series an up-ramp, a maximum and a down-ramp, at least one of the up-ramp and the down-ramp forming an inflection point or curvature discontinuity point on the profile, and the inflection point or curvature discontinuity point being axially located between the highlight and a position which is axially rearward therefrom by a distance of no more than 0.6 L, where L is the axial distance between the highlight and the front fan face.

2. The nacelle according to claim 1, wherein the bump or series of bumps is configured such that radial inwards displacement of the intake by the bump or series of bumps varies, in the circumferential direction, from a maximum at about the centre of the bump or series of bumps to a minimum at the ends of the bump or series of bumps.

3. The nacelle according to claim 1, wherein the bump or series of bumps is centred, in the circumferential direction, on the bottom dead centre position of the intake.

4. The nacelle according to claim 1, wherein on a closed loop around the intake coinciding with the bump or series of bumps, a first portion of the loop is without the bump or series of bumps, and a remaining second portion of the loop contains the bump or series of bumps

5. The nacelle according to claim 4, wherein the first portion is one half of the circumferential extent of the closed loop, and the second portion is the other half of the circumferential extent of the closed loop.

6. The nacelle according claim 1, wherein the or each bump is smoothly curved and is tangential to the surrounding parts of the intake.

7. The nacelle according to claim 1, wherein the or each bump is facetted.

8. The nacelle according to claim 1, wherein the series of circumferentially spaced shock control bumps is formed on the intake, the bumps being configured such that, between the bumps, the intake displaces radially outwards to compensate for fan frontal area blocked by the bumps.

9. 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, the fan comprising a plurality of fan blades; and the nacelle defining an intake according to claim 1 for channelling and conditioning freestream airflow to the fan.

10. The gas turbine engine according to claim 9, 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.

11. The gas turbine engine according to claim 10, 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.

12. A nacelle defining an intake for channelling and conditioning freestream airflow to a fan of a ducted fan gas turbine engine, the intake comprising in flow series an intake lip, and a diffuser, the intake lip having at its forwardmost end a highlight which is a closed loop defining a boundary between inner and outer surfaces of the nacelle, and the diffuser terminating at its rearwardmost end at a front face of the fan; wherein the intake has a circumferentially extending series of circumferentially spaced shock control bumps formed thereon; and wherein the series of bumps is configured such that radial inwards displacement of the intake by the series of bumps varies, in the circumferential direction, from a maximum at about the centre of the bump or series of bumps to a minimum at the ends of the series of bumps.

13. The nacelle according to claim 12, wherein the bump or series of bumps is centred, in the circumferential direction, on the bottom dead centre position of the intake.

14. The nacelle according to claim 12, wherein, on longitudinal cross sections through the intake containing the series of bumps: each bump of the series of bumps has a profile which forms in flow series an up-ramp, a maximum and a down-ramp, at least one of the up-ramp and the down-ramp forming an inflection point or curvature discontinuity point on the profile, and the inflection point or curvature discontinuity point being axially located between the highlight and a position which is axially rearward therefrom by a distance of no more than 0.6 L, where L is the axial distance between the highlight and the front fan face.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0049] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0050] FIG. 1 shows schematically a bottom half of a longitudinal cross section of an intake of a ducted fan gas turbine engine at high angle of attack;

[0051] FIG. 2 is a sectional side view of a gas turbine engine;

[0052] FIG. 3 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0053] FIG. 4 is a partially cut-away view of a gearbox for a gas turbine engine;

[0054] FIG. 5 shows schematically a longitudinal cross section through the nacelle air intake of the engine shown in FIG. 2;

[0055] FIG. 6 shows schematically a front view of the intake of FIG. 5;

[0056] FIG. 7 shows schematically a bottom half of the longitudinal cross section of FIG. 5 at high angle of attack;

[0057] FIG. 8 shows plots of variation in pressure loss at the fan face against incidence angle for a conventional nacelle and a nacelle with a shock bump;

[0058] FIG. 9 shows schematically a front view of a variant intake;

[0059] FIG. 10 shows schematically a front view of another variant intake;

[0060] FIG. 11 shows schematically a front view of another variant intake;

[0061] FIG. 12 shows schematically a front view of another variant intake;

[0062] FIG. 13 shows schematically a bottom half of a longitudinal cross section of the intake of FIG. 12 at high angle of attack;

[0063] FIG. 14 shows schematically a bottom half of a longitudinal cross section of another variant intake; and

[0064] FIG. 15 shows schematically a front view of another variant intake.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0065] FIG. 2 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.

[0066] 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.

[0067] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 3. The low pressure turbine 19 (see FIG. 2) 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.

[0068] 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.

[0069] The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 4. 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. 4. 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.

[0070] 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.

[0071] 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. 3 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. 3. 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. 3.

[0072] 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.

[0073] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0074] 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. 2 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle 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.

[0075] 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. 2), and a circumferential direction (perpendicular to the page in the FIG. 2 view). The axial, radial and circumferential directions are mutually perpendicular.

[0076] A longitudinal cross section through the nacelle air intake 12 of the engine 10 of FIG. 2 is shown schematically in FIG. 5, and a front view of the intake is shown schematically in FIG. 6. The lightly scarfed intake comprises in flow series an intake lip 40, and a diffuser 42. The forwardmost end of the lip is a highlight H, which is a closed loop defining a boundary between inner and outer surfaces of the nacelle 21. The diffuser terminates at its rearwardmost end at a front face of the fan. The intake uses a circumferentially extending shock control bump 48 to weaken shock waves at high incidence angles , therefore reducing shock-induced separation and extending the safe operating range of the engine.

[0077] Shock control bumps can weaken shock waves on aerodynamic surfaces to reduce entropy generation and therefore drag. See Qin, N., W. S. Wong, and A. Le Moigne, Three-dimensional contour bumps for transonic wing drag reduction, Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 222.5 (2008): 619-629, and Bruce, P. and Colliss, S. (2015), Review of research into shock control bumps, Shock Waves, 25(5):451-471.

[0078] In FIGS. 5 and 6, the shock control bump 48 is added to the lower half of the intake 12. The bump displaces the intake radially inwards relative to a datum geometry D (indicated by dotted line) where it would otherwise be.

[0079] The bump 48 is smoothly curved and is preferably tangential to the upstream and downstream parts of the intake 12. Nonetheless, it forms a pre-compression ramp/bump that decelerates the flow upstream of the shock, thus weakening it. On longitudinal cross sections, such as shown in FIG. 5, the bump has a profile which forms in flow series an up-ramp 46 with a certain curvature distribution, a maximum M and a down-ramp 44 with a certain curvature distribution, with at least one of the up-ramp and the down-ramp forming an inflection point or curvature discontinuity point in the intake. Moreover, the inflection or curvature discontinuity point (or both points if there is one on the up-ramp and one on the down-ramp) is axially located between the highlight H and a position which is axially rearward therefrom by a distance of no more than 0.6 L (and preferably no more than 0.5 L or 0.4 L), where L is the axial distance between the highlight and the front fan face.

[0080] In the example of FIG. 5, there is an inflection or curvature discontinuity point on the down-ramp at about point P. Additionally or alternatively, the bump 48 may have an inflection or curvature discontinuity point on the up-ramp, However, in this case, because the curvature of the intake 12 at the highlight H and at the maximum M of the bump generally have the same sign, the bump typically requires two such points sequentially on the up-ramp. Moreover, particularly if the bump only has inflection or curvature discontinuity point(s) on the up-ramp, the intake may have a further inflection point or curvature discontinuity point in the diffuser 42 relatively close to the fan face (i.e. outside the specified the 0.6 L distance from the highlight H). However, this further point relates to diffuser functionality and is not generally relevant for shock control.

[0081] The bump 48 is configured such that the radial inwards displacement of the intake 12 produced by the bump is a maximum at about the centre of the lower half (i.e. at bottom dead centre) and decays in magnitude in the circumferential direction (i.e. angular direction ) to zero at about the 3 o'clock and 9 o'clock positions. This matches shock strength/separation, which is typically maximum at or close to bottom dead centre to coincide with maximum acceleration due to local peak incidence angle .

[0082] Referring to FIG. 7, which is for comparison with FIG. 1, the bump 48 weakens the shock shown in FIG. 1, and turns it into a series of isentropic compression waves Y, usually terminated by a much weaker shock. These compression waves (and weaker shock) no longer cause separation, such that the boundary layer Z stays attached. This results in reduced pressure loss, decreases the blockage caused to the fan, increases engine performance and allows the engine to operate without separation occurring at higher incidences than would otherwise be possible.

[0083] Thus, as shown in FIG. 8, the operating range of engine intakes (i.e. the maximum angle of incidence before catastrophic separation and pressure loss occurs) can be increased through the use of shock control bumps. Alternatively, the use of shock control bumps can allow intake geometries to be made thinner and lighter, whilst maintaining the operating range.

[0084] In FIGS. 5 and 6, the height of the shock control bump 48 varies smoothly from a maximum at bottom dead centre to zero at 3 o'clock and 9 o'clock positions. However, any one or more of the bump height, position, shape and width (axial extent) can vary arbitrarily with circumferential position (i.e. with ), as shown in FIG. 9. This allows the bump to be optimised a for a specific nacelle geometry and/or flow condition.

[0085] Indeed, instead of a single bump, a series of circumferentially spaced shock control bumps 48 can be used, as shown in FIG. 10. In this example, the intake returns to a datum geometry between the bumps. This blocks less of the fan frontal area, than a single continuous bump, and can thereby increase mass flow through the fan while still weakening the shock. In the further example of FIG. 11, this concept is extended such that the intake displaces radially outwards of the datum geometry between the bumps. In this way the intake flow area can be made at least the same as a conventional intake without shock control bumps.

[0086] Yet another option is to use a facetted bump or a series of facetted bumps instead of the smoothly curved bump(s) discussed above in respect of FIGS. 5 to 11. FIG. 12 shows schematically a front view of an intake with a series of circumferentially spaced facetted shock control bumps 48, and FIG. 13 shows schematically, for comparison with FIGS. 1 and 7, a bottom half of a longitudinal cross section of the intake of FIG. 12 at high angle of attack. Each facetted bump 48 has a straight-lined up-ramp and a straight-lined down-ramp with a flat therebetween, the up-ramp and down-ramp both being non-tangential to the upstream and downstream parts of the intake 12.

[0087] The vertical structures produced by facetted bumps 48 enhance boundary layer mixing to help keep flow attached, and the facetted bumps create a lambda type shock Y that is weaker than the normal shock Y shown in FIG. 1, thereby generating less loss and causing less separation. In general, however, due to entropy increase, the lambda shock Y has a higher loss than the compression waves Y of FIG. 7. Facetted (or ramp) bumps are discussed in Bruce, P. and Colliss, S., (2015), Review of research into shock control bumps, Shock Waves, 25(5):451-471.

[0088] FIG. 14 shows a variant of the facetted bump of FIG. 13. In this case, the bump 48 has a smoothly curved up-ramp and down-ramp. However, the ramps are both non-tangential to the upstream and downstream parts of the intake 12, and help to create a lambda type shock.

[0089] The shock control bump 48 or series of shock control bumps discussed above in respect of FIGS. 5 to 14 extend(s) over the lower half of the intake 12. However, the circumferential extent of the bump or series of bumps can be less than or greater than this, depending on the nacelle geometry and flow conditions. Additionally or alternatively, as shown in FIG. 15, the midpoint of the bump 48 or series of bumps can be varied, and does not have to be bottom dead centre. For example, the bump or series of bumps can be to a side of the intake to reduce separation due to runway cross-winds.

[0090] 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.