GAS TURBINE ENGINE OPERATING POINT

Abstract

A gas turbine engine has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The bypass flow rate at the optimum bypass efficiency is appreciably lower than the maximum bypass flow rate at the given conditions. This results in increased design flexibility and improved overall engine performance.

Claims

1. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a bypass duct defined radially outside the engine core and radially inside a nacelle, such that a proportion of the fan flow flows through the bypass duct as bypass flow, and a further proportion of the fan flow flows through the engine core as core flow; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein: the gas turbine engine is operable at a reference operating point at cruise conditions for the engine, the reference operating point defining a fan reference rotational speed; a bypass efficiency is defined as the efficiency of the fan compression of the bypass flow, the bypass efficiency being a function of the fan bypass inlet mass flow rate; and at the fan reference rotational speed and cruise conditions, the ratio of the fan bypass inlet mass flow rate that would give the peak bypass efficiency to the maximum possible fan bypass inlet mass flow rate is no greater than 0.96.

2. A gas turbine engine according to claim 1, wherein at the reference operating point, the gas turbine engine delivers the thrust required to maintain the cruise conditions.

3. A gas turbine engine according to claim 1, wherein a reference operating mass flow rate, defined as the fan bypass inlet mass flow rate at the reference operating point, is between the fan bypass inlet mass flow rate that would give the peak bypass efficiency and the maximum possible fan bypass inlet mass flow rate at the fan reference rotational speed and cruise conditions.

4. A gas turbine engine according to claim 1, wherein the fan bypass inlet mass flow rate at the reference operating point is at least 2%, optionally 2.5%, higher than the fan bypass inlet mass flow rate that would give the peak bypass efficiency at the fan reference rotational speed and cruise conditions.

5. A gas turbine engine according to claim 1, wherein: the bypass duct comprises a throat having a throat area that defines the minimum flow area through the bypass duct; and the fan bypass inlet mass flow rate at the reference operating point is determined by the throat area, and wherein, optionally: the throat area is greater than the throat area that would be required to give peak bypass efficiency at the fan reference rotational speed and cruise conditions; and/or further optionally the throat is a fixed-area.

6. A gas turbine engine according to claim 1, wherein at the fan reference rotational speed and cruise conditions, the ratio of the fan bypass inlet mass flow rate that would give the peak bypass efficiency to the maximum possible fan bypass inlet mass flow rate is no greater than 0.94.

7. A gas turbine engine according to claim 1, wherein the bypass efficiency at the reference operating point is within 0.5% of the peak bypass efficiency at the fan reference rotational speed and cruise conditions.

8. A gas turbine engine according to claim 1, wherein a quasi-non-dimensional mass flow rate Q is defined as: $Q=W.Math.\frac{\sqrt{T.Math..Math.0}}{P.Math..Math.0.Math.A_{\mathrm{fan}}}.$ where: W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of the air at the fan face in Kelvin; P0 is average stagnation pressure of the air at the fan face in Pa; A.sub.fan is the area of the fan face in m.sup.2, and at engine cruise conditions: Kgs.sup.1N.sup.1K.sup.1/2Q0.0.6 Kgs.sup.1N.sup.NK.sup.1/2.

9. A gas turbine engine according to claim 1, wherein a specific thrust is defined as net engine thrust divided by mass flow rate through the engine and, at cruise conditions, the specific thrust is in the range of from 70 Nkg.sup.1s to 110 Nkg.sup.1s, optionally 70 Nkg.sup.1s to 90 Nkg.sup.1s.

10. A gas turbine engine according to claim 1, wherein: a fan pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit to the mean total pressure of the flow at the fan inlet, is no greater than 1.5, optionally in the range of from 1.35 to 1.45, at the reference operating point; and/or a fan root pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core to the mean total pressure of the flow at the fan inlet, is no greater than 1.25 at the reference operating point, wherein, optionally, the ratio between the fan root pressure ratio to a fan tip pressure ratio at the reference operating point is no greater than 0.95, where the fan tip pressure ratio is defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct to the mean total pressure of the flow at the fan inlet.

11. A gas turbine engine according to claim 1, wherein the forward speed of the gas turbine engine at the cruise conditions is in the range of from Mn 0.75 to Mn 0.85, and, optionally, the forward speed of the gas turbine engine at the cruise conditions is Mn 0.8.

12. A gas turbine engine according to claim 1, wherein the cruise conditions correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 11582 m and a forward Mach Number of 0.8.

13. A gas turbine engine according to claim 1, wherein the cruise conditions correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 10668 m and a forward Mach Number of 0.85.

14. A fan for a gas turbine engine having an engine core and a bypass duct, the fan being designed to operate at a reference operating point at cruise conditions for the engine, the reference operating point defining a fan reference rotational speed, wherein: a bypass efficiency is defined as the efficiency of the fan compression of the flow that would subsequently flow through the bypass duct, the bypass efficiency being a function of the mass flow through the bypass duct; and at the fan reference rotational speed and cruise conditions, the ratio of the fan bypass inlet mass flow rate that would give the peak bypass efficiency to the maximum possible fan bypass inlet mass flow rate that could be supported by the fan is no greater than 0.96.

15. A gas turbine engine comprising: a fan according to claim 14 comprising the engine core and bypass duct, wherein the bypass duct has a throat having a throat area that defines the minimum flow area through the bypass duct; and at the fan reference rotational speed and cruise conditions, the mass flow rate through the bypass duct is dependent on the throat area, wherein, optionally, the throat area is fixed.

16. A gas turbine engine according to claim 15, wherein the throat area is between the area of a throat that would result in the peak bypass efficiency and the throat area that would result in the maximum possible fan bypass inlet mass flow rate at the fan reference rotational speed and cruise conditions.

17. A gas turbine engine according to claim 16, wherein the throat area results in a mass flow rate through the bypass duct at the reference operating point that is at least 2% higher than the mass flow rate through the bypass duct that would give the peak bypass efficiency at the fan reference rotational speed and cruise conditions.

18. A gas turbine engine according to claim 15 wherein the throat area gives a bypass efficiency at the reference operating point that is within 0.5% of the maximum possible bypass efficiency at the cruise conditions and fan reference rotational speed.

19. A method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a bypass duct defined radially outside the engine core and radially inside a nacelle, such that a proportion of the fan flow flows through the bypass duct as bypass flow, and a further proportion of the fan flow flows through the engine core as core flow; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein: a bypass efficiency is defined as the efficiency of the fan compression of the bypass flow, the bypass efficiency being a function of the fan bypass inlet mass flow rate, the method comprising: using the engine to propel an aircraft in a climb phase, a cruise phase, and a descent phase, the cruise phase being directly after the climb phase and directly before the descent phase, and covering all operation between the end of the climb phase and the start of the descent phase; and for at least 80% of the time that the gas turbine engine is operating in the cruise phase, the ratio of the fan bypass inlet mass flow rate that would give the peak bypass efficiency to the maximum possible fan bypass inlet mass flow rate is no greater than 0.96.

20. An aircraft comprising at least two gas turbine engines, each gas turbine engine being in accordance with claim 1, wherein at the cruise conditions, each gas turbine engine provides thrust equal to the thrust required to maintain the cruise Mach Number divided by the number of gas turbine engines attached to the aircraft.

Description

[0109] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0110] FIG. 1 is a sectional side view of a gas turbine engine;

[0111] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0112] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0113] FIG. 4 is a graph showing bypass flow against bypass efficiency for a gas turbine engine in accordance with an example of the present disclosure;

[0114] FIG. 5 is a graph showing bypass flow against bypass efficiency for a gas turbine engine in accordance with an example of the present disclosure; and

[0115] FIG. 6 is a graph showing bypass flow against bypass efficiency for a gas turbine engine in accordance with an example of the present disclosure.

[0116] 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 1, 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 a gearbox 30.

[0117] The bypass duct 22 has a throat 100 which is defined by the minimum flow area A.sub.N through the bypass duct 22. In use, for example at certain operating conditions such as cruise conditions, the flow through the bypass duct 22 may be choked at the throat 100. For a given set of conditions (for example cruise conditions and a fixed fan rotational speed) the mass flow rate through the bypass duct 22 and/or over the fan 23 may be determined at least in part (for example solely or substantially solely determined by) the area A.sub.N of the throat 100.

[0118] 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. A throttle 161 is provided to control the fuel supply to the combustor. The amount of fuel supplied is dependent on the throttle position. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the core exhaust nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The gearbox 30 is a reduction gearbox, and may be an epicyclic gearbox.

[0119] 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.

[0120] 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.

[0121] 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.

[0122] 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.

[0123] 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.

[0124] 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.

[0125] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0126] 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 exhaust nozzle 20, 18 meaning that the flow through the bypass duct 22 has its own exhaust that is separate to and radially outside the core engine exhaust 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 exhaust nozzle, which may be referred to as a mixed flow exhaust 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 comprising a gearbox, the disclosure may apply, for example, to any type of gas turbine engine. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

[0127] 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.

[0128] The gas turbine engine 10 is required to operate in a range of different conditions, or states. Such conditions may be dependent upon the point in a flight cycle of an aircraft to which the gas turbine engine 10 may be attached. For example, the atmospheric conditions may vary significantly between take off and cruise. By way of further example, the forward speed and/or thrust requirement to accelerate up to or maintain a given forward speed may vary during the flight cycle.

[0129] Typically, gas turbine engine operation may be described at various points in the flight cycle, such as: [0130] maximum take-offwhich describes the performance of the engine at runway conditions (which may be taken as 101.3 kPa and 30 deg C.), for example at maximum throttle; [0131] Top of climbwhich describes the performance of the engine at the transition between the climb phase of the aircraft cycle and the start of the cruise phase; and [0132] Cruisewhich describes the performance of the engine during the cruise phase. By way of example, the cruise conditions may be as described/defined elsewhere herein.

[0133] Typically, a gas turbine engine will be designed for optimum efficiency at cruise, because this is the condition at which the engine operates for the longest period of time, and thus the condition at which most fuel is burned. Furthermore, because the engine operating point can be known and controlled accurately at cruise (for example through an understanding of the environmental cruise conditions, fan rotational speed at cruise and/or control of the throttle position 161), conventional engines are simply designed such that the fan operates at peak efficiency at cruise, with little consideration given to the characteristics of the fan performance away from the cruise operating point. For example, gas turbine engine fans may typically be designed to provide peak efficiency at cruise by selecting a given bypass mass flow rate at cruise, which may be determined by selection of an appropriate bypass geometryfor example having an appropriate throat 100 and associated throat area A.sub.Nwith little consideration given to the fan efficiency away from that cruise point (for example, away from the selected cruise mass flow rate).

[0134] However, the gas turbine 10 according to the present disclosure has been developed using a novel approach that considers a wider range of engine and fan characteristics. The resulting gas turbine engine 10 has a fan 23 that exhibits different characteristics to those of conventional engines. In this regard, FIGS. 4 to 6 are graphs showing fan bypass inlet mass flow rate (i.e. mass flow that subsequently flows down the bypass duct 22) as a percentage of the maximum possible fan bypass inlet mass flow rate at cruise conditions for a given fan reference rotational speed (x-axis) vs fan bypass efficiency (y-axis, which may be referred to herein simply as bypass efficiency, as defined elsewhere herein). The three graphs (FIGS. 4 to 6) represent three different examples of gas turbine engines 10 in accordance with the present disclosure. In each case, the curves A, B, C show the relationship between the fan bypass inlet mass flow rate and the bypass efficiency when the respective engine is operating at cruise conditions and fan reference rotational speed (as defined elsewhere herein). References herein to fan bypass inlet mass flow rate may mean quasi non-dimensional mass flow rate given by:

[00005]$\stackrel{.}{m}.Math.\frac{\sqrt{T.Math..Math.0.Math.\mathrm{in}}}{P.Math..Math.0.Math.\mathrm{in}}$

[0135] Where:

[0136] {dot over (m)} is absolute mass flow rate

[00006]$\left(\frac{\mathrm{Kg}}{s}\right)$

[0137] T0in is the stagnation temperature at the inlet to the fan

[0138] P0in is the stagnation pressure at the inlet to the fan

[0139] At the cruise conditions and fan reference rotational speed, the fan bypass inlet mass flow rate may be changed by changing the area A.sub.N of the throat 100. Accordingly, for gas turbine engines 10 that have fixed geometry nozzles (i.e. fixed bypass duct 22 shapes) and thus a fixed area A.sub.N of the throat 100, the operating point of the fan 23 on the curves A, B, C at the cruise conditions is fixed by the selected bypass duct geometry, for example by the selected throat area A.sub.N.

[0140] In this regard, the solid vertical lines A1, B1, C1 represent the chosen fan bypass inlet mass flow rate of the respective gas turbine engine 10 at the chosen operating point of the fan 23 (and thus the gas turbine engine 10) at cruise. Thus, the point (X, Y, Z respectively) at which the vertical line A1, B1, C1 crosses the respective curve A, B, C indicates the operating point of the fan 23 (and thus the gas turbine engine 10) at the cruise conditions, which may be referred to herein as the reference operating point X, Y, Z. The reference operating point X, Y, Z may be determined by the throttle position 161 required at cruise conditions to generate the fan reference rotational speed and give a desired level of thrust.

[0141] The dashed vertical lines A2, B2, C2 indicate the maximum possible fan bypass inlet mass flow rate of a respective engine 10 at the cruise conditions and fan reference rotational speed using the fan 23. In other words, for a given fan 23, it is not possible for the fan bypass inlet mass flow rate to increase beyond the dashed vertical lines A2, B2, C2 at the cruise conditions and fan reference rotational speed regardless of the geometry of the bypass duct 22, for example regardless of the area A.sub.N of the throat 100.

[0142] The dot-chain vertical lines A3, B3, C3 indicate the fan bypass inlet mass flow rate of the respective engine 10 required to give the maximum bypass efficiency at the cruise conditions and fan reference rotational speed using the fan 23. In other words, for a given fan 23, it is not possible to improve the bypass efficiency above that corresponding to the mass flow rate at the dot-chain vertical lines A3, B3, C3 at the cruise conditions and fan reference rotational speed regardless of the geometry of the bypass duct 22, for example regardless of the area A.sub.N of the throat 100.

[0143] For conventional engines, the position of the dot chain lines A3, B3, C3 would be very close to the position of dashed lines A2, B2, C2. However, for fans 23 and gas turbine engines 10 comprising fans 23 in accordance with the present disclosure (such as the three corresponding to the graphs shown by way of example in FIGS. 4 to 6), at the fan reference rotational speed and cruise conditions, the ratio of the fan bypass inlet mass flow rate that would give the peak bypass efficiency (dot-chain lines A3, B3, C3) to the maximum possible fan bypass inlet mass flow rate (dashed vertical lines A2, B2, C2) is no greater than 0.96. The fan bypass inlet mass flow rate that would give the peak bypass efficiency may be said to be 96% or less of the maximum possible fan bypass inlet mass flow rate at the fan reference rotational speed and cruise conditions.

[0144] In other words, the dot chain lines A3, B3, C3 and the dashed lines A2, B2, C2 are further apart for fans 23 and gas turbine engines 10 according to the present disclosure than they are for conventional fans and gas turbine engines.

[0145] This increased separation of the mass flow rate at cruise conditions and fan reference rotational speed between the maximum possible and that required to give the maximum bypass efficiency opens up greater design freedom. Purely by way of example, it may allow considerable separation between the mass flow rate at the reference operating point (i.e. the point at which the engine 10 is designed to operate at cruise conditions, solid vertical lines A1, B1, C1) and the mass flow rate at the peak bypass efficiency (dot-chain lines A3, B3, C3). As noted elsewhere herein, in turn this may result in greater stall and/or flutter margin, which may allow wider design freedom in other areas, such as fan blade geometry and/or construction and/or mass. In a conventional engine, such separation would result in unacceptably low efficiency at the reference operating point (i.e. for the point at which the engine is operating at cruise).

[0146] In conventional engines, the solid vertical lines A1, B1, C1, the dashed vertical lines A2, B2, C2 and the dot-chain vertical lines A3, B3, C3 would typically all be much closer together, because the focus would conventionally be on maximizing the peak efficiency value, and then setting the reference operating point to coincide as closely as possible with that peak efficiency. Thus, in a conventional engine, the solid vertical lines A1, B1, C1 (reference operating mass flow rate) and the dot-chain vertical lines A3, B3, C3 (peak bypass efficiency) would be necessarily much closer together, and typically substantially overlapping.

[0147] The greater design freedom offered by the fans 23 and gas turbine engines 10 according to the present disclosure may allow throat areas A.sub.N of the bypass duct 22 to be selected that are different (for example larger or smaller) to the throat area that would be selected purely in order to maximize the bypass efficiency at the cruise conditions and fan reference rotational speed. For example, a given difference (for example given percentage increase) between the bypass mass flow rate resulting from a selected throat area A.sub.N of the bypass duct 22 and the bypass mass flow rate resulting from a throat area selected purely in order to maximize the bypass efficiency would result in a greater reduction in bypass efficiency for a conventional fan or engine than for fans 23 or engines 10 in accordance with the present disclosure.

[0148] The arrangements and advantages associated therewith (for example in terms of the increased separation between the maximum mass flow rate and the mass flow rate at the peak bypass efficiency) may be particularly effective for gas turbine engines 10 in which the fan 23 is linked to a turbine 19 via a gearbox 30. In such gas turbine engines 10, there may be greater opportunity to take advantage of the increased design freedom. Purely by way of example, the lower rotational speed and/or larger diameter of the fan for a given power of engine may result in different fan design challenges compared with engines that do not have a gearboxsuch as, for example, different flutter and/or stall and/or surge characteristicswhich may be addressed more effectively through the design freedom offered by gas turbine engines according the present disclosure.

[0149] A further example of a feature that may be better optimized for gas turbine engines 10 according to the present disclosure compared with conventional gas turbine engines is the intake region, for example the ratio between the intake length L and the fan diameter D. Referring to FIG. 1, the intake length L is defined as the axial distance between the leading edge of the intake and the leading edge of the tip of the fan blades, and the diameter D of the fan 23 is defined at the leading edge of the fan 23. Gas turbine engines 10 according to the present disclosure, such as that shown by way of example in FIG. 1, may have values of the ratio L/D as defined herein, for example less than or equal to 0.4. Without being bound by any particular theory, reducing the ratio of L/D may be more attractive for gas turbine engines 10 in accordance with the present disclosure because of extra operational stability afforded by the lower than conventional ratio of the mass flow rate through the bypass duct at the peak bypass efficiency to the maximum possible mass flow through the bypass duct. This may make the fan 23 more resilient to distortion in the inlet flow, thereby requiring a relatively shorter intake length L to cope with the range of flow distortions that may be experienced in use.

[0150] The gas turbine engine 10 shown in FIG. 1 and described herein may comprise any one or more of the features described and/or claimed herein. For example, where compatible, such a gas turbine engine 10 may have any one or more of the features or values described herein of: quasi-non-dimensional mass flow rate Q; specific thrust; maximum thrust, turbine entry temperature; overall pressure ratio; bypass ratio; fan diameter; fan rotational speed; fan hub to tip ratio; fan pressure ratio; fan root pressure ratio; ratio between the fan root pressure ratio to the fan tip pressure ratio; fan tip loading; number of fan blades; construction of fan blades; and/or gear ratio.

[0151] 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.