Aircraft engine operability

Abstract

A gas turbine engine has a cycle operability parameter β in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined range of cycle operability parameter β may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

Claims

1. A method of operating a gas turbine engine for an aircraft, the gas turbine engine having (i) an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, (ii) a fan located upstream of the engine core, the fan comprising a plurality of fan blades, an annular fan face being defined at a leading edge of the fan, and (iii) 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, the gearbox having a reduction ratio GR, the method comprising: operating the gas turbine engine to propel the aircraft at a forward Mach number of 0.8, at an atmospheric pressure of 23000 Pa and an atmospheric temperature of −55 deg. C., such that a cycle operability parameter β satisfies the following:
1.2K.sup.−1/2≤β≤2.0K.sup.−1/2 where: a quasi-non-dimensional mass flow rate Q is defined as: $Q=W\frac{\sqrt{T0}}{P0.Math.A_{\mathrm{fan}}}$ W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of 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 an area of the fan face in m.sup.2; a specific thrust ST is defined as net engine thrust (N) divided by mass flow rate (Kg/s) through the engine; and the cycle operability parameter β is defined as: $\beta =\frac{\mathrm{GR}}{Q.Math.\mathrm{ST}}.$

2. The method of claim 1, further comprising: operating the gas turbine engine to propel the aircraft at the forward Mach number of 0.8, at the atmospheric pressure of 23000 Pa and the atmospheric temperature of −55 deg. C., such that the cycle operability parameter β satisfies the following:
1.3K.sup.−1/2≤β≤2.0K.sup.−1/2.

3. The method of claim 1, further comprising: operating the gas turbine engine to propel the aircraft at the forward Mach number of 0.8, at the atmospheric pressure of 23000 Pa and the atmospheric temperature of −55 deg. C., such that the cycle operability parameter β satisfies the following:
1.4K.sup.−1/2≤β≤2.0K.sup.−1/2.

4. The method of claim 1, further comprising: operating the gas turbine engine to propel the aircraft at the forward Mach number of 0.8, at the atmospheric pressure of 23000 Pa and the atmospheric temperature of −55 deg. C., such that the quasi-non-dimensional mass flow rate Q satisfies the following:
0.029Kgs.sup.−1N.sup.−1K.sup.1/2≤Q≤0.036Kgs.sup.−1N.sup.−1K.sup.1/2.

5. The method of claim 1, further comprising: operating the gas turbine engine to propel the aircraft at the forward Mach number of 0.8, at the atmospheric pressure of 23000 Pa and the atmospheric temperature of −55 deg. C., such that the specific thrust ST satisfies the following:
70Nkg.sup.−1s≤ST≤110Nkg.sup.−1s.

6. The method of claim 1, wherein: a fan tip loading is defined as dH/Utip.sup.2, where dH is an enthalpy rise across the fan and Utip is a translational velocity of the fan blades at a tip of the leading edge; and the method further comprises operating the gas turbine engine to propel the aircraft at the forward Mach number of 0.8, at the atmospheric pressure of 23000 Pa and the atmospheric temperature of −55 deg. C., such that that the fan tip loading dH/Utip.sup.2 satisfies the following:
0.28Jkg.sup.−1K.sup.−1/(ms.sup.−1).sup.2<dH/Utip.sup.2<0.36Jkg.sup.−1K.sup.−1/(ms.sup.−1).sup.2.

7. The method of claim 1, wherein: a fan pressure ratio, defined as a ratio of a mean total pressure of a flow at a fan exit to a mean total pressure of a flow at a fan inlet; and the method further comprises operating the gas turbine engine to propel the aircraft at the forward Mach number of 0.8, at the atmospheric pressure of 23000 Pa and the atmospheric temperature of −55 deg. C., such that the fan pressure ratio is no greater than 1.5.

8. The method of claim 1, wherein: the gas turbine engine further comprises an annular splitter at which divides flow between a core flow that flows through the engine core, and a bypass flow that flows along a bypass duct; a fan root pressure ratio is defined as a ratio of a mean total pressure of a flow at a fan exit that subsequently flows through the engine core to a mean total pressure of a flow at a fan inlet; and the method further comprises operating the gas turbine engine to propel the aircraft at the forward Mach number of 0.8, at the atmospheric pressure of 23000 Pa and the atmospheric temperature of −55 deg. C., such that the fan root pressure ratio is no greater than 1.25.

9. The method of claim 8, wherein: a fan tip pressure ratio is defined as a ratio of a mean total pressure of flow at the fan exit that subsequently flows through the bypass duct to the mean total pressure of the flow at the fan inlet; and the method further comprises operating the gas turbine engine to propel the aircraft at the forward Mach number of 0.8, at the atmospheric pressure of 23000 Pa and the atmospheric temperature of −55 deg. C., such that a ratio between the fan root pressure ratio to the fan tip pressure ratio is less than 0.95.

10. The method of claim 1, wherein a diameter of the fan is in a range of from 250 cm to 390 cm.

11. The method of claim 1, wherein the fan blades comprise a main body attached to a leading edge sheath, the main body and the leading edge sheath being formed of different materials.

12. The method of claim 11, wherein: the leading edge sheath material comprises titanium; and the main body material comprises carbon fibre or an aluminium alloy.

13. The method of claim 1, wherein: the gas turbine engine further comprises an intake that extends upstream of the fan blades; an intake length L is defined as an axial distance between a leading edge of the intake and a leading edge of a tip of the fan blades; a fan diameter D is a diameter of the fan at the leading edge of the tips of the fan blades; and a ratio L/D is in a range of from 0.2 to 0.45.

14. The method of claim 1, wherein the gearbox reduction ratio GR is in a range of from 3.1 to 3.8.

15. The method of claim 1, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

16. A method of operating a gas turbine engine for an aircraft, the gas turbine engine having (i) an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, (ii) a fan located upstream of the engine core, the fan comprising a plurality of fan blades, an annular fan face being defined at a leading edge of the fan, and (iii) 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, the gearbox having a reduction ratio GR, the method comprising: operating the gas turbine engine to propel the aircraft at a forward Mach number of 0.8, at an atmospheric pressure of 23000 Pa and an atmospheric temperature of −55 deg. C., such that: a cycle operability parameter β satisfies the following:
1K.sup.−1/2≤β≤2K.sup.−1/2 where: a quasi-non-dimensional mass flow rate Q is defined as: $Q=W\frac{\sqrt{T0}}{P0.Math.A_{\mathrm{fan}}}$ W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of 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 an area of the fan face in m.sup.2; a specific thrust ST is defined as net engine thrust (N) divided by mass flow rate (Kg/s) through the engine; the cycle operability parameter β is defined as: $\beta =\frac{GR}{Q.Math.\mathrm{ST}};$ and the specific thrust ST satisfies the following:
70Nkg.sup.−1s≤ST≤85Nkg.sup.−1s.

17. A method of operating a gas turbine engine for an aircraft, the gas turbine engine having (i) an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, (ii) a fan located upstream of the engine core, the fan comprising a plurality of fan blades, an annular fan face being defined at a leading edge of the fan, and (iii) 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, the gearbox having a reduction ratio GR, the method comprising: operating the gas turbine engine to propel the aircraft at a forward Mach number of 0.8, at an atmospheric pressure of 23000 Pa and an atmospheric temperature of −55 deg. C., such that: a cycle operability parameter β satisfies the following:
1K.sup.−1/2≤β≤2K.sup.−1/2 where: a quasi-non-dimensional mass flow rate Q is defined as: $Q=W\frac{\sqrt{T0}}{P0.Math.A_{\mathrm{fan}}}$ W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of 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 an area of the fan face in m.sup.2; a specific thrust ST is defined as net engine thrust (N) divided by mass flow rate (Kg/s) through the engine; the cycle operability parameter β is defined as: $\beta =\frac{GR}{Q.Math.\mathrm{ST}};$ and the quasi-non-dimensional mass flow rate Q satisfies the following:
0.032Kgs.sup.−1N.sup.−1K.sup.1/2≤Q≤0.036Kgs.sup.−1N.sup.−1K.sup.1/2.

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

(5) FIG. 4 is another close up sectional side view of an upstream portion of a gas turbine engine showing flow parameters.

(6) 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.

(7) 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 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, having a reduction ratio GR. In other words, the ratio between the rotational speed of the low pressure turbine 19 and the rotational speed of the fan 23 is GR.

(8) 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.

(9) 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.

(10) 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.

(11) 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.

(12) 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.

(13) 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.

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

(15) 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 20, 18 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.

(16) 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.

(17) Referring to FIG. 2 and FIG. 4, an example of a gas turbine engine 10 in accordance with the present disclosure is shown. Engine dimensions are shown in FIG. 2 and flow parameters at cruise conditions are shown in FIG. 4. The fan 23 comprises a plurality of fan blades 230. Each fan blade 230 has a tip 231, a leading edge 232, a trailing edge 234, and a hub (which may be referred to as a root) 235. The hub 235 may be said to define the radially inner boundary of the gas washed surfaces of the fan blade 230. The fan blade 230 (i.e. the gas washed surfaces of the fan blade 230, which include a pressure surface and a suction surface) extend in a generally radial direction from the root 235 to the tip 231.

(18) A fan face area A.sub.fan is shown in FIG. 2 and calculated as:

(19) $A_{\mathrm{fan}}=\frac{\pi D^2}{4}\left(1-{\left(\frac{h}{t}\right)}^2\right)$

(20) FIG. 2 also shows how the values D, h and t relate to the gas turbine engine 10. In particular:

(21) D is the diameter (in metres) of the fan at the leading edge 232 (i.e. at the tips 231 of the leading edge 232 of the fan blades 230);

(22) h is the distance (in metres) between the centreline 9 of the engine 10 and the radially inner point on the leading edge 232 of the fan blade 230 (i.e. the intersection of the leading edge 232 and the hub 235); and

(23) t is the distance (in metres) between the centreline 9 of the engine 10 and the radially outer point (i.e. at the tip 231) on the leading edge 232 of the fan blade (i.e. t=D/2).

(24) The value (h/t) may be referred to elsewhere herein—and in other literature in the field—as the hub-to-tip ratio.

(25) As noted elsewhere herein, a quasi non-dimensional mass flow rate Q is defined as:

(26) $Q=W\frac{\sqrt{T0}}{P0.Math.A_{\mathrm{fan}}}.$

(27) Where:

(28) W is mass flow rate through the fan in Kg/s;

(29) T0 is average stagnation temperature of the air at the fan face in Kelvin;

(30) P0 is average stagnation pressure of the air at the fan face in Pa; and

(31) A.sub.fan is the area of the fan face in m.sup.2, as defined above.

(32) The parameters W, T0, P0 and A.sub.fan are all shown schematically in FIGS. 2 and 4.

(33) At cruise conditions of the gas turbine engine 10 (which may be as defined elsewhere herein), the value of Q may be in the ranges described and/or claimed herein, for example in the range of from 0.029 to 0.034 Kgs.sup.−1N.sup.−1K.sup.1/2.

(34) Also at cruise conditions, the gas turbine engine 10 generates a thrust T (which may be referred to as a cruise thrust), shown schematically in FIG. 4. This thrust may be equal to the thrust required to maintain the cruise forward speed of an aircraft to which the gas turbine engine 10 is attached, divided by the number of engines 10 provided to the aircraft.

(35) At cruise conditions, the specific thrust ST, which is calculated as the thrust T divided by the mass flow rate W through the engine (which is equal to the mass flow rate W at the fan inlet) may be in the ranges described and/or claimed herein, for example in the range of from 70 Nkg.sup.−1s to 110 Nkg.sup.−1s.

(36) At cruise conditions, a cycle operability parameter β is in the ranges described and/or claimed herein, for example in the range of from 1 K.sup.−1/2 to 2 K.sup.−1/2, where the cycle operability parameter β is defined as:

(37) $\beta =\frac{\mathrm{GR}}{Q.Math.\mathrm{ST}}$

(38) As noted above, downstream of the fan 13 the air splits into two separate flows: a first air flow A into the engine core and a second air flow B which passes through a bypass duct 22 to provide propulsive thrust. The first and second airflows A, B split at a generally annular splitter 140, for example at the leading edge of the generally annular splitter 140 at a generally circular stagnation line.

(39) A stagnation streamline 110 stagnates on the leading edge of the splitter 140. The stagnation streamlines 110 around the circumference of the engine 10 form a streamsurface 110. All of the flow A radially inside this streamsurface 110 ultimately flows through the engine core. The streamsurface 110 forms a radially outer boundary of a streamtube that contains all of the flow that ultimately flows through the engine core, which may be referred to as the core flow A. All of the flow B radially outside the streamsurface 110 ultimately flows through the bypass duct 22. The streamsurface 110 forms a radially inner boundary of a streamtube that contains all of the flow B that ultimately flows through the bypass duct 22, which may be referred to as the bypass flow B.

(40) The ratio of the mass flow rate of the bypass flow B to the core flow A may be as described and/or claimed herein, for example at least 10, 11, 12 or 13.

(41) In use, the fan blades 230 of the fan 23 do work on the flow, thereby raising the total pressure of the flow. A fan root pressure ratio is defined as the mean total pressure of the flow at the fan exit that subsequently flows (as flow A) through the engine core to the mean total pressure at the inlet to the fan 23. With reference to FIG. 4, the mean total pressure of the flow at the fan exit that subsequently flows through the engine core is the mean total pressure P.sub.A of the flow that is just downstream of the fan 23 and radially inside the streamsurface 110. Also in FIG. 4, the mean total pressure P0 at the inlet to the fan 23 is the mean total pressure over the surface that extends across the engine (for example from the hub 235 to the tip 231 of the fan blade 230) and is immediately upstream of the fan 23.

(42) The value of the fan root pressure ratio (P.sub.A/P0) may be described and/or claimed herein, for example less than 1.25, for example less than 1.22.

(43) A fan tip pressure ratio is defined as the mean total pressure P.sub.B of the flow at the fan exit that subsequently flows (as flow B) through the bypass duct 22 to the mean total pressure at the inlet to the fan 23. With reference to FIG. 4, the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct 22 is the mean total pressure over the surface that is just downstream of the fan 23 and radially outside the streamsurface 110.

(44) The ratio between the fan root pressure ratio (P.sub.A/P0) and the fan tip pressure ratio (P.sub.B/P0) may be as described and/or claimed herein, for example less than 0.95, and/or less than 0.9 and/or less than 0.85. This ratio may alternatively be expressed simply as the ratio between the mean total pressure (P.sub.A) of the flow at the fan exit that subsequently flows (as flow A) through the engine core to the mean total pressure (P.sub.B) of the flow at the fan exit that subsequently flows (as flow B) through the bypass duct 22.

(45) The fan blades 230 may be manufactured using any suitable material or combination of materials, as described elsewhere herein. Purely by way of example, in the FIG. 4 example, the fan blade 330 has a main body 350 attached to a leading edge sheath 360. The main body 350 and the leading edge 360 in the FIG. 4 example are manufactured using different materials. Purely by way of example, the main body 350 may be manufactured using a carbon fibre composite material or an aluminium alloy material (such as an aluminium lithium alloy), and the leading edge sheath 360 may be manufactured from a material that is better able to withstand being struck by a foreign object (such as a bird). Again, purely by way of example, the leading edge sheath may be manufactured using a titanium alloy.

(46) As explained elsewhere herein, gas turbine engines having a cycle operability parameter β in the ranges outlined herein may provide various advantages, such as improving the bird strike capability whilst retaining the efficiency advantages associated with geared and/or low specific thrust gas turbine engines. This may allow greater design freedom in other aspects of the fan system (including fan blades), such as weight, aerodynamic design, complexity and/or cost.

(47) 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 232 of the tip 231 of the fan blades 230, 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 0.45. This may lead to further advantages, such as installation and/or aerodynamic benefits.

(48) The gas turbine engine 10 shown by way of example in FIG. 1 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.

(49) 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.