Gas turbine engine with a double wall core casing

Abstract

A gas turbine engine includes an engine core including: a compressor system including first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, wherein the first flange connection is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor; a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; wherein an axial midpoint of the radially outer edge is defined as the fan OGV tip centrepoint.

Claims

1. A gas turbine engine for an aircraft comprising: an engine core comprising: a compressor system with compressor blades comprising respective aerofoils, the compressor system comprising a first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system and comprising a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, the first flange connection having a first flange radius, the first flange connection being the first flange connection that is downstream of an axial position defined by an axial midpoint between a mid-span axial location on a trailing edge of a most downstream aerofoil of the first compressor and a mid-span axial location on a leading edge of a most upstream aerofoil of the second compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades and having a fan diameter; a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; and a fan outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and an inner surface of the nacelle, the fan OGV having a radially inner edge and a radially outer edge, an axial midpoint of the radially outer edge being defined as a fan OGV tip centrepoint, wherein a fan OGV tip position ratio of: $\frac{\begin{array}{c}\mathrm{an}\mathrm{axial}\mathrm{distance}\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\mathrm{and}\mathrm{the}\mathrm{fan}\mathrm{OGV}\mathrm{root}\mathrm{centrepoint}\end{array}}{\mathrm{the}\mathrm{first}\mathrm{flange}\mathrm{radius}}$ is equal to or less than 1.8.

2. The gas turbine engine of claim 1, wherein the fan OGV tip position ratio is greater than or equal to 0.6.

3. The gas turbine engine of claim 1, wherein the fan OGV tip position ratio is less than or equal to 1.20.

4. The gas turbine engine of claim 1, wherein the fan diameter is greater than 240 cm and less than or equal to 380 cm.

5. The gas turbine engine of claim 1, wherein the fan diameter is between 330 cm and 380 cm.

6. The gas turbine engine of claim 1, wherein a number of the fan blades is between 16 and 22.

7. The gas turbine engine of claim 1, further comprising a gearbox that is configured to receive an input from a core shaft, and output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, and optionally wherein a gear ratio of the gearbox is between 3.1 and 4.0.

8. The gas turbine engine of claim 1, wherein the first flange connection is at, or axially downstream of, the leading edge of the most upstream aerofoil of the second compressor.

9. The gas turbine engine of claim 1, wherein the first flange connection is at, or axially upstream of, the leading edge of the most upstream aerofoil of the second compressor.

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

11. The gas turbine engine according to claim 1, wherein a fan OGV tip position to fan diameter ratio of: $\frac{\begin{array}{c}\mathrm{the}\mathrm{axial}\mathrm{distance}\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\mathrm{and}\mathrm{the}\mathrm{fan}\mathrm{OGV}\mathrm{tip}\mathrm{centrepoint}\end{array}}{\mathrm{the}\mathrm{fan}\mathrm{diameter}}$ is less than or equal to 0.22.

12. The gas turbine engine according to claim 11, wherein the fan OGV tip position to fan diameter ratio is greater than or equal to 0.095.

13. The gas turbine engine of claim 12, wherein the fan OGV root position ratio is greater than or equal to 0.8.

14. The gas turbine engine according to claim 12, wherein a fan OGV root position to fan diameter ratio of: $\frac{\begin{array}{c}\mathrm{the}\mathrm{axial}\mathrm{distance}\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\mathrm{and}\mathrm{the}\mathrm{fan}\mathrm{OGV}\mathrm{root}\mathrm{centrepoint}\end{array}}{\mathrm{the}\mathrm{fan}\mathrm{diameter}}$ is less than or equal to 0.33.

15. The gas turbine engine according to claim 1, further comprising a front mount arranged to be connected to a pylon, wherein a front mount position ratio of: $\frac{\begin{array}{c}\mathrm{an}\mathrm{axial}\mathrm{distance}\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\mathrm{and}\mathrm{the}\mathrm{front}\mathrm{mount}\end{array}}{\mathrm{the}\mathrm{first}\mathrm{flange}\mathrm{radius}}$ is equal to or less than 1.18.

16. The gas turbine engine according to claim 15, wherein a front mount position to fan diameter ratio of: $\frac{\begin{array}{c}\mathrm{the}\mathrm{axial}\mathrm{distance}\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\mathrm{and}\mathrm{the}\mathrm{front}\mathrm{mount}\end{array}}{\mathrm{the}\mathrm{fan}\mathrm{diameter}}$ is less than or equal to 0.145.

17. The gas turbine engine according to claim 1, wherein: the engine core further comprises an inner core casing provided radially inwardly of the compressor blades of the compressor system, the inner core casing and the outer core casing defining a core working gas flow path therebetween, a gas path radius is defined as an outer radius of the core working gas flow path at the axial position of the first flange connection, and a gas path ratio of: $\frac{\mathrm{the}\mathrm{first}\mathrm{flange}\mathrm{radius}}{\mathrm{the}\mathrm{gas}\mathrm{path}\mathrm{radius}}$ is equal to or greater than 1.10 and less than or equal to 2.0.

18. The gas turbine engine according to claim 1, wherein a fan diameter ratio of: $\frac{\mathrm{the}\mathrm{first}\mathrm{flange}\mathrm{radius}}{\mathrm{the}\mathrm{fan}\mathrm{diameter}}$ is equal to or greater than 0.125 and less than or equal to 0.17.

19. The gas turbine engine according to claim 1, wherein a fan blade mass ratio of: $\frac{\mathrm{the}\mathrm{first}\mathrm{flange}\mathrm{radius}}{a\mathrm{mass}\mathrm{of}\mathrm{each}\mathrm{fan}\mathrm{blade}}$ is equal to or less than 19.0 mm/lb.

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. 3A is a partially cut-away view of a gearbox for a gas turbine engine;

(5) FIG. 3B shows a schematic view of flange connections of a gas turbine engine;

(6) FIG. 4A shows a schematic view of an engine;

(7) FIG. 4B shows a shear force diagram corresponding to FIG. 4A;

(8) FIG. 4C shows a bending moment diagram corresponding to FIG. 4A;

(9) FIG. 5 is a close up sectional side view of an upstream portion of a gas turbine engine with an intercase portion highlighted;

(10) FIG. 6A is a close up sectional side view of an intercase portion;

(11) FIG. 6B is a close up sectional side view of a different intercase portion;

(12) FIG. 7 is a close up sectional side view of a portion of a gas turbine engine behind the fan, with component spacings and radii marked;

(13) FIG. 8 is an enlarged view of a portion of FIG. 7;

(14) FIG. 9 shows an aircraft with two engines mounted thereon;

(15) FIG. 10 is a schematic sectional side view of the mounting of an engine to a wing of the aircraft;

(16) FIG. 11A is a schematic sectional side view illustrating first flange position in one embodiment; and

(17) FIG. 11B is a schematic sectional side view illustrating first flange position in another embodiment.

(18) 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.

(19) 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.

(20) 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.

(21) 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.

(22) The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3A. 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. 3A. 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.

(23) 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.

(24) 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.

(25) 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.

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

(27) 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.

(28) 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.

(29) The engine 10 can be subject to bending due to both static and dynamic loading conditions. A simplified engine bending scenario is show in FIGS. 4A, 4B and 4C. In the embodiment being described, the engine 10 is a medium-large, geared gas turbine engine 10, having a fan diameter 112 greater than 240 cm, and more particularly greater than 300 cm. The engine 10 of the embodiment being described may therefore be described as a large engine, for example having a fan diameter 112 between 330 cm and 380 cm, and optionally between 335 cm and 360 cm, a gear ratio between 3.1 and 4.0, and a number of fan blades between 16 and 22.

(30) A schematic side view of the gas turbine engine 10 is shown in FIG. 4A, with the rotational axis 9 extending horizontally. FIG. 4B shows a diagram of shear force with distance along the rotational axis 9, aligned with the schematic engine view of FIG. 4A. FIG. 4C shows a diagram of bending moment with distance along the rotational axis, aligned with the schematic engine view of FIG. 4A.

(31) Arrow X in FIG. 4A indicates intake upload on the fan 23. The skilled person would appreciate that take-off is generally the most severe condition where the intake upload is a maximum.

(32) Arrow Y in FIG. 4A indicates reaction load at the front mount 50 due to intake upload. The skilled person would appreciate that gas turbine engines 10 are generally mounted to the wing 52 of an aircraft 90 by one or more pylons 53, as illustrated in FIGS. 9 and 10. The pylons 53 may be secured to the engine core 11, to the nacelle 21, or to both. The or each pylon 53 may be secured to the engine 10 at multiple points—multiple mounts may therefore be provided for the pylon(s). The front mount 50 is the forward-most mount on the engine 10, and may be located on the core 11 or on the nacelle 21 in various embodiments. The rear mount is the rear-most mount on the engine 10, and may be located on the core 11 or on the nacelle 21 in various embodiments. In FIG. 4A, the front and rear mount are both indicated to be on the engine core 11—one or both may instead be on the nacelle 21 in other embodiments.

(33) Arrow Z in FIG. 4A indicates reaction load at the rear mount. In FIG. 4A, the front and rear mounts are both indicated to be on the core 11. In alternative embodiments, such as that shown in FIG. 10, the front mount 53a may be on the nacelle 21, and the rear mount 53b may be on the core 11. In such embodiments, multiple mounts may be present on the core 11—the mount 53b may be referred to as the front core mount as it is the forward-most (or only) mount present on the core 11.

(34) The engine core 11 is therefore designed to react the bending moment with sufficient resistance to reduce or minimise performance losses due to casing deformations. The skilled person would appreciate that increased deformations leads to increased tolerances being needed, such as an increased blade tip to casing gap, so potentially resulting in decreased efficiency. Additionally or alternatively, casing deformations may result in increased wear on bearings, joints and the like, so potentially reducing engine lifespan.

(35) The skilled person would appreciate that the structural load path of a gas turbine engine 10 generally comprises bearing structures, which are relatively high in stiffness, and rotor/combustor casings, which are relatively weaker. Flanges that join bearing structures to casings, and/or casing portions to other casing portions, are therefore likely to be areas where significant changes in stiffness occur. Regions containing one or more flanges may therefore be regions where the slope (dw/dl—delta deformation over delta length) tends to be severe.

(36) The casing surrounding the core 11 is arranged to be separated at one or more positions along its length. A flange connection may be provided to allow separation of the casing into different portions. The positioning of such a flange connection may be constrained by flange integrity considerations. In design studies it has been observed that engine stiffness can be improved by moving a flange connection provided to connect portions of the casing (referred to as the first flange connection 60 in the embodiments described herein) further from the engine axis 9—i.e. to a higher diameter relative to the gas path.

(37) Referring to FIG. 5, the low pressure compressor 14 and the high pressure compressor 15 together form a compressor system. The compressor system is shown in FIG. 5, and in the close up views of FIGS. 6A and 6B.

(38) Each compressor 14, 15 of the compressor system comprises a respective axial compressor having one or more compressor stages, in the embodiments being described. In alternative embodiments, one or more centrifugal compressors may be used. In the embodiments being described, each compressor stage comprises a rotor and a stator. In the described embodiment, each of the high pressure compressor 15 and the lower pressure compressor 14 comprise two stages formed by a respective first rotor 62a, 62b, first stator 64a, 64b, second rotor 66a, 66b and second stator 68a, 68b. Each of the rotors provided in the compressors 14, 15 are formed from an annular array of rotor blades arranged to rotate in order to provide compression of airflow through the engine 10. Each of the stators comprises an annular array of stator blades that are stationary. The rotor blades and stator blades can each be described as aerofoils provided in the compressors 14, 15.

(39) In the described embodiment two stages are provided in each compressor 14, 15. In other embodiments, any other suitable number of stages may be provided such as a single stage or three or more stages. The number of stages in each compressor may be the same, as illustrated, or different from each other.

(40) The engine core 11 further comprises a radially inner core casing 70, which is provided radially outwardly of the interconnecting shafts 26, 27 connecting the low and high pressure compressors 14, 15 to the respective low and high pressure turbines 17, 19. The inner core casing 70 is provided radially inwardly of the blades of the compressors 14, 15. The inner core casing 70 extends in a generally axial direction between an inlet 72 downstream of the fan 23 and upstream of the low pressure compressor 15 to an outlet 74 downstream of the high pressure compressor 15 and upstream of the combustion equipment 16.

(41) The engine core 11 further comprises an outer core casing 76 that generally surrounds the compressor system. The outer core casing 76 is provided radially outwardly of the inner core casing 70 and the tips of the stators and rotors provided in the compressors 14, 15. The core airflow path A is defined between a radially outer surface of the inner core casing 70 and a radially inner surface of the outer core casing 76. The engine outer core casing 76 extends between the inlet 72 and the outlet 74 similarly to the inner core casing 70.

(42) The outer core casing 76 comprises a single wall in a forward region of the engine 10, and a first outer core casing 78 and a second outer core casing 80 in a rearward region of the engine 10, in the embodiment being described. As can be seen in FIG. 5 and the close up view of FIG. 6A, the outer core casing 76 bifurcates into the first and second outer core casings 78, 80 at a point along its axial length downstream (rearward) of the low pressure compressor 14 and upstream (forward) of the high pressure compressor 15. The first and second outer core casings 78, 80 are spaced apart by a gap extending along the axis 9. In the described embodiment, therefore, only part of the axial length of the outer core casing 76 is formed from the first and second outer core casings 78, 80. In other embodiments, separate first and second core casings 78, 80 could also extend across the low pressure compressor 14, and optionally across the full length of the outer core casing 76, or a single wall outer core casing 76 may extend the full length.

(43) The first outer core casing 78 is provided radially inwardly of the second outer core casing 80. The inner surface of the first outer core casing 78 forms the inner surface of the outer core casing 76 which contains gas flow within the core airflow A. The first and second outer core casings 78, 80 each provide a separate function within the engine 10. The first outer core casing 78 is adapted to contain the core airflow A. It may therefore be wholly annular and is generally airtight (save for access for bleed ports or the like). The second outer core casing 80 is instead adapted to provide structural support (i.e. it may provide only structural support). It may not therefore need to be wholly annular or airtight. In other embodiments, both pressure containment and structural support may be provided by both the first and second core casings 78, 80.

(44) The first outer core casing 78 extends radially inwardly in a downstream direction towards the engine centreline 9 in a part of the core 11 between the low pressure compressor 14 and the high pressure compressor 15 (e.g. in a diffuser section between the compressors 14, 15). The second outer core casing 80 on the other hand is relatively straight, and extends radially inwardly in a downstream direction to a lesser extent than the first core casing 78. As can be seen in the close up of FIG. 6A, downstream of the point at which the outer core wall 76 splits into the first and second outer core casings 78, 80, the first outer core casing extends radially inwardly in a downstream direction to a greater extent (i.e. at a steeper angle towards the engine centreline) than the second outer core casing 80. This results in an annular inter-casing gap 82 being defined by a radially outer surface of the first outer core casing 78 and a radially inner surface of the second outer core casing 80. In the embodiment being described, this arrangement of the first and second outer core casings 78, 80 may therefore provide narrowing of the core airflow path A without narrowing the outer surface of the outer core casing 76 to the same extent.

(45) In an alternative embodiment, as illustrated in FIG. 6B, the outer core casing 76 does not bifurcate such that first outer core casing 78 and second outer core casing 80 are not present. In this embodiment, the compressors 14, 15 are surrounded by a single casing formed by the outer core casing 76. In such embodiments, the single wall 76 may increase in width and/or change shape at a point along its axial length upstream of the low pressure compressor 14 and downstream of the high pressure compressor 15.

(46) First Flange Connection

(47) The first flange connection 60 forms a connection at one end region of the “intercase” 76b of the engine 10—i.e. a part of the outer core casing 76 between the casing 76a of the low pressure compressor 14 and the casing 76c of the high pressure compressor 15, as illustrated in FIG. 3B.

(48) In the embodiment being described, the first flange connection 60 comprises two flanges 60a, 60b that extend radially outward from adjacent portions of the outer core casing 76, and which extend circumferentially around the casing 76. The two flanges of the first flange connection 60 extend radially outward from the second outer core casing 80 in the embodiment shown in FIG. 6A, and radially outward from the single wall outer casing 76 in the embodiment shown in FIG. 6B. In alternative embodiments, the first flange connection 60 may comprise a single flange arranged to be connected to a connection block, hollow portion of the casing 76 or the likes, instead of to a second flange. The first flange connection 60 may therefore comprise one or more flanges.

(49) The intercase 76b may be arranged to be removable or detachable so as to allow access to the first and second compressors 14, 15.

(50) The first flange connection 60 is arranged to allow separation of the outer core casing 76 at the axial position of the first flange 60 connection, for example to facilitate access for servicing and maintenance—the first flange connection 60 therefore defines a separation point of the engine 10. Two portions 10a, 10b of the casing 76 of the engine 10 may be separated by disconnection of the first flange 60 connection (where portion 10a may correspond to the low pressure compressor casing 76a and the intercase 76b, and portion 10b to the high pressure compressor casing 76c, in the examples shown in FIGS. 3B and 6B).

(51) The first flange connection 60 comprises a two-part connection formed by a flange 60a and a respective connection structure 60b (i.e. another flange, bulkhead, or other structure) to which the flange 60a is connected. In the embodiment being described, the flange 60a of the first flange connection 60 is a flange extending from the intercase 76b, and the connection structure 60b is a flange extending from the casing 76c of the high pressure compressor 15. In the embodiment being described, the flange 60a of the first flange connection 60 is the rearmost flange of the intercase 76b; in alternative embodiments, a or the flange forming a part of the first flange connection 60 may be integral with the intercase 76 but not the rearmost flange of the intercase, may be integral with the casing 76a of the low pressure compressor (e.g. being the most downstream flange of the low pressure compressor casing 76a), or may be integral with the casing 76c of the high pressure compressor (e.g. being the most upstream flange of the high pressure compressor casing 76c).

(52) The axial position of the first flange connection 60 is defined as the axial position of the contact surface of the one or more flanges 60a, 60b from which it is formed. The axial position therefore corresponds to the axial position of the separation point formed by the first flange connection 60.

(53) For example, in one embodiment, the first flange connection 60 is formed by a pair of cooperating flanges 60a, 60b via which the two portions 10a, 10b are connected. An example of this is shown in FIGS. 3B and 6A and described in more detail later. In this embodiment, the axial position of the first flange connection 60 is defined as the axial position of the contact surface at which one of the pair of flanges is connected to, and in contact with, the other.

(54) In other embodiments, the first flange connection 60 comprises a single flange 60a that is connected to another structure such as a bulkhead, box-portion or similar structure. An example of this is shown in FIG. 6B and described in more detail later. In this embodiment, the axial position of the first flange connection 60 is defined as the axial position of the contact surface of the single flange 60a from which the first flange connection 60 is formed.

(55) In the embodiment being described, the first flange 60a of the first flange connection 60 forms part of a first engine casing portion 10a, and is connected to a second engine casing portion 10b by a flange connector 61.

(56) The two parts 60a,b of the first flange connection 60 are connected by a flange connector 61. In the embodiment being described the flange connector 61 comprises a plurality of bolts passing through the first flange 60a of the first flange connection 60 and into a second opposing flange 60b provided on the second engine casing portion 10b. In this embodiment, the first flange 60a comprises a plurality of holes therethrough arranged to receive the bolts 61, with corresponding holes provided in the second flange 60b. In alternative embodiments, one or more clamps, clips and/or fasteners may be used in addition to, or instead of, bolts 61. In such embodiments, the first and/or second flange 60a,b may not have holes therethrough. In other embodiments, the bolts may pass through holes provided in a single flange 60a forming the first flange connection 60 into a bulkhead or other structure to which the flange is connected.

(57) The first flange connection 60 is the first flange connection that is downstream of an axial position, X.sub.2, defined by the axial midpoint between the mid-span axial location, X.sub.1, on the trailing edge of the most downstream low pressure aerofoil of the low pressure compressor 14 (the first compressor 14) and the mid-span axial location, X.sub.3, on the leading edge of the most upstream high pressure aerofoil of the high pressure compressor 15 (the second compressor 15). I.e. it is the flange 60 connection closest to that axial midpoint, X.sub.2, in a downstream direction (as marked by arrow C in FIGS. 11A and 11B) from the axial midpoint, X.sub.2, the axial midpoint being the midpoint between the rear of the forwardmost compressor 14 and the front of the rearmost compressor 15 in the embodiments being described.

(58) The skilled person would appreciate that flange connection arrangements may vary in various embodiments. For example, in some embodiments the first flange connection 60 may be the first flange connection downstream of the first compressor 14, whereas in other embodiments an additional one or more flange connections 63 may be present between the first compressor 14 and the first flange 60 connection, and/or downstream (rearward) of the first flange 60 connection.

(59) In various embodiments, the additional flange connections 63 may be located anywhere along the length of a casing of the (first) low pressure compressor 14. In some embodiments, the additional flange connection 63 is located downstream of the first compressor 14. In some engine designs, for example, presence of a core mount 53b connecting the core 11 to the pylon and torque box may necessitate a joint in the core casing at the start of the torque box support structure (rearward of the first compressor 14). There may be no barrel-shaped casing extending along the length of the first compressor 14 to meet a different compressor casing and/or forward support structure.

(60) In other embodiments, the additional flange connection 63 may be located at a position along the axial length of first compressor 14. In some engine designs, for example, where the only mount(s) provided may be to the nacelle 21 rather than to the core 11, no torque box or torque panel may be provided within the engine core 11—in such embodiments, the casing may extend further forward—for example to half way along the length of the first compressor 14.

(61) In some embodiments, the additional flange connection 63 may not be present. In such an embodiment, the low pressure compressor casing 76a and the intercase 76b may form a single casing rather than being split into separate sections. The low pressure compressor casing 76b then extends up to, and is connected to, the high pressure compressor casing 76c (e.g. via the first flange connection 60).

(62) In alternative embodiments, such as that shown in FIG. 11A, the first flange connection 60 is axially upstream of a leading edge of a first (or most upstream) aerofoil of the second compressor 15.

(63) In some embodiments, such as that shown in FIG. 11B, the first flange connection 60 is axially downstream of a leading edge of a first aerofoil of the second compressor 15. In the embodiment of FIG. 11B, the first flange connection 60 has an axial position part way along the second compressor 15. In alternative embodiments, the first flange connection 60 may be axially aligned to the mid-span leading edge of the most upstream high pressure aerofoil of the high pressure compressor 15.

(64) In FIGS. 11A and 11B, the span 103 between X.sub.1 and X.sub.3 is marked. However, this span only serves to define the axial position X.sub.2, and does not limit the position of the flange connection 60—the flange connection 60 may be anywhere downstream of the axial position X.sub.2, as illustrated by arrow C.

(65) In the embodiment being described, the intercase 76b comprises two flanges—a forward flange nearer the first compressor 14 and a rearward flange 60a nearer the second compressor 15. The two flanges may each form a part of a different flange connection, and may allow an intercase portion 76b of casing 76 to be lifted away to facilitate access to the compressors 14, 15. In the embodiment being described, the rearward flange 60a of the intercase forms part of the first flange connection 60 (as the forward flange lies forward of the axial midpoint X.sub.2). In alternative embodiments, the intercase 76b may be divided into two or more portions, and/or a larger number of flanges may be present—the first flange 60a may therefore not be the rearward, or rearmost, flange of the intercase portion 76b in all embodiments.

(66) In the embodiment illustrated in FIG. 6A the first flange connection 60 is provided in the second outer core casing 80. In this embodiment, the second outer core casing 80 is separated into two portions at the separation point formed by the first flange connection 60. A first flange 60a forming the first flange connection 60 is provided on the downstream of those portions. An opposing second flange 60b is provided on the other portion of the second outer core casing 80 with which the first flange is coupled via the flange connector 61. In other embodiments, any other suitable structure may be provided to provide a connecting point for the flange connector 61.

(67) In the embodiment of FIG. 6B, the first flange connection 60 is provided in the outer core casing 76. In this case, the first and second outer core casings are not provided at the axial position of the first flange connection 60—the outer core casing 76 instead comprises a single wall. In this embodiment, the first flange connection 60 comprises a single flange 60a that is arranged to couple to an adjacent portion of the outer core casing 76, in this embodiment a box-type portion of the outer casing (which may be described as the outer core casing 76 separating into first and second outer core casings over a relatively short axial length of the engine 10). In this embodiment, no second, opposing flange is provided—the flange connector 61 connects the flange 60a of the first flange connection 60 directly to the opposing casing surface.

(68) In the embodiment being described, the opposing casing surface comprises threaded holes arranged to align with threaded holes in the flange 60a; bolts 61 may then be used to join the flange 60 to the opposing casing surface.

(69) First Flange Radius

(70) The first flange radius 104 is the radial distance between the engine centre line 9 and the flange connector 61. In the embodiment being described, the flange connector 61 comprises a plurality of bolts, and the first flange radius 104 is defined as the distance between the engine centreline 9 and a centreline of each bolt (the bolts being oriented axially and located at the same radial distance from the engine centreline 9).

(71) The skilled person would appreciate that the flange connector location (i.e. bolt location in the embodiment being described) affects stress and strain distribution and may therefore be a more relevant parameter than the location of the radially outer edge of the first flange connection 60.

(72) An increase in first flange radius 104 therefore corresponds to moving the first flange connection 60 further from the engine centreline 9, and/or moving the flange connector 61 further up the flange provided in the first flange connection 60 (e.g. by providing bolt holes at a higher radius).

(73) In the embodiments being described, the first flange radius 104 is in the range of 15 cm to 90 cm, and more particularly in the range from 25 cm to 60 cm, for example from 30 cm to 55 cm.

(74) Gas Path Radius and Gas Path Ratio

(75) Referring to FIG. 7, a gas path radius 102 is defined as the outer radius of the core gas flow path A at the axial position of the first flange connection 60. The gas path radius is measured in the same plane as the first flange radius 104, and is measured from the engine centreline 9. In the described embodiment, the gas path radius 102 is defined as the radius of the radially inner surface of the first outer core casing 78 which defines the core gas flow path A measured from the engine centreline 9. In other embodiments, the gas path radius may be measured to the radially inner surface of the outer core casing 76 which defines the core has flow path A (e.g. in embodiments where the outer core casing 76 is not bifurcated into the first and second outer core casings 78, 80 at the position of the first flange connection 60).

(76) A gas path ratio is defined as:

(77) $\frac{\mathrm{first}\mathrm{flange}\mathrm{radius}\left(104\right)}{\mathrm{gas}\mathrm{path}\mathrm{radius}\left(102\right)}$

(78) In the embodiment being described, the gas turbine engine 10 is configured such that the gas path ratio is equal to or greater than 1.10, and more particularly equal to or greater than 1.50. In both cases, the gas path ratio may be less than 2.0. It may therefore be in an inclusive range between 1.10 and 2.0 or in an inclusive range between 1.50 and 2.0.

(79) The radial positioning of the first flange connection 60 relative to the radius of the gas flow path may contribute to reducing or minimising engine bending whilst maintaining flange integrity. By configuring the gas turbine engine 10 so that the gas path ratio is within the range above the appropriate stiffness may be provided to the engine core 11.

(80) The gas path ratio may be equal to or greater than 1.10 for a medium sized engine (i.e. fan diameter 112 greater than 240 cm). The gas path ratio may be equal to or greater than 1.50 for a large sized engine (i.e. fan diameter 112 greater than 300 cm). These values may however be associated with other fan sizes.

(81) In various embodiments, the gas path ratio may have a value of 1.10, 1.15. 1.20, 1.25, 1.30, 1.35, 1.40, 1.45, 1.50, 1.55, 1.60, 1.65, 1.70, 1.75, 1.80, 1.85, 1.90, 1.95 and 2.00. The gas path ratio may be, for example, between any two of the values in the previous sentence.

(82) Fan Diameter Ratio

(83) As already described elsewhere herein, the gas turbine engine 10 comprises a fan 23 located upstream of the engine core 11. The fan 23 comprises a plurality of rotor blades 23a, also referred to as fan blades 23a, one of which is shown in FIG. 5. The plurality of rotor blades form a rotor blade set in an annular array around a central hub.

(84) A fan diameter ratio is defined as:

(85) $\frac{\mathrm{first}\mathrm{flange}\mathrm{radius}\left(104\right)}{\mathrm{fan}\mathrm{diameter}\left(112\right)}$

(86) In the embodiment being described, the gas turbine engine is configured such that the fan diameter ratio is equal to or greater than 0.125, and more particularly less than or equal to 0.17. It may therefore be in an inclusive range between 0.125 and 0.17.

(87) The fan diameter is equal to twice the radius 101 of the fan 23. In the embodiment being described, the fan diameter is greater than 240 cm, and more particularly greater than 300 cm (in both cases it may be no more than a maximum of 380 cm). In the embodiment being described, the fan diameter is between 330 cm and 380 cm, and more particularly between 335 cm and 360 cm.

(88) The radial positioning of the first flange connection 60 relative to the fan 23 contributes to reducing or minimising engine bending whilst maintaining flange integrity. By configuring the gas turbine engine 10 so that the fan diameter ratio is within the range above the appropriate stiffness may be provided to the engine core 11.

(89) In various embodiments, the fan diameter ratio may have a value of 0.125, 0.130, 0.135, 0.140, 0.145, 0.150, 0.155, 0.160, 0.165 and 0.170. The fan diameter ratio may be, for example, between any two of the values in the previous sentence.

(90) Fan Blade Mass and Blade Set Ratio

(91) A fan blade mass ratio is defined as:

(92) $\frac{\mathrm{first}\mathrm{flange}\mathrm{radius}\left(104\right)}{\mathrm{mass}\mathrm{of}\mathrm{each}\mathrm{fan}\mathrm{blade}}$

(93) The fan blade mass ratio relates the mass of each fan blade 23a provided on the fan 23 to the first flange radius 104. The skilled person would appreciate that each fan blade 23a generally has the same mass, within manufacturing tolerances. If the mass of each fan blade differs significantly, a fan blade mass ratio for each fan blade may be determined separately and configured to fall within the ranges defined herein. In the embodiments being described, the gas turbine engine 10 is configured such that the fan blade mass ratio is equal to or less than 19.0 mm/pound (41.9 mm/kg). More particularly, the fan blade mass ratio is equal to or greater than 5 mm/pound (11 mm/kg) (or 5.0 mm/pound (11.0 mm/kg)). It may therefore be in an inclusive range between 19.0 mm/pound (41.9 mm/kg) and 5.0 mm/pound (11.0 mm/kg). The mass of each fan blade may be in a range between 20 lb (9 kg) and 70 lb (32 kg).

(94) In various embodiments, fan blade mass ratio may have a value of 5.0 mm/lb (11.0 mm/kg), 6.0 mm/lb (13.2 mm/kg), 7.0 mm/lb (15.4 mm/kg), 8.0 mm/lb (17.6 mm/kg), 9.0 mm/lb (19.8 mm/kg), 10.0 mm/lb (22.1 mm/kg), 11.0 mm/lb (24.3 mm/kg), 12.0 mm/lb (26.5 mm/kg), 13.0 mm/lb (28.7 mm/kg), 14.0 mm/lb (30.9 mm/kg), 15.0 mm/lb (33.1 mm/kg), 16.0 mm/lb (35.3 mm/kg), 17.0 mm/lb (37.5 mm/kg), 18.0 mm/lb (39.7 mm/kg) and 19.0 mm/lb (41.9 mm/kg). The blade set mass ratio may be, for example, between any two of the values in the previous sentence.

(95) The radial positioning of the first flange connection 60 (as determined by the first flange radius 104) and the fan blade mass may also contribute to minimising engine bending whilst maintaining flange integrity. By configuring the gas turbine engine 10 so that the fan blade mass ratio is within the range above the appropriate stiffness may be provided to the engine core 11.

(96) A blade set mass ratio is defined as

(97) $\frac{\mathrm{first}\mathrm{flange}\mathrm{radius}\left(104\right)}{\mathrm{total}\mathrm{mass}\mathrm{of}\mathrm{the}\mathrm{plurality}\mathrm{of}\mathrm{fan}\mathrm{blades}}$

(98) The blade set mass ratio relates the total mass of the plurality of fan blades 23a forming the fan 23 (i.e. the blade set) and the first flange radius (104). In the embodiments being described, the blade set ratio is the inclusive range between 0.95 mm/pound (2.09 mm/kg) and 0.35 mm/pound (0.77 mm/kg).

(99) In various embodiments, the blade set mass ratio may have a value of 0.35 mm/lb (0.77 mm/kg), 0.40 mm/lb (0.88 mm/kg), 0.45 mm/lb (0.99 mm/kg), 0.50 mm/lb (1.10 mm/kg), 0.55 mm/lb (1.21 mm/kg), 0.60 mm/lb (1.32 mm/kg), 0.65 mm/lb (1.43 mm/kg), 0.70 mm/lb (1.54 mm/kg), 0.75 mm/lb (1.65 mm/kg), 0.80 mm/lb (1.76 mm/kg), 0.85 mm/lb (1.87 mm/kg), 0.90 mm/lb (1.98 mm/kg) and 0.95 mm/lb (2.09 mm/kg). The blade set mass ratio may be, for example, between any two of the values in the previous sentence.

(100) As discussed elsewhere herein, each of the fan blades 23a is at least partly formed from a metallic material. The metallic material may be titanium based metal or an aluminium based material such as aluminium lithium alloy.

(101) In other embodiments, each of the fan blades 23a may be at least partly formed from a composite material. The composite material may be, for example, a metal matrix composite and/or an organic matrix composite, such as carbon fibre.

(102) Fan Outlet Guide Vane

(103) A fan outlet guide vane (OGV) 58 is provided that extends radially across the bypass duct 22, between an outer surface of the engine core 11 (e.g. the outer core casing 76) and the inner surface of the nacelle 21.

(104) The fan outlet guide vane 58 connects the engine core 11 to the nacelle 21. The fan OGV 58 may additionally remove or reduce the swirl from the flow coming from the fan 23.

(105) The fan OGV 58 extends between a radially inner edge 58a (adjacent the engine core 11) and a radially outer edge 58b (adjacent the nacelle 21) and has a leading (or upstream) edge and a trailing (or downstream) edge relative to the direction of gas flow B through the bypass duct 22.

(106) An axial position of the radially inner edge 58a of the OGV 58 is defined at the axial mid-point of the radially inner edge 58a. This may be referred to as the inner axial centrepoint of the OGV 58, or the root centrepoint of the OGV 58.

(107) An axial position of the radially outer edge 58b of the OGV 58 is defined at the axial mid-point of the radially inner outer edge 58b. This may be referred to as the outer axial centrepoint of the OGV 58, or the tip centrepoint of the OGV 58.

(108) The axial distance 108 between the root centrepoint of the OGV 58a and the first flange connection 60 is defined as the distance along the axis 9 between the axial position of the root centrepoint 58a of the OGV 58 and the axial position of the axial centre point of the first flange connection 60. The axial distance 108 between the root centrepoint of the OGV 58a and the first flange connection 60 is less than or equal to 135 cm, and more particularly in the range of 30 cm to 130 cm in the embodiment being described. More particularly, it may be in the range of 30 cm to 105 cm, more specifically in the range of 50 cm to 105 cm.

(109) The axial distance 110 between the tip centrepoint 58b of the OGV 58 and the first flange connection 60 is defined as the distance along the axis 9 between the axial position of the tip centrepoint of the OGV 58b and the axial position of the axial centre point of the first flange connection 60. The axial distance 110 between the root centrepoint of the OGV 58a and the first flange connection 60 is less than or equal to 90 cm, and more particularly in the range of 20 cm to 90 cm in the embodiment being described. Yet more particularly, it may be in the range of 40 cm to 90 cm.

(110) The axial positioning of the fan outlet guide vanes (fan OGVs) 58 may have an effect in reducing or minimising engine bending whilst maintaining flange integrity.

(111) In particular, the engine 10 may be designed such that the axial distance 108 between the fan OGV root centrepoint 58a and the first flange connection 60 is relatively short. A ratio of the axial distance 108 between the fan OGV root centrepoint 58a and the first flange connection 60 centre to the first flange radius 104 of 2.6 or less may provide an appropriate stiffness for the engine core 11—this ratio may be referred to as a fan OGV root position ratio, and may be represented as:

(112) $\frac{\begin{array}{c}\mathrm{axial}\mathrm{distance}\left(108\right)\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\left(60\right)\mathrm{and}\mathrm{the}\mathrm{fan}\mathrm{OGV}\mathrm{root}\mathrm{centrepoint}\left(58a\right)\end{array}}{\mathrm{first}\mathrm{flange}\mathrm{radius}\left(104\right)}$

(113) In the embodiment being described, the engine 10 is configured such that the fan OGV root position ratio has a value of less than or equal to 2.6, and more particularly between 2.6 and 0.8 (inclusive).

(114) In various embodiments, the fan OGV root position ratio may have a value of 2.6, 2.5, 2.4, 2.2, 2.0, 1.8, 1.6, 1.5, 1.4, 1.2, 1.0, or 0.8. The fan OGV root position ratio may be, for example, between any two of the values in the previous sentence.

(115) In some embodiments, a fan OGV root position to fan diameter ratio of:

(116) $\frac{\begin{array}{c}\mathrm{axial}\mathrm{distance}\left(108\right)\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\left(60\right)\mathrm{and}\mathrm{the}\mathrm{fan}\mathrm{OGV}\mathrm{root}\mathrm{centrepoint}\left(58a\right)\end{array}}{\mathrm{the}\mathrm{fan}\mathrm{diameter}}$

(117) is less than or equal to 0.33. The fan diameter is equal to twice the radius 101 of the fan 23. In the embodiment being described, the fan diameter is greater than 240 cm, and more particularly greater than 300 cm (in both cases it may be no more than a maximum of 380 cm). In the embodiment being described, the fan diameter is between 330 cm and 380 cm, and more particularly between 335 cm and 360 cm.

(118) In the embodiment being described, the engine 10 is configured such that the fan OGV root position to fan diameter ratio is greater than or equal to 0.12.

(119) In various embodiments, the fan OGV root position to fan diameter ratio may have a value of 0.33, 0.32, 0.30, 0.27, 0.25, 0.22, 0.20, 0.17, 0.15, or 0.12. The fan OGV root position to fan diameter ratio may be, for example, between any two of the values in the previous sentence.

(120) In some embodiments, the fan OGV root position to fan diameter ratio may take a value, or fall in a range, as listed above whilst the fan OGV root position ratio may not take a value, or fall in a range, as listed above, or vice versa. In other embodiments, both fan OGV root position ratios may take a value, or fall in a range, as listed above.

(121) Additionally or alternatively, the engine 10 may be designed such that the axial distance 110 between the fan OGV tip centrepoint 58b and the first flange connection 60 is relatively short. A ratio of the axial distance 110 between the fan OGV tip centrepoint 58b and the first flange connection 60 centre to the first flange radius 104 of 1.8 or less may provide an appropriate stiffness for the engine core 11—this ratio may be referred to as a fan OGV tip position ratio, and may be represented as:

(122) $\frac{\begin{array}{c}\mathrm{axial}\mathrm{distance}\left(110\right)\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\left(60\right)\mathrm{and}\mathrm{the}\mathrm{fan}\mathrm{OGV}\mathrm{tip}\mathrm{centrepoint}\left(58b\right)\end{array}}{\mathrm{first}\mathrm{flange}\mathrm{radius}\left(104\right)}$

(123) In the embodiment being described, the engine 10 is configured such that the fan OGV tip position ratio has a value of less than or equal to 1.8, and more particularly between 1.8 and 0.6 (inclusive).

(124) In various embodiments, the fan OGV tip position ratio may have a value of 1.8, 1.7, 1.6, 1.5, 1.4, 1.3, 1.2, 1.1, 1.0, 0.9, 0.8, 0.7, or 0.6. The fan OGV tip position ratio may be, for example, between any two of the values in the previous sentence.

(125) In some embodiments, a fan OGV tip position to fan diameter ratio of:

(126) $\frac{\begin{array}{c}\mathrm{axial}\mathrm{distance}\left(110\right)\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\left(60\right)\mathrm{and}\mathrm{the}\mathrm{fan}\mathrm{OGV}\mathrm{tip}\mathrm{centrepoint}\left(58b\right)\end{array}}{\mathrm{the}\mathrm{fan}\mathrm{diameter}}$

(127) is less than or equal to 0.22. The fan diameter is equal to twice the radius 101 of the fan 23. In the embodiment being described, the fan diameter is greater than 240 cm, and more particularly greater than 300 cm (in both cases it may be no more than a maximum of 380 cm). In the embodiment being described, the fan diameter is between 330 cm and 380 cm, and more particularly between 335 cm and 360 cm.

(128) In the embodiment being described, the engine 10 is configured such that the fan OGV tip position to fan diameter ratio is greater than or equal to 0.095.

(129) In various embodiments, the fan OGV tip position to fan diameter ratio may have a value of 0.22, 0.21, 0.20, 0.19, 0.18, 0.17, 0.16, 0.15, 0.14, 0.13, 0.12, 0.11, 0.10 or 0.095. The fan OGV tip position to fan diameter ratio may be, for example, between any two of the values in the previous sentence.

(130) In some embodiments, the fan OGV tip position to fan diameter ratio may take a value, or fall in a range, as listed above whilst the fan OGV tip position ratio may not take a value, or fall in a range, as listed above, or vice versa. In other embodiments, both fan OGV tip position ratios may take a value, or fall in a range, as listed above.

(131) Front Mount

(132) The engine 10 is arranged to be mounted to a wing 52 of an aircraft 90 by means of one or more pylons 53 (a pylon may also be referred to as an airframe strut).

(133) In the embodiments being described with respect to FIG. 10, the engine 10 is arranged to be connected to a pylon 53 in a minimum of two places. In the embodiment being described, the two places comprise a nacelle mount 53a connecting the nacelle 21 to the pylon 53 and a core mount 53b connecting the core 11 to the pylon 53. The nacelle mount 53a is forward of the core mount 53b in this embodiment. The front mount 50 is therefore the nacelle mount 53a in the embodiment being described.

(134) In the embodiment shown in FIG. 7, the front mount 50 is a core mount, and two core mounts are provided. The front mount 50 is the front core mount 50.

(135) In some embodiments, the front mount 50 may be a nacelle mount 53a and may be located at the axial position of the fan OGV tip centrepoint 58b.

(136) In various embodiments, there may be only one core mount, or there may be multiple core mounts 53b—for example, the pylon 53 may be connected to the core 11 in multiple places, or multiple pylons 53 may each be connected to the core 11.

(137) In various embodiments, there may be only one nacelle mount 53a, or there may be multiple nacelle mounts 53a—for example, the pylon 53 may be connected to the nacelle 21 in multiple places, or multiple pylons 53 may each be connected to the nacelle 21.

(138) The forward-most mount 50, whether it is a nacelle mount 53a or a core mount 53b, is defined as the front mount 50.

(139) The axial distance 106 between the front mount 50 and the first flange connection 60 is defined as the distance along the axis 9 between the axial position of the axial centre point of the front mount 50 and the axial position of the axial centre point of the first flange connection 60.

(140) The skilled person would appreciate that the axial positioning of the front mount 50 may be important for reducing or minimising engine bending whilst maintaining flange integrity. In particular, the engine 10 may be designed such that the axial distance 106 between the front mount 50 and the first flange connection 60 is relatively short to increase stiffness (in particular increasing intercase stiffness). Keeping the distance 106 relatively short may also improve ease of assembly and core inspection. In the embodiments being described the first flange connection 60 is located at a point where the bending moment on the engine core 11 is quite high. The skilled person would appreciate that bending moment is generally higher nearer to the front mount 50. Increasing the first flange radius 104, so providing a larger diameter for the first flange connection 60, may facilitate reacting the relatively high bending moment.

(141) A ratio of the axial distance 106 between the front mount 50 and the first flange connection 60 centre to the first flange radius of 1.18 or less may provide an appropriate stiffness for the engine core 11—this ratio may be referred to as a front mount position ratio, and may be represented as:

(142) $\frac{\begin{array}{c}\mathrm{axial}\mathrm{distance}\left(106\right)\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\left(60\right)\mathrm{and}\mathrm{the}\mathrm{front}\mathrm{mount}\left(50\right)\end{array}}{\mathrm{first}\mathrm{flange}\mathrm{radius}\left(104\right)}$

(143) In the embodiment being described, the engine 10 is configured such that the front mount position ratio has a value of less than or equal to 1.18, and more particularly between 1.18 and 0.65.

(144) In various embodiments, the front mount position ratio may have a value of 1.18, 1.14, 1.10, 1.05, 1.00, 0.95, 0.90, 0.85, 0.80, 0.75, 0.70, 0.67, or 0.65. The front mount position ratio may be, for example, between any two of the values in the previous sentence.

(145) The axial distance 106 between the first flange connection 60 and the front mount 50 is between 30 cm and 75 cm in the embodiment being described, and more particularly around 30 cm.

(146) In some embodiments, a front mount position to fan diameter ratio of:

(147) 0$\frac{\begin{array}{c}\mathrm{the}\mathrm{axial}\mathrm{distance}\left(106\right)\mathrm{between}\mathrm{the}\mathrm{first}\mathrm{flange}\\ \mathrm{connection}\left(60\right)\mathrm{and}\mathrm{the}\mathrm{front}\mathrm{mount}\left(50\right)\end{array}}{\mathrm{the}\mathrm{fan}\mathrm{diameter}\left(112\right)}$

(148) is less than or equal to 0.145. The fan diameter 112 is equal to twice the radius 101 of the fan 23. In the embodiment being described, the fan diameter 112 is greater than 240 cm, and more particularly greater than 300 cm (in both cases it may be no more than a maximum of 380 cm). In the embodiment being described, the fan diameter 112 is between 330 cm and 380 cm, and more particularly between 335 cm and 360 cm.

(149) In the embodiment being described, the engine 10 is configured such that the front mount position to fan diameter ratio is greater than or equal to 0.07.

(150) In various embodiments, the front mount position to fan diameter ratio may have a value of 0.145, 0.140, 0.135, 0.130, 0.125, 0.120, 0.115, 0.110, 0.105, 0.100, 0.095, 0.090, 0.085, 0.080, 0.075, or 0.070. The front mount position to fan diameter ratio may be, for example, between any two of the values in the previous sentence.

(151) In some embodiments, the front mount position to fan diameter ratio may take a value, or fall in a range, as listed above whilst the front mount position ratio may not take a value, or fall in a range, as listed above, or vice versa. In other embodiments, both front mount position ratios may take a value, or fall in a range, as listed above.

(152) In the present disclosure, upstream and downstream are with respect to the air flow through the compressor system; and front and rear is with respect to the gas turbine engine, i.e. the fan being in the front and the turbine being in the rear of the engine.

(153) 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.