GAS TURBINE ENGINE

Abstract

A gas turbine engine for an aircraft includes: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein the engine core extends along a rotational axis, and has an engine core diameter perpendicular to the rotational axis; and a fan having a plurality of fan blades extending radially from the rotational axis, wherein the fan has a fan diameter perpendicular to the rotational axis, wherein a ratio of the engine core diameter to the fan diameter is between 1:1.7 and 1:2.2.

Claims

1. 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, wherein the engine core extends along a rotational axis, and has an engine core diameter perpendicular to the rotational axis; and a fan comprising a plurality of fan blades extending radially from the rotational axis, wherein the fan has a fan diameter perpendicular to the rotational axis, wherein a ratio of the engine core diameter to the fan diameter is between 1:1.7 and 1:2.2.

2. The gas turbine engine of claim 1, wherein the ratio of the engine core diameter to the fan diameter is between 1:1.7 and 1:1.8.

3. The gas turbine engine of claim 1, wherein the fan diameter is at least 2.2 metres.

4. 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, wherein the engine core extends along a rotational axis, and has an engine core diameter perpendicular to the rotational axis; and a fan comprising a plurality of fan blades extending radially from the rotational axis, wherein the fan has a fan diameter perpendicular to the rotational axis, wherein: a ratio of the engine core diameter to the fan diameter is between 1:1.65 and 1:2.2; and the fan diameter is at least 2.2 metres.

5. A gas turbine engine as claimed in claim 4, wherein the ratio of the engine core diameter to the fan diameter is between 1:1.65 and 1:1.8.

6. A gas turbine engine as claimed in claim 3, wherein the fan diameter is greater than or equal to 2.5 metres.

7. A gas turbine engine as claimed in claim 3, wherein the fan diameter is greater than or equal to 3.3 metres.

8. A gas turbine engine as claimed in claim 7, wherein the fan diameter is less than or equal to 3.7 metres.

9. A gas turbine engine as claimed in claim 1, wherein the gas turbine engine generates a maximum thrust up to 225 kN.

10. A gas turbine engine as claimed in claim 1, wherein the gas turbine engine generates a maximum thrust of 310 kN or more.

11. A gas turbine engine according to claim 1, wherein the engine core diameter varies along the rotational axis and the ratio of the engine core diameter to the fan diameter comprises the ratio of a maximum engine core diameter to the fan diameter.

12. A gas turbine engine according to claim 1, 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.

13. A gas turbine engine according to claim 12, wherein the gearbox is an epicyclic gearbox.

14. A gas turbine engine according to claim 13, wherein the gearbox comprises: a central sun gear coupled to the core shaft, and arranged to rotate around the rotational axis; a planet gear carrier arranged to rotate about the rotational axis and coupled to the fan; a plurality of planet gears mounted on the planet gear carrier, the planet gears radially outwards of and intermeshing with the sun gear; and a ring gear radially outward of and intermeshing with the planet gears, and held stationary with respect to the sun gear wherein the planet carrier is arranged to hold the planet gears in a fixed spacing relative to each other, and to enable each planet gear to rotate about its own axis, such that rotation of the sun gear drives rotation of the planet gears about their own axes, causing precession of the planet gears about the sun gear, in synchronicity, to drive rotation of the planet carrier.

15. A gas turbine engine according to claim 1, wherein: the turbine is a first turbine, the compressor is a first compressor 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.

16. A gas turbine engine according to claim 1, comprising a fan case arranged around the fan and extending along the rotational axis, wherein: the engine core is comprised in a first module of the gas turbine engine, the fan is comprised in a second module of the gas turbine engine, and the fan case is comprised in a third module of the gas turbine engine; one or more of the first, second and third modules are interchangeable with different first, second or third modules.

Description

[0053] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0054] FIG. 1 is a sectional side view of a gas turbine engine;

[0055] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0056] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0057] FIG. 4A illustrates a schematic view of the gas turbine engine of FIG. 1, illustrating the separate modules of the engine; and

[0058] FIG. 4B illustrates the modules of FIG. 4A, in exploded form.

[0059] 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. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0060] The propulsive fan 23 includes a plurality of fan blades 25 extending radially outward from a hub 29 mounted on an output shaft of the gearbox 30. The radially outer tips of the fan blades 25 are surrounded by a fan casing 42, which extends downstream behind the fan 23. The fan casing 42 will be discussed in more detail below, in relation to FIGS. 4A and 4B. Behind the fan casing 42, in the axial flow direction (downstream), a nacelle 21 surrounds the engine core 11. The fan casing 42 and nacelle 21 define a bypass duct 22 and a bypass exhaust nozzle 18 around the engine core 11.

[0061] The bypass airflow B flows through the bypass duct 22. At an upstream end of the bypass duct 22, adjacent an intake 31 of the bypass duct 22, and downstream of the fan 23, a plurality of outlet guide vanes 33 extend radially between the engine core 11 and the fan casing 42. The outlet guide vanes 33 reduce swirl and turbulence in the bypass airflow B, providing improved thrust.

[0062] 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.

[0063] 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.

[0064] 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.

[0065] 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.

[0066] 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.

[0067] 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.

[0068] 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.

[0069] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0070] 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. The gas turbine engine 10 may also be arranged in the pusher configuration, in which the fan 23 is located downstream of the core 11. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

[0071] 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.

[0072] FIG. 4A schematically illustrates the constituent components of the gas turbine engine 10 of FIGS. 1 to 3, with the nacelle 21 removed. As shown in FIG. 4B, the gas turbine engine 10 is formed of a number of separate modules 11, 23, 35. The engine 10 may thus be considered modular.

[0073] The first module is an engine core module 11. This typically includes the gearbox 30, low pressure compressor 14, high-pressure compressor 15, combustion equipment 16, high-pressure turbine 17, and low pressure turbine 19. The engine core module 11 can also be referred to as a propulsor. The second module, also referred to as the fan module 23, includes the fan blades 25. The third module 35 includes the fan case 42.

[0074] The outlet guide vanes 33 extend inwardly from the fan case 42, and typically form part of the fan case module 35. The hub 29 and gearbox 30 may be part of the fan module 23 or the engine core module 11. The gearbox 30 may additionally be configured as a separable module in its own right or part of the fan case module 35.

[0075] As shown in FIG. 4B, the fan module 35 can be removed from the engine core module 11, and the engine core module 11 and fan case module 35 can be separated from one another. This facilitates easy delivery and transport of the engine 10, as the separate modules 11, 23 35. Any suitable connections may be used to join the modules. For example, the fan case module 35 may be bolted to the engine core 11 by bolted connections at the radially inner ends of the outlet guide vanes 33. Further connecting/support struts may also be provided between the fan case 42 and the engine core 11.

[0076] The modules 11, 23, 35 may be interchangeable, such that, for example, a gas turbine engine 10 that includes a first engine core module 11, a first fan module 23 and a first fan case module 35 may have the first engine core module 11 removed, and replaced with a second engine core module 11 having the same design. The second engine core module 11 may have the same design at least with respect to the interfaces between the modules.

[0077] It will be appreciated that any one or more of the modules 11, 23, 35 may be interchanged. There may be a plurality of engine core modules 11, a plurality of fan modules 23 and a plurality of fan case modules 35. An engine 10 may include any one of each of the modules, rather than each engine 10 comprising dedicated sets of modules that can only be used together (i.e. the first engine core module 11 only works with the first fan module 23 and the first fan case module 35, the second engine core module 11 only works with the second fan module 23 and the second fan case module 35, and the like).

[0078] The interchangeability of modules allows the first engine core module 11 to be serviced, replaced or repaired, whilst the engine 10 remains functional. The engine core module 11 is smaller than the fan case 42, and also requires more regular maintenance. Therefore, by using a modular engine with interchangeable modules, the smaller, easier to transport parts, are shipped more easily, whilst the larger parts are kept with the aircraft, or in an aircraft maintenance facility.

[0079] In some cases, the nacelle 21 remains with the aircraft. In other instances, the whole or part of the nacelle 21, such as the region around the air intake, could be removed from the aircraft with the fan case module 35.

[0080] The engine core 11 is encased in an engine core housing 44. The engine core 11 has a radius measured from the centre line of the engine 10 (the principal axis 9), to the engine core housing 44, in a radial direction perpendicular to the housing. The diameter of the engine core 11 is twice the radius.

[0081] As shown in FIGS. 4A and 4B, the diameter of the engine core 11 varies along the axial length of the engine 10. The engine core 11 has a maximum diameter 46 at a point along its length. In one example, the maximum diameter 46 may be in the region of the intermediate pressure compressor 15, but this need not necessarily be the case.

[0082] The fan 23 also has a radius. The radius of the fan 23 is measured between the engine centreline 9 and the tip of a fan blade 25 at its leading edge. As with the engine core 11, the diameter 48 of the fan 23 is twice the radius.

[0083] The size of the engine 10 is described by the diametral ratio of the engine 10, which is the ratio of the maximum engine core diameter 46 to the fan diameter 48.

[0084] The diametral ratio of the engine 10 discussed above and shown in the Figures may be between 1:1.65 and 1:2.2. For example, the diametral ratio may be between 1:1.7 and 1:2.2, or may further be between 1:1.65 and 1:1.8 or 1:1.7 and 1:1.8, or indeed any other range defined and/or claimed herein.

[0085] The fan diameter 48 may be above 90 inches (2.286 metres). For example, the fan diameter 48 may be around 101 inches (2.565 metres), or above 130 inches (3.302 metres). For example, the fan diameter 48 may be between 130 inches and 145 inches (3.683 metres). For example, the fan diameter 48 may be on the order of 140 inches (3.556 metres).

[0086] When the diametric ratio is within any of the ranges discussed above, the engine 10 may be arranged to generate a maximum thrust of around 50,000 lbf (222 kN). In alternative examples, the engine 10 may be arranged to generate a maximum thrust of over 70,000 lbf (311 kN). In one such example, the engine 10 may be arranged to generate in the range of from 80,000 lbf thrust (356 kN) to 90,000 lbf thrust (400 kN).

[0087] The diametral ratios discussed above may be applicable, however, to any fan diameter 48, and any thrust level. The thrust level will be dependent, at least in part, on the fan diameter 48.

[0088] The efficiency of a gas turbine engine 10 can be characterised by the bypass ratio. The bypass ratio is the ratio of the air mass flow through the bypass duct, flow B, to the air mass flow through the core, flow A. As the bypass ratio increases, the proportion of the thrust generated by the core engine 11 decreases.

[0089] Under mid-cruise conditions (when the engine is flying a stable altitude and thrust levels), and with the diametral ratios discussed above, as an example the engine 10 can have a bypass ratio of 14.5 or more, or indeed any other bypass ratio described and/or claimed herein.

[0090] Although the engine 10 has been described in relation to a modular geared gas turbine engine 10 for an aircraft, it may in some cases be used on any suitable gas turbine engine, including non-geared and/or non-modular engines.

[0091] Furthermore, the architecture of the engine 10 discussed above is given by way of example only. The engine 10 may have any suitable architecture. For example, there may be any number of compressor stages, turbine stages, core shafts and the like.

[0092] The engine core module 11 may include features on the outside of the housing 44, or protruding from the housing 44. These protruding features may be, for example, mounting the engine core 11 to the fan case module 35 and support structure of the engine, fixing outlet guide vanes 33 to the engine core 11, and features associated with accessories of the engine and the like. The protruding features may be integral with the engine core 11, or removable, and are typically discontinuous along the length and/or around the circumference of the engine core 11. The protruding features are typically used when the engine is operational. Whilst the features may be used for storage or transportation, they are not provided solely for this purpose. The core diameter 46 is measured from the core housing 44, and does not include such features.

[0093] 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.