EFFICIENT GAS TURBINE ENGINE

Abstract

A highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

Claims

1. A method of operating an aircraft that includes at least one gas turbine engine, the gas turbine engine including: an engine core comprising: a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft, the gas turbine engine further comprising: a fan comprising a plurality of fan blades; and a gearbox that receives an input from the first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, wherein the method includes operating the aircraft to take-off such that during take-off the gas turbine engine exhibits a maximum thrust, and wherein, at the maximum thrust, the gas turbine engine is configured to exhibit a fan to core efficiency ratio (FC) that is in a range of from 1.9×10.sup.5 mkg.sup.−1sPa to 3.5×10.sup.5 mkg.sup.−1sPa, where FC is defined as $\mathrm{FC}=\left(\mathrm{Fan}\mathrm{Diameter}\right).Math.\frac{\sqrt{T0\mathrm{turb\_in}}}{\mathrm{CS}},$ in which (i) T0.sub.turb_in is a turbine entry temperature that is defined as a temperature (K) at an inlet to a most axially upstream turbine rotor; (ii) CS is a core size defined as $\mathrm{CS}={\mathrm{Wcomp}}_{\mathrm{in}}.Math.\frac{\sqrt{T0\mathrm{comp\_out}}}{P0\mathrm{comp\_out}}$ where: Wcomp.sub.in is a mass flow rate (kg/s) at entry to the engine core; T0comp_out is a stagnation temperature at an exit to the second compressor; and P0comp_out is a stagnation pressure at the exit to the second compressor.

2. The method according to claim 1, wherein at the maximum thrust, the gas turbine engine is configured so that FC is in a range of from 2.0×10.sup.5 mkg.sup.−1sPa to 3×10.sup.5 mkg.sup.−1sPa.

3. The method according claim 1, wherein the fan has a diameter that is in a range of from 225 cm to 400 cm.

4. The method according to claim 1, wherein: the second turbine comprises at least one ceramic matrix composite component.

5. The method according to claim 4, wherein a mass of ceramic matrix composite in the second turbine is in a range of from 2% to 15% of a total mass of the second turbine.

6. The method according to claim 4, wherein: the first turbine comprises at least one ceramic matrix composite component.

7. The method according to claim 1, wherein: the second turbine comprises at least one row of stator vanes; and a most axially upstream row of the at least one row of stator vanes is metallic.

8. The method according to claim 1, wherein: the second turbine comprises at least one row of rotor blades; and a most axially upstream row of the at least one row of rotor blades is metallic.

9. The method according to claim 1, wherein: the second turbine comprises (i) at least one row of rotor blades, and (ii) seal segments that radially surround a most axially upstream row of the at least one row of rotor blades; and the seal segments comprise a ceramic matrix composite.

10. The method according to claim 1, wherein: the second turbine comprises at least two rows of stator vanes; and a second most axially upstream row of the at least two rows of stator vanes comprises a ceramic matrix composite.

11. The method according to claim 1, wherein: the second turbine comprises at least two rows of rotor blades; and a second most axially upstream row of the at least two rows of rotor blades comprises a ceramic matrix composite.

12. The method according to claim 11, wherein: the second most axially upstream row of the at least two rows of rotor blades is radially surrounded by seal segments that comprise a ceramic matrix composite.

13. The method according to claim 1, wherein: the first turbine comprises at least one row of stator vanes; and an axially most upstream row of the at least one row of stator vanes in the first turbine comprises a ceramic matrix composite.

14. The method according to claim 1, wherein: the first turbine comprises at least one row of rotor blades; and an axially most upstream row of the at least one row of rotor blades in the first turbine comprises a ceramic matrix composite.

15. The method according to claim 1, wherein at the maximum thrust, the gas turbine engine is configured so that T0.sub.turb_in is in a range of from 1800K to 2100K.

16. The method according to claim 1, wherein at the maximum thrust, the gas turbine engine is configured so that T0.sub.turb_in is in a range of from 1950K to 2100K.

17. The method according to claim 1, wherein a gear reduction ratio of the gearbox is in a range of from 3.3 to 4.

18. The method according to claim 1, wherein at the maximum thrust, the gas turbine engine is configured so that FC is in the range of from 2.1×10.sup.5 mkg.sup.−1sPa to 2.5×10.sup.5 mkg.sup.−1sPa.

Description

[0186] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0187] FIG. 1 is a sectional side view of a gas turbine engine;

[0188] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0189] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0190] FIG. 4 is a schematic showing an enlarged view of an upstream portion of the turbine of the gas turbine engine.

[0191] 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 (which may be referred to herein as a first compressor 14), a high-pressure compressor 15 (which may be referred to herein as a second compressor), combustion equipment 16, a high-pressure turbine 17 (which may be referred to herein as a second turbine), a low pressure turbine 19 (which may be referred to herein as a first turbine) 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.

[0192] 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.

[0193] 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.

[0194] 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.

[0195] 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.

[0196] 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.

[0197] 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.

[0198] 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.

[0199] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0200] 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.

[0201] 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.

[0202] FIG. 4 shows a part of the turbine in greater detail. In particular, FIG. 4 shows a downstream portion of the combustor 16, the second (high pressure) turbine 17, and an upstream portion of the first (low pressure) turbine 19. The high pressure turbine 17 is connected to the second core shaft 27. The low pressure turbine 19 is connected to the first core shaft 26.

[0203] In the illustrated example, the high pressure turbine 17 comprises, in axial-flow series, a first (most axially upstream) stator vane row 171, a first (most axially upstream) rotor blade row 172, a second (second most axially upstream) stator vane row 173, and a second (second most axially upstream) rotor blade row 174.

[0204] The first rotor blade row 172 is connected to a rotor disc 177. The second rotor blade row 174 is connected to a rotor disc 178. The two rotor discs 177, 178 are rigidly connected together by a link member 179. At least one of the rotor discs (in the illustrated example the first rotor disc 177) is connected to the second core shaft 27 via an arm 271. Accordingly, in use, the second core shaft 27, rotor discs 177, 178 and rotor blades 172, 174 all rotate together, at the same rotational speed.

[0205] The gas turbine engine 10 also comprises seal segments 175 provided radially outside the first rotor blade row 172. The gas turbine engine 10 also comprises seal segments 176 provided radially outside the second rotor blade row 174. The seal segments 175, 176 form the radially outer flow boundary (which may be referred to as the radially outer annulus line) in the region of the respective rotor blade row 172, 174, for example over the axial extent of the tips of the rotor blades 172, 174. The seal segments 175, 176 may form a seal with the tips of the rotor blades to prevent—or at least restrict—flow passing over or past the tips of the rotor blades. The seal segments 175, 176 may be abradable by the rotor blades. Thus, for example, the seal segments 175, 176 may be abraded by the rotor blades in use so as to form an optimal seal therebetween. Each segment may form an annular segment or a frusto-conical segment.

[0206] In the illustrated example, the high pressure turbine 17 is a two-stage high pressure turbine, in that it comprises two stages of vanes and blades, each stage comprising a stator vane row followed by a rotor blade row. However, it will be appreciated that gas turbine engines 10 in accordance with the present disclosure may comprise a high pressure turbine with any number of stages, for example 1, 2, 3, 4, 5 or more than 5 stages of stator vanes and rotor blades.

[0207] The low pressure turbine 19 is provided downstream of the high pressure turbine 17. An axially most upstream row of stator vanes 191 in the low pressure turbine 19 is provided immediately downstream of the final row of rotor blades 174 of the high pressure turbine 17. An axially most upstream row of rotor blades 192 in the low pressure turbine 19 is provided immediately downstream of the axially most upstream row of stator vanes 191. The axially most upstream row of rotor blades 192 is connected to the first core shaft 26 via a rotor disc. In use, the rotor blades 192 of the low pressure turbine 19 drive the first core shaft 26, which in turn drives the low pressure compressor 14, and also drives—via a gearbox 30—the fan 23.

[0208] FIG. 4 only shows an upstream portion of the low pressure turbine 19. However, it will be appreciated that downstream of the illustrated portion there may be provided further rows of stator vanes and rotor blades. For example, the low pressure turbine 19 may comprise 1, 2, 3, 4, 5 or more than 5 stages of stator vanes and rotor blades. The axially most upstream row of rotor blades 192 are connected to one or more (not shown) downstream rotor blade rows via a linkage 199 that is connected to the disc 197 on which the rotor blades 192 are supported.

[0209] At least a part of the high pressure turbine 17 and/or the low pressure turbine 19 comprises a CMC in the illustrated example. Purely by way of example, the CMC material may be silicon carbide fibres and/or a silicon carbide matrix (SiC—SiC), although it will be appreciated that other CMCs may be used, such as an oxide-oxide (Ox-Ox CMC material), a monolithic ceramic, and/or the like.

[0210] As noted elsewhere herein, CMCs have different properties to conventional turbine materials, such as nickel alloys. For example, CMCs typically have lower density and are able to withstand higher temperatures than metals such as nickel alloys. The present inventors have understood that these properties can be useful in some areas of the turbine 17, 19, but other properties—such as lower thermal conductivity of CMCs compared to nickel alloys—mean that their use is not appropriate in all areas of the turbine 17, 19.

[0211] For example, depending on the type of engine (for example in terms of architecture and/or maximum thrust), any one or more of the first (most axially upstream) stator vane row 171, first (most axially upstream) rotor blade row 172, second (second most axially upstream) stator vane row 173, second (second most axially upstream) rotor blade row 174 and first or second set of seal segments 175, 176 of the high pressure turbine may be manufactured using CMCs. Components in the above list that are not manufactured using CMCs may be manufactured using a metal, such as a nickel alloy. Optionally, in any aspect or arrangement described and/or claimed herein and regardless of the number of stages in the high pressure turbine 17, the rotor blades of each stage in the high pressure turbine 17 may be surrounded by seal segments, and the seal segments surrounding any one or more stage (for example all stages) may be made from a CMC.

[0212] Purely by way of non-limitative example, in the FIG. 4 arrangement, the second stator vane row 173, second rotor blade row 174 and first set of seal segments 175 and second set of seal segments 176 of the high pressure turbine are manufactured using CMCs, whereas the first stator vane row 171 and the first rotor blade row 172 are manufactured using a nickel alloy. In this particular example, the temperature experienced by the first stator vane row 171 and the first rotor blade row 172 may be even higher than that which can be tolerated by CMCs. Accordingly, for this particular example, this means that the first stator vane row 171 and the first rotor blade row 172—which experience higher temperatures than downstream components due to their proximity to the combustor exit 16—can take advantage of the relatively high thermal conductivity of the nickel alloy so as to be cooled more effectively using cooling air (taken from the compressor, for example) which may be provided to passages running through the components.

[0213] The total mass of the high pressure turbine 17 may include the masses of the stator vanes 171, 173, rotor blades 172, 174, seal segments 175, 176, rotor discs 177, 178, one or more radially inner casing elements that form the inner flow boundary 220 over the axial extent of the high pressure turbine 17, and one or more radially outer casing elements that form the outer flow boundary 230 over the axial extent of the high pressure turbine 17.

[0214] CMCs may be used in appropriate parts of the low pressure turbine 19, although in some engines 10 their use in the low pressure turbine 19 may not be appropriate, and thus they may not be used. Purely by way of non-limitative example, in the FIG. 4 arrangement, the axially most upstream row of stator vanes 191 is manufactured using a CMC, whereas the axially most upstream row of rotor blades 192 is manufactured using a metal alloy (such as a nickel alloy). In this particular example, the temperature experienced by the axially most upstream row of rotor blades 192 may not be sufficiently high to benefit from the use of CMCs, although it will be appreciated that in other engines 10 in accordance with the present disclosure, the axially most upstream row of rotor blades 192 and/or the associated seal segments 193 may be manufactured using CMCs. Indeed, in some engines, components (such as vanes, blades and seals) downstream of the axially most upstream row of rotor blades 192 in the low pressure turbine 19 may be manufactured using CMCs.

[0215] Any component manufactured using CMCs may also be provided with an environmental barrier coating (EBC). Such an EBC may cover at least the gas washed surface of such components. Such an EBC may protect the CMC from environmental deterioration, for example deterioration due to the effects of water vapour. Such an EBC may be, for example ytterbium disilicate (Yb.sub.2Si.sub.2O.sub.7), which may be applied by any suitable method, such as air plasma spray.

[0216] As noted elsewhere herein, CMCs have a higher temperature capability than conventional materials, such as metal alloys. This means that selective use of CMCs in the turbine can mean that some components that would need to be cooled if they were to be made from a metal alloy do not need to be cooled because they are made from a CMC and/or some components manufactured using a CMC require less cooling than if they were to be made from a metal alloy. Additionally or alternatively, through use of CMCs it may be possible to expose some components to a higher temperature than would otherwise be possible.

[0217] Purely by way of non-limitative example, optimizing the use of CMCs in the engine (for example in one or more components of the turbine 17, 19 as described herein) may reduce the cooling flow C requirement, which may result in a more efficient engine core (because less flow is bypassing the combustor), meaning that for a given amount of core power, the mass flow entering the core can be reduced and/or the size and/or mass of the turbine(s) 17, 19 can be reduced.

[0218] FIGS. 1 and 4 schematically show turbine cooling apparatus 50. The turbine cooling apparatus extracts cooling flow C from the compressor 14, 15. The cooling flow C bypasses the combustor 16. The cooling flow C is then delivered to the high pressure turbine 17 and optionally the low pressure turbine 19. Although the turbine cooling apparatus 50 is shown in FIGS. 1 and 4 as extracting cooling flow C from a specific position in the high pressure compressor 15 and delivering it to a specific position in the high pressure turbine 17, it will be appreciated that this is merely for ease of schematic representation, and that the cooling flow C may be extracted from any suitable locations (for example multiple locations in the high pressure compressor 15 and/or the low pressure compressor 14) and delivered to any desired locations (for example multiple locations in the high pressure turbine 17 and/or the low pressure turbine 19).

[0219] A reduction in the amount of cooling flow C is desirable, because the cooling flow is not combusted and thus the amount of work that can be extracted from it is lower than for the flow that passes through the combustor 16. With reference to FIG. 1, the gas turbine engine 10 has a bypass ratio defined as the mass flow rate of the flow B through the bypass duct 22 divided by the mass flow rate of the flow A entering the engine core at cruise conditions. As the bypass ratio is increased—for example to increase engine efficiency—proportionally less flow A goes through the core. This means that for a given size of engine and/or to be able to withstand a given turbine entry temperature, a higher proportion of the core flow A may be required to be used as turbine cooling flow C. In this regard, as used herein, turbine entry temperature (T0turb_in) may be the maximum stagnation temperature measured at a position 60 in the gas flow path that is immediately upstream of the most axially upstream rotor blade row 172, i.e. at a so-called “red-line” operating condition of the engine at which the engine is certified.

[0220] A cooling to bypass efficiency ratio may be defined as the ratio of the mass flow rate C of the turbine cooling flow to the mass flow rate B of the bypass flow at cruise conditions. Using an understanding of the constraints and/or technologies described by way of example herein, the cooling to bypass efficiency ratio may be optimized to be as described and/or claimed herein. Additionally or alternatively, the mass of the high pressure turbine 17 and/or the low pressure turbine 19 may be optimized (for example reduced) relative to a conventional engine. In turn, this may reduce the mass of the high pressure turbine 17 and/or the low pressure turbine 19 as a proportion of the overall mass of the gas turbine engine 10.

[0221] Using an understanding of the constraints and/or technologies described by way of example herein, the normalized thrust may be optimized. In this regard, the normalized thrust is defined as the maximum net thrust of the engine 10 at sea level divided by the total mass of the turbines 17, 19 in the engine 10. The illustrated example has a high pressure turbine 17 and a low pressure turbine 19, however, it will be appreciated that where further turbines are included in the engine the total turbine mass includes the mass of all turbines.

[0222] As noted elsewhere herein, the optimized use of CMCs may result in a reduced turbine cooling flow requirement. Additionally or alternatively, through use of CMCs it may be possible to expose some components to a higher temperature than would otherwise be possible. This may result in the ability to increase the turbine entry temperatures relative to engines 10 that do not include optimized use of CMCs. In this regard, it has been found that higher turbine entry temperatures are desirable from an engine efficiency perspective.

[0223] Using an understanding of the constraints and/or technologies described by way of example herein, the cooling efficiency ratio may be optimized. In this regard, the cooling efficiency ratio is defined as the ratio between the turbine entry temperature (as defined elsewhere herein) and the cooling flow requirement. The cooling flow requirement may be defined as the mass flow rate of the turbine cooling flow C divided by the mass flow rate of the flow A entering the engine core at cruise conditions.

[0224] A core size CS may be defined for the gas turbine engine 10 as:

[00005]$\mathrm{CS}={\mathrm{Wcomp}}_{\mathrm{in}}.Math.\frac{\sqrt{T0\mathrm{comp\_out}}}{P0\mathrm{comp\_out}}$ [0225] where:

[0226] Wcomp_in is the mass flow rate (kg/s) at entry to the engine core, i.e. the mass flow rate of the core flow A measured at position 70 in FIG. 1;

[0227] T0comp_out is the stagnation temperature (K) at exit to the compressor, i.e. at exit of the highest pressure compressor 15, indicated by position 80 in FIG. 1;

[0228] P0comp_out is the stagnation pressure (Pa) at exit to the compressor i.e. at exit of the highest pressure compressor 15, indicated by position 80 in FIG. 1.

[0229] Using an understanding of the constraints and/or technologies described by way of example herein may allow a thrust to core efficiency ratio TC and/or a fan to core efficiency ratio FC to be optimised to be in the ranges described and/or claimed herein, where the thrust to core efficiency ratio TC and the fan to core efficiency ratio FC are as defined below (with T0turb_in being the turbine entry temperature at position 60 shown in FIG. 4, as described above).

[00006]$\mathrm{TC}=\left(\mathrm{Max}\mathrm{Net}\mathrm{Thrust}\mathrm{at}\mathrm{Sea}\mathrm{Level}\right).Math.\frac{\sqrt{T0\mathrm{turb\_in}}}{\mathrm{CS}}\mathrm{FC}=\left(\mathrm{Fan}\mathrm{Diameter}\right).Math.\frac{\sqrt{T0\mathrm{turb\_in}}}{\mathrm{CS}}$

[0230] It will be appreciated that the understanding and/or technology described and/or claimed herein results in a particularly efficient gas turbine engine 10. For example, the understanding and/or technology described and/or claimed herein may provide a particularly efficient gas turbine engine 10 in which a fan 23 that is driven by a gearbox 30 is complemented by an optimized engine core.

[0231] 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 and aspects 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.