Compression in a gas turbine engine

Abstract

A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

Claims

1. A gas turbine engine for an aircraft comprising: an engine core comprising a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine being a higher pressure turbine than the first turbine and the second compressor being a higher pressure compressor than the first compressor, and wherein a second turbine entrance temperature (T40) is defined as an average temperature of airflow at the entrance to the second turbine at cruise conditions, a first turbine entrance temperature (T44) is defined as an average temperature of airflow at the entrance to the first turbine at cruise conditions, a second turbine exit temperature (T42) is defined as an average temperature of airflow at the exit from the second turbine at cruise conditions, and a first turbine exit temperature (T50) is defined as an average temperature of airflow at the exit from the first turbine at cruise conditions, and wherein a low pressure turbine temperature change is defined as: $\frac{\mathrm{the}\mathrm{first}\mathrm{turbine}\mathrm{entrance}\mathrm{temperature}\left(T44\right)}{\mathrm{the}\mathrm{first}\mathrm{turbine}\mathrm{exit}\mathrm{temperature}\left(T50\right)},$ and a high pressure turbine temperature change is defined as: $\frac{\mathrm{the}\mathrm{second}\mathrm{turbine}\mathrm{entrance}\mathrm{temperature}\left(T40\right)}{\mathrm{the}\mathrm{second}\mathrm{turbine}\mathrm{exit}\mathrm{temperature}\left(T42\right)}\text{;}$ and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, and wherein a low to high pressure turbine temperature change ratio of: $\frac{\mathrm{the}\mathrm{low}\mathrm{pressure}\mathrm{turbine}\mathrm{temperature}\mathrm{change}}{\mathrm{the}\mathrm{high}\mathrm{pressure}\mathrm{turbine}\mathrm{temperature}\mathrm{change}}$ is in the range from 1.09 to 1.25, wherein the second turbine entrance temperature (T40) is in the range from 1400K to 1700K, and wherein the cruise conditions means the conditions at mid-cruise of the aircraft to which the gas turbine engine is attached, and the cruise conditions at mid-cruise means the conditions experienced by the aircraft and the gas turbine engine at a point between top of climb and start of descent at which 50% of the total fuel that is burned between the top of climb and the start of descent has been burned.

2. The gas turbine engine according to claim 1, wherein the low to high pressure turbine temperature change ratio is in the range from 1.10 to 1.25.

3. The gas turbine engine according to claim 1, wherein (i) the low pressure turbine temperature change (T44/T50) is in the range from 1.6 to 1.85; and/or (ii) the high pressure turbine temperature change (T40/T42) is in the range from 1.40 to 1.55.

4. The gas turbine engine according to claim 1, wherein the first turbine is arranged to receive airflow from the exit of the second turbine, such that the first turbine entrance temperature (T44) is at least substantially equal to the second turbine exit temperature (T42).

5. The gas turbine engine according to claim 1, wherein the engine comprises more than two turbines, and wherein the highest pressure turbine of the engine is selected as the second turbine and the lowest pressure turbine of the engine is selected as the first turbine.

6. The gas turbine engine according to claim 1, wherein a specific thrust of the engine at cruise conditions, defined as net engine thrust divided by mass flow rate through the engine, is in the range from 50 to 100 Nkg.sup.−1 s.

7. The gas turbine engine according to claim 1, wherein a quasi-non-dimensional mass flow rate Q is defined as: $Q=W\frac{\sqrt{T_0}}{P_0.Math.A_{\mathrm{fan}}}.$ where: W is mass flow rate through the fan in Kg/s; T.sub.0 is average stagnation temperature of the air at the fan face in Kelvin; P.sub.0 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 has a value in the range from 0.025 to 0.038 Kgs.sup.−1 N.sup.−1 K.sup.1/2 at cruise conditions.

8. The gas turbine engine according to claim 1, wherein the first turbine entrance temperature (T44) is in the range from 960K to 1150K.

9. The gas turbine engine according to claim 1, wherein the first turbine exit temperature (T50) is in the range from 590K to 640K.

10. The gas turbine engine according to claim 1, wherein (i) the engine comprises a total of two turbines and the first turbine entrance temperature (T44) is at least substantially equal to the second turbine exit temperature (T42); or (ii) the engine comprises more than two turbines and the high pressure turbine temperature change provides a measure of the temperature change across all turbines except the lowest pressure turbine.

11. The gas turbine engine according to claim 1, wherein the fan has a fan tip radius and: (i) the fan tip radius is in the range from 110 cm to 150 cm; or (ii) the fan tip radius is in the range from 155 cm to 200 cm.

12. The gas turbine engine according to claim 1, further comprising 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.

13. The gas turbine engine according to claim 12, wherein the gearbox has a gear ratio in the range from 3.2 to 5.

14. The gas turbine engine according to claim 1, wherein the first turbine comprises at least four rotor stages.

15. A method of operating a gas turbine engine on an aircraft, wherein the method comprises: operating the gas turbine engine to provide propulsion under cruise conditions according to claim 1.

16. The method according to claim 15, further comprising operating the gas turbine engine to provide propulsion under cruise conditions such that the first turbine entrance temperature (T44) is in the range from 960K to 1150K.

17. The method according to claim 15, further comprising operating the gas turbine engine to provide propulsion under cruise conditions such that 5 the first turbine exit temperature (T50) is in the range from 590K to 640K.

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;

(5) FIG. 4A is a close up sectional side view of an upstream portion of the gas turbine engine shown in FIG. 2, with indications of where various temperatures are to be measured marked;

(6) FIG. 4B is the close up sectional side view of FIG. 4A with regions within which the various temperatures may be measured marked;

(7) FIG. 5A is a sectional side view of the gas turbine engine shown in FIG. 1, with indications of where various temperatures are to be measured marked;

(8) FIG. 5B is the sectional side view of FIG. 5A with regions within which the various temperatures may be measured marked;

(9) FIG. 6 is a schematic side view of a gas turbine engine:

(10) FIG. 7 is the schematic side view of a gas turbine engine as shown in FIG. 1, with turbine details highlighted;

(11) FIG. 8 illustrates a method;

(12) FIG. 9 is a close up sectional side view of an upstream portion of the gas turbine engine shown in FIG. 2, with indications of flows and areas marked; and

(13) FIG. 10 is a perspective view of an aircraft with two gas turbine engines mounted thereon.

(14) 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.

(15) In the embodiment being described, the nacelle inner radius at the axial position of the leading edge blade tips 68a is arranged to be slightly larger than the fan tip radius 102, such that the fan 23 can fit within the nacelle 21 without the blade tips 68 rubbing the nacelle 21. More particularly, in the embodiment being described the engine 10 comprises an engine fancasing 21a adjacent the blade tips 68a; the nacelle 21 is mounted on/around the engine fancasing 21a such that the engine fancasing 21a and the nacelle 21 form and surround an outer surface of the gas path though the engine 10. Fancasing inner radius at the axial position of the leading edge blade tips 68a is arranged to be slightly larger than the fan tip radius 102, such that the fan 23 can fit within the engine fancasing 21a without the blade tips 68 rubbing the fan casing 21a. In some alternative embodiments, the blade tips 68a may be arranged to rub the fancasing 21a.

(16) In the embodiments shown in the Figures, the engine fancasing 21a extends only in the region of the fan 23. In alternative embodiments, the fancasing 21a may extend rearwardly, for example to the axial location of a bypass duct outlet guide vane (OGV) 58.

(17) 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.

(18) 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 process 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.

(19) 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.

(20) Each of the compressors provided in the gas turbine engine 10 (e.g. the lower pressure compressor 14 and the high pressure compressor 15) comprises any number of compression stages, for example multiple compression stages. Each compression stage may comprise a row of rotor blades and a row of stator vanes that are axially offset from each other. The fan 23 also provides compression of airflow, and so provides an additional compression stage separate from those of the low and high pressure compressors 14, 15. A compression stage number is defined as the total number of compression stages provided by the fan 23 and the one or more compressors 14, 15 provided in the gas turbine engine 10. In the presently described embodiment, the compression stage number is therefore the sum of the compression stages provided in the low pressure compressor 14, the high pressure compressor 15 and the fan 23.

(21) In other embodiments, the compression stages provided in the compressors 14, 15 of the gas turbine engine may not be axial compression stages. In some embodiments, one or more radial compression stages may be provided in addition, or alternatively, to the axial compression stages provided in each compressor. For example, in one embodiment, the low pressure compressor 14 and/or the high pressure compressor 15 may comprise one or more axial compression stages (each formed by a row of rotor blades and stators) followed by a radial compression stage provided downstream of the axial compression stage or stages. In yet other embodiments, each of the compressors may comprise only radial compression stages. The compression stage number is defined as the total number of compression stages, including both radial and axial compression stages (including the fan 23). In all of the embodiments described above, each radial compression stage may comprise a centrifugal compressor.

(22) 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.

(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) As noted above, downstream of the fan 23 the air splits into two separate flows: a first air flow A into the engine core 11 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. Referring to FIGS. 4A and 4B, the first and second airflows A, B split at a generally annular splitter 70, for example at the leading edge of the generally annular splitter 70 at a generally circular stagnation line. The splitter 70 is provided by a forwardmost portion of the core casing 11a in the embodiments being described, and may alternatively be referred to as a forwardmost tip 70 of the engine core 11 in some embodiments.

(30) A stagnation streamline 110 stagnates on the leading edge of the splitter 70. 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 11. 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. The streamsurface 110, and correspondingly the streamtube, may be defined at cruise conditions.

(31) The flow at the fan exit that subsequently flows through the engine core 11 may therefore be defined by a streamtube that extends from the fan exit to the engine core 11. Such a streamtube may be bounded by a radially outer surface that terminates at the splitter 70, i.e. by a radially outer surface that is formed by streamlines 110 that terminate at a stagnation point on the splitter 70. Such a streamtube may be generally annular. For example a cross-section through such a streamtube may be substantially annular at any given cross-section perpendicular to the engine (rotational) axis 9.

(32) Definitions of various temperatures, radii, and other parameters are provided below for ease of reference.

(33) Fan Tip Radius:

(34) The radius 102 of the fan 23, also referred to as the fan tip radius 102, or R.sub.fan tip, may be measured between the engine centreline 9 and the tip 68a of a fan blade 64 at its leading edge 64a (in a radial direction). The fan diameter (D) may simply be defined as twice the radius 102 of the fan 23.

(35) In the embodiments being described, the fan tip radius 102 is in the range from 95 cm to 200 cm, or from 110 cm to 200 cm. In some embodiments, the fan tip radius is in the range from 95 cm to 150 cm or from 110 cm to 150 cm. In some alternative embodiments, the fan tip radius is in the range from 155 cm to 200 cm

(36) In some embodiments, the fan diameter is in the range from 190 cm to 300 cm, or 220 cm to 300 cm. In some alternative embodiments, the fan diameter is in the range from 310 cm to 400 cm.

(37) The skilled person would appreciate that fan blades 64 may expand in operation, and that the fan tip radius 102 under cruise conditions may be slightly greater than the fan tip radius 102 measured when the fan 23 is not in use. Fan tip radius 102 may be defined under cruise conditions. However, the skilled person would appreciate that the change in fan tip radius 102 is generally small compared to the fan tip radius and that the radius as measured when not in operation may be used.

(38) Hub Radius:

(39) The hub radius, R.sub.hub, is the (radial) distance 103 (in metres) between the centreline 9 of the engine 10 and the radially inner point on the leading edge 64a of the fan blade (i.e. of radially inner point of the gas-washed surface of the fan blade)—this is equivalent to the radius of the hub 66 of the fan 23 at the point at which the leading edge of each blade 64 extends therefrom.

(40) Fan Area:

(41) The fan face area, A.sub.fan, which may also be described as the flow area of the fan, is defined as the annular area between fan blade tips 68 and the hub 66 at the axial position of the fan blade leading edge tip 68a. The fan face area is measured in a radial plane (i.e. a plane perpendicular to the engine axis 9 and containing radii of the engine at the axial position of the plane). The skilled person will appreciate that A.sub.fan is at least substantially equivalent to the area of the annulus formed between the hub 66 of the fan 23 and the inner surface of the nacelle 21 immediately adjacent the leading edge blade tips (as the blade tip leading edges 64a are arranged to lie very close to the inner surface of the nacelle 21—noting the above comments about the fancasing 21a) for the fan engine 10 being described, and is therefore equivalent to the fan face area minus the area taken by the hub 66.

(42) As referred to herein, the flow area of the fan (A.sub.fan) is defined as:
A.sub.fan=π(R.sub.fan tip.sup.2−R.sub.hub.sup.2)
Where:
R.sub.fan tip is the radius 102 (in metres) of the fan 23 at the leading edge (i.e. at the tips 68a of the leading edge of the fan blades 64).
R.sub.hub is the distance 103 (in metres) between the centreline of the engine and the radially inner point on the leading edge of the fan blade (i.e. of radially inner point of the gas-washed surface of the fan blade)—this is equivalent to the radius of the hub 66 of the fan 23 at the point at which the leading edge of each blade 64 is connected thereto, and may be referred to as the hub radius 103.

(43) In various embodiments, the ratio of the radius of fan blade 64 at its hub 66 to the radius of the fan blade at its tip 68 may be less than 0.33.

(44) In the embodiment being described, the flow area is defined in a radial plane, and can therefore be calculated using the fan tip radius 102 and the hub radius 103.

(45) In the embodiments being described, the fan tip radius 102 is in the range from 95 cm to 200 cm, or from 110 cm to 200 cm. In some embodiments, the fan tip radius is in the range from 95 cm to 150 cm or from 110 cm to 150 cm. In some alternative embodiments, the fan tip radius is in the range from 155 cm to 200 cm

(46) In some embodiments, the fan diameter (twice the fan radius 102) is in the range from 190 cm to 300 cm, or 220 cm to 300 cm. In some alternative embodiments, the fan diameter is in the range from 310 cm to 400 cm.

(47) Q:

(48) A quasi-non-dimensional mass flow rate, Q, may be defined as:

(49) $Q=W\frac{\sqrt{T_0}}{P_0.Math.A_{\mathrm{fan}}},$
where:
W is mass flow rate through the fan in Kg/s;
T.sub.0 is average stagnation temperature of the air at the fan face in Kelvin;
P.sub.0 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.

(50) As referred to herein, the area of the fan face (A.sub.fan) is defined as:

(51) $A_{\mathrm{fan}}=\frac{\pi D^2}{4}\left(1-{\left(\frac{h}{t}\right)}^2\right)$
Where:
D is the diameter (in metres) of the fan at the leading edge (i.e. at the tips of the leading edge of the fan blades);
h is the distance (in metres) between the centreline of the engine and the radially inner point on the leading edge of the fan blade (i.e. of radially inner point of the gas-washed surface of the fan blade); and
t is the distance (in metres) between the centreline of the engine and the radially outer point on the leading edge of the fan blade (i.e. t=D/2)

(52) A.sub.fan may also be referred to as a fan flow area as it corresponds to the gas-washed area of the fan (the blade-swept area outside of the hub). This may be equivalently represented as:
A.sub.fan=π(R.sub.fan tip.sup.2−R.sub.hub.sup.2)
as described above.

(53) At cruise conditions, the value of Q may be in the range of from: 0.0295 to 0.0335; 0.03 to 0.033; 0.0305 to 0.0325; 0.031 to 0.032 or on the order of 0.031 or 0.032 Kgs.sup.−1 N.sup.−1 K.sup.1/2. Thus, it will be appreciated that the value of Q may be in a range having a lower bound of 0.029, 0.0295, 0.03, 0.0305, 0.031, 0.0315 or 0.032 and/or an upper bound of 0.031, 0.0315, 0.032, 0.0325, 0.033, 0.0335, 0.034, 0.0345 or 0.035. All values of Q referred to herein are provided in units of Kgs.sup.−1 N.sup.−1 K.sup.1/2.

(54) Temperatures

(55) All temperatures referred to herein are total temperatures; the sum of static temperature plus velocity/kinetic energy effects. Total temperatures may also be referred to as stagnation temperatures. All temperature values are listed in Kelvin, unless otherwise stated and all temperature ratios and rises are likewise calculated in Kelvin. All temperatures are defined at cruise conditions, as defined above. In particular, the ISA standards for cruise conditions may provide an indication of an expected ambient temperature. “Average” temperature is used to indicate a mean temperature.

(56) In the embodiments being described, temperatures may be defined or measured at the mid-cruise aerodynamic design point, which is defined as Mn 0.85, and an altitude of 10700 m (35,000 ft), and optionally more particularly of 10668 m, for engines 10 of the embodiments being described. The skilled person would appreciate that these cruise conditions are provided by way of example only and may vary for engines 10 of other embodiments. Under differing conditions, the absolute temperature values may vary whilst the ratios remain within the ranges described.

(57) The following temperatures are referred to herein, and a more detailed description of each is provided below in Table 1. The numbering used for the temperatures corresponds to that provided in SAE standard AS755F.

(58) TABLE-US-00001 TABLE 1 Temperatures Approximate temperature in various embodiments (Kelvin) T120 the fan (tip) rotor 23 entry In the range 235 to 265; optionally in the temperature range 242 to 252 (approximately equal across full May be, for example 244 or 250 blade length - therefore generally (altitude- and Mach number-dependent) equal to the fan root rotor entry temperature T20) T125 the fan tip rotor 23 exit temperature In the range 260 to 285; optionally in the range (may vary along the blade length - 270 to 280 defined as an average across a May be, for example 270 for a T120 of 244, or radially outer portion of the blade 278 for a T120 of 250 unless otherwise specified) T30 the compressor 15 exit temperature In the range 750 to 1050; optionally in the range (at the exit from the highest pressure of 780 or 815 to 1000 compressor in embodiments with May be, for example, 834, 835 or 1000 multiple compressors) T21 the core 11 entry temperature (may In the range 245 to 270; optionally in the range be equivalent to a fan root 69 exit 260 to 270 temperature) May be, for example, 260, 268 or 252 T42 a second (higher/highest pressure) In the range 960 to 1150; optionally in the range turbine 17 exit temperature T42 (also 1030 to 1100 or 1030 to 1090. at least substantially equal to the first May be, for example, 995, 1030 or 1100 (low pressure) turbine 19 entrance temperature T44 in the embodiments described) T50 a first (low pressure) turbine 19 exit In the range 590 to 640; optionally in the range temperature 600 to 630 or 605 to 615. May be, for example, 600, 612 or 630 T40 the second (high pressure) turbine 17 In the range 1400 to 1700; optionally in the entry temperature range 1450 to 1650, or 1520 to 1570 May be, for example, 1480, 1560 or 1650

(59) For example, in one embodiment having a fan diameter in the range from 330 to 380 cm, T120 may be 250 K and T125 may be 278 K, giving a temperature increase across the fan 23 of 28 Kelvin. The compressor exit temperature (T30) may be 834 K. The core entry temperature (T21) may be 268 K. The second turbine exit temperature (T42) may be 1030 K and the first turbine exit temperature (T50) may be 612 K. The second turbine entry temperature (T40) may be 1560 K, giving a temperature decrease across the second turbine 17 of 530 K.

(60) For example, in an alternative embodiment having a fan diameter in the range from 240 cm to 280 cm, T120 may be 245 K and T125 may be 270 K, giving a temperature increase across the fan 23 of 25 Kelvin. The compressor exit temperature (T30) may be 780 K. The core entry temperature (T21) may be 260 K. The second turbine exit temperature (T42) may be 1000 K and the first turbine exit temperature (T50) may be 630 K. The second turbine entry temperature (T40) may be 1480 K, giving a temperature decrease across the second turbine 17 of 480 K.

(61) The skilled person would appreciate that one or more of the temperatures listed in Table 1 may be measured or otherwise determined in various ways, for example by use of a temperature probe or rake, by modelling, or by indirect determination from a temperature measured (or otherwise determined) elsewhere in the engine 10. For example, T125 (fan tip exit temperature) may be measured by one or more probes mounted on a leading edge of an outlet guide vane 59 in the bypass duct 22 (for example the closest OGV 59 to the fan 23, if multiple bypass duct OGVs are present), or by a rake anywhere in the region labelled in FIG. 4B. Similarly, T21 (fan root exit temperature) may be measured by one or more probes mounted on a leading edge of an outlet guide vane 24 in the core duct, or by a rake anywhere in the region labelled in FIG. 4B.

(62) The skilled person would appreciate that one or more of the temperatures listed in Table 1 may be difficult to measure practically, for example the relatively high temperature T40. Various temperatures may therefore be inferred from temperature measurements elsewhere and a knowledge of engine properties and temperature relationships.

(63) The fan 23, which is located upstream of the engine core 11, comprises a hub 66 and a plurality of fan blades 64 extending from the hub 66. Each fan blade 64 has a leading edge 64a and a trailing edge 64b.

(64) The fan rotor entry temperature (T120) is defined as an average temperature of airflow across the leading edge 64a of each fan blade 64 at cruise conditions; in particular, the temperature may be defined across a radially outer portion of the leading edge 64a of each fan blade 64 at cruise conditions. The skilled person would appreciate that the temperature across the leading edge 64a of each fan blade 64 may be at least substantially equal across the entire leading edge 64a of the fan blade 64, and that an average across the outer portion (or across the whole blade length) may be taken. More specifically, T120 may be used to refer to the temperature across a radially outer (fan tip) portion of the leading edge 64a of each fan blade 64 at cruise conditions and T20 may be used to refer to the temperature across a radially inner (fan root) portion of the leading edge 64a of each fan blade 64 at cruise conditions. The term fan rotor entry temperature may therefore be used generally for T120 or T20.

(65) The fan rotor entry temperature T120 may be higher than the ambient temperature, for example by around 30 K in some embodiments, due to dynamic head/Mach number-related effects. The fan rotor entry temperature (T120) may therefore be measured or calculated anywhere within a relatively large region, as illustrated in FIG. 4B (anywhere within the nacelle 21 and up to the fan blades 64, as the increase of around 30 K from ambient may occur at or near the forward-most edge of the nacelle 21), but may more specifically be measured at or adjacent the fan's leading edge 64a. The Fan Rotor Entry Temperature, T120, is around 250K in the embodiment being described. The skilled person would appreciate that this may vary in other embodiments, based on factors such as altitude of cruise.

(66) A radially outer portion of each fan blade 64 is selected in line with standards for the definition of T120, and for convenience of comparison with other temperatures as described below. The skilled person would appreciate that, for the engine 10 shown at cruise, T10 (temperature at the nacelle's forward-most point) is at least substantially equal to T120 (temperature at the leading edge 64a of an outer region of the fan blade 64/near the fan tip), as the Mach number based temperature increase takes effect from that point, and that T10 and T120 may also be substantially equal to T20 (temperature at the leading edge 64a of an inner region of the fan blade 64/near the fan root).

(67) The radially outer portion of each fan blade 64 may be defined as the portion of each fan blade 64 washed by the bypass airflow B, which flows around the outside of the engine core 11 after passing the fan 23 (as opposed to the core airflow, A, which passes through the core 11). This bypass airflow B flows through the bypass duct 22 in the embodiments described herein.

(68) In the embodiment being described, the engine core 11 has a core radius 105 defined between the centreline 9 of the engine 10 and a forwardmost tip of the engine core 11; the forwardmost tip may be referred to as the splitter 70, as it divides the core airflow A from the bypass airflow B. The radially outer portion of each fan blade 64 is generally the portion of each fan blade 64 at a radial distance from the centreline 9 of the engine 10 greater than the core radius 105.

(69) The skilled person would appreciate that, in reality the streamsurface 110 may slope and/or curve relative to the engine axis 9, such that some of the gas stream passing the fan blade 64 at a radial distance from the centreline 9 of the engine 10 slightly less than or equal to the core radius 105 may still enter the bypass stream B in some embodiments. In the embodiments being described, the slope and/or curvature of the streamsurface 110 relative to the engine axis 9 is relatively small, such that using the radial position of the splitter 70 provides an at least substantially equivalent temperature to using the streamsurface 110, within measurement errors. The division at a set radial position may therefore provide an equivalent value which may be easier to determine than streamtube shape in some scenarios.

(70) The gas turbine engine 10 of the embodiment being described comprises a nacelle 21 surrounding the fan 23 and the engine core 11 and defining a bypass duct 22 outside of the engine core 11. The bypass airflow B flows through the bypass duct 22 after leaving the fan 23 in the embodiments being described. The radially outer portion of each fan blade 64 is therefore the portion of each fan blade 64 extending across the entrance to the bypass duct 22 in the embodiments being described.

(71) The fan tip rotor exit temperature (T125) is defined as an average temperature of airflow over the radially outer portion of each fan blade 64 at the trailing edge 64b of each fan blade 64 at cruise conditions. The radially outer portion is as defined for the leading edge 64a. T125 therefore corresponds to the temperature of the bypass stream B on leaving the fan 23.

(72) The fan tip rotor exit temperature (T125) and the fan rotor entry temperature (T120) both therefore refer to airflow temperatures across the fan blade portion which is located in a bypass stream of air B about to enter the bypass duct 22 (the radially outer fan blade portion).

(73) The Fan Rotor Entry Temperature, T20, may also refer to airflow temperatures across the leading edge of the fan blade portion which is located in a core stream of air A about to enter the engine core 11 (the radially inner fan blade portion/the portion of the gas stream radially inward of the streamsurface 110), as the temperature is equivalent across the leading edge 64a of the blade 64. This is generally not the same for the exit temperature T125, T21, as the temperature generally varies with radius across the trailing edge 64b of each blade.

(74) The skilled person would appreciate that airflow temperature generally increases across the fan 23 when at cruise conditions, as work is done on the air by the fan blades 64 and some of this work generally manifests as heat. The exit temperature T125, T21 is therefore generally higher than the entry temperature T120, T20.

(75) The ratio of the fan tip rotor exit temperature T125 to the fan tip rotor entry temperature T120, T125/T120 may therefore be referred to as the fan tip temperature rise. The fan tip temperature rise may be defined as the average temperature rise across the fan rotor portion in the bypass stream (B in FIG. 1). The fan tip temperature rise is greater than one, and more specifically is in the range from 1.11 to 1.05, in the embodiments being described. The temperature rise may be defined as the ratio of the mean total temperature of the flow at the fan exit that subsequently flows (as flow B) around the outside of the engine core 11 to the mean total temperature at the inlet to the fan 23.

(76) The fan tip temperature rise is relatively low in the embodiment being described. In various embodiments, the fan 23 may be rotated at a relatively low speed at cruise to facilitate the low temperature rise. For example, the fan 23 may rotate at less than 2000 rpm, and/or may have a tip speed below Mn 1.1. The fan 23 of such embodiments may have a fan diameter equal to or greater than 230 cm. The skilled person would appreciate that, in various embodiments, a gearbox 30 may be provided to facilitate slower rotation of the fan 23, and that the engine cycle may be designed around these parameters.

(77) The aerodynamic design of the fan 23 may be selected to facilitate obtaining relatively low temperature rises across the fan as described herein. For example, the fan 23 may be designed to have pressure ratios at cruise of: Fan Tip Pressure Ratio: in the range of 1.2-1.45; optionally in the range of 1.35-1.44; and further optionally equal to 1.41; Fan Root Pressure Ratio at cruise: in the range of 1.18-1.30, and optionally equal to 1.24; and/or Fan Pressure Ratio: in the range of 1.35-1.43, and optionally equal to 1.39.

(78) The fan tip pressure ratio is defined as the mean total pressure 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 FIGS. 4A and 4B, 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 immediately downstream of the fan 23 and radially outside the streamsurface 110.

(79) The 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 11 to the mean total pressure at the inlet to the fan 23. With reference to FIGS. 4A and 4B, the mean total pressure of the flow at the fan exit that subsequently flows through the engine core is the mean total pressure of the flow that is immediately downstream of the fan 23 and radially inside the streamsurface 110. The mean total pressure 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 66 to the tip 68 of the fan blade 66) and is immediately upstream of the fan 23.

(80) The fan pressure ratio is defined as the ratio of the mean total pressure of the air flow at the exit of the fan 23 to the mean total pressure of the air flow at the inlet of the fan 23.

(81) The hub to tip ratio of the fan, defined as R.sub.Hub/R.sub.fan tip, may also be selected as part of aerodynamic design considerations for the fan 23—for the engine 10 being described, the hub to tip ratio of the fan 23 is in the range from 0.285 to 0.2, and optionally in the range from 0.24 to 0.27.

(82) As mentioned above, the temperature across the trailing edge 64b of each blade 64 generally varies with radius—the average temperature of airflow B entering the bypass duct 22 is different from the average temperature of airflow A entering the engine core 11. A further temperature, T21, is therefore defined as the average temperature of airflow A entering the engine core 11, which corresponds to the average temperature across the radially inner portion of the trailing edge 64a of each fan blade 64 (the radially inner portion of each blade 64 being the remainder of the blade once the radially outer portion defined above is removed from consideration). T21 may be referred to as the inner fan rotor exit temperature or fan root exit temperature. T21 may be referred to as the core entry temperature, as it is the average temperature of airflow entering the engine core 11 at cruise conditions. As the temperature of the core airflow does not vary significantly between the trailing edge 64b of the fan blades 64 and the first stator/guide vane 24 within the core 11, T21 may be measured anywhere within that region as marked in FIG. 4B. For example, the core entry temperature (T21) may be measured at/defined as any or all of the following:

(83) (i) the temperature of the core airflow at the axial position of the forwardmost point 70 of the core casing 11a (the splitter 70);

(84) (ii) the temperature of the core airflow at the axial position of the leading edge of the forwardmost stator or rotor of the (forwardmost/lowest pressure) compressor 14; and/or

(85) (iii) the temperature of the airflow across the trailing edge 64b of a radially inner portion of each fan blade 64, the airflow across the radially inner portion of each fan blade 64 being arranged to provide the core airflow A.

(86) The compressor exit temperature (T30) is defined as an average temperature of airflow at the exit from the compressor 15. T30 is defined at the axial position of the trailing edge of the rearmost rotor of the compressor 15. In the embodiment being described, combustion equipment 16 located between the exit from the compressor 15 and the entrance to the turbine 17 located downstream of the compressor 15 provides heat to the gas flow leaving the compressor 15, so increasing the temperature of flow into the turbine 17 from T30 to T40—the difference between T30 and T40 may be around 800 K, or more in some embodiments.

(87) In the embodiment being described, the gas turbine engine comprises more than one compressor 14, 15, and more specifically comprises two compressors. In such embodiments, the compressor exit temperature T30 is defined at the exit from the highest pressure compressor 15.

(88) A core temperature rise may be defined as:

(89) $\frac{\mathrm{the}\mathrm{compressor}\mathrm{exit}\mathrm{temperature}\left(T30\right)}{\mathrm{the}\mathrm{fan}\mathrm{rotor}\mathrm{entry}\mathrm{temperature}\left(T120\mathrm{or}T20\right)}.$
The core temperature rise may therefore measure the change in core airflow A temperature caused by both the fan 23 and the compressor(s) 14, 15. The core temperature rise is defined as the temperature ratio across the core compression system; this may be thought of as a measure of core thermal efficiency. In the embodiments being described, the core temperature rise is in the range from 3.1 to 4.0, and optionally in the range from 3.3 to 3.5. The core temperature rise may be equal to 3.33.

(90) A core to fan tip temperature rise ratio may be defined as:

(91) $\frac{\mathrm{the}\mathrm{core}\mathrm{temperature}\mathrm{rise}}{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{temperature}\mathrm{rise}}=\frac{T30/T120}{T125/T120}=\frac{T30}{T125}$

(92) As the temperature of both the core flow A and the bypass flow B is the same at the leading edge 64a of the fan blades 64 (T20=T120), the same temperature value, T120, can be used for each flow A, B, so allowing T120 to cancel out as shown above.

(93) The core to fan tip temperature rise ratio may be in the range from 2.845 to 3.8, and optionally in the range from 2.9 to 3.2. The core to fan tip temperature rise ratio may be equal to 3 in some embodiments. The core to fan tip temperature rise ratio is therefore relatively high by virtue of the relatively high core temperature rise and relatively low fan tip temperature rise. The engine cycle may be devised, and/or engine parameters selected, based on these parameters.

(94) In the embodiment being described, a geared architecture and gearbox 30 are used to facilitate the lower fan tip temperature rise as described above. In addition, a compressor design is selected to provide a compressor 14, 15 with an aerodynamic design that is efficient at a high level of loading. In the embodiments being described, the compressor design comprises 13 or more stages of compression (including the fan 23 as the first stage) so as to provide the desired effects. The compressor design may comprise a maximum of 16 stages of compression (including the fan 23) in some such embodiments. Each stage may be defined as a rotor, or a rotor-stator pair. In the embodiment being described, with 13 stages of compression—the fan 23 provides the first stage, the low pressure compressor 14 provides the subsequent three stages, and the high pressure compressor 15 provides the final nine stages. In alternative embodiments, the total number of compression stages may vary, the number of compressors 14, 15 may vary, and/or the split between the one or more compressor(s) may vary.

(95) In the embodiment being described the compressors 14, 15 are axial compressors. In various alternative embodiments, one or more of the compressors 14, 15 may be a centrifugal compressor.

(96) A high pressure ratio across the compressor(s) and an efficient level of core compression may therefore be achieved.

(97) The core to fan tip temperature rise ratio may be thought of as a relationship between the temperature rise across the core compression system (including the fan) and that across the bypass compression system at the cruise operating conditions (i.e. the mid-cruise operating point, which is an altitude of 10700 m (35,000 ft), or optionally more particularly of 10668 m, and a speed of 0.85 Mn in the embodiment being described.

(98) A core compressor temperature rise may be defined as:

(99) $\frac{\mathrm{the}\mathrm{compressor}\mathrm{exit}\mathrm{temperature}\left(T30\right)}{\mathrm{the}\mathrm{core}\mathrm{entry}\mathrm{temperature}\left(T21\right)}$
The core compressor temperature rise may be in the range from 2.9 to 4.0, and optionally in the range from 3.1 to 3.3. The core compressor temperature rise may be, for example, equal to 3.12.

(100) A core compressor to fan ti temperature rise ratio may be defined as:

(101) $\begin{array}{c}\frac{\mathrm{the}\mathrm{core}\mathrm{compressor}\mathrm{temperature}\mathrm{rise}}{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{temperature}\mathrm{rise}}=\frac{T30/T21}{T125/T120}\\ =\frac{T120\times T30}{T125\times T21}\end{array}$
The core compressor to fan tip temperature rise ratio may be in the range from 2.67 to 3.8, and more specifically in the range from 2.67 to 3.7 or from 2.67 to 3.5. The core compressor to fan tip temperature rise ratio may be in the range from 2.80 to 2.95, and optionally may be equal to 2.81. The core compressor to fan temperature rise ratio may be relatively high, resulting from a relatively low fan tip temperature rise, and/or a relatively high core compressor temperature rise.

(102) As compared to the core temperature rise, the core compressor temperature rise excludes the temperature rise across the fan root 69, and therefore only measures the heat imparted to the airflow A by the compressor(s) 14, 15.

(103) In the embodiment being described, obtaining a relatively low fan tip temperature rise is facilitated by use of a geared architecture and a gearbox 30, allowing the fan 23 to rotate more slowly than other drivetrain components. The relatively high core compressor temperature rise may be provided by having a core compressor aerodynamic design that is efficient at a high level of loading, which typically can be achieved with 13 stages of compression or greater as discussed above.

(104) The core compressor to fan tip temperature rise ratio may be thought of as a relationship between the temperature rise across the core compression system (excluding the fan) and that across the bypass compression system at the cruise operating condition.

(105) The core compressor temperature rise is defined as the temperature ratio across the core compression system; this may be thought of as a measure of core compressor thermal efficiency, by which the core compressor pressure rise is achieved.

(106) The skilled person would appreciate that one or more of the following engine features may be adjusted to obtain an engine 10 with a core compressor to fan tip temperature rise ratio within the specified range: A large flow area fan design, with the fan 23 being arranged to rotate at a relatively slow speed (optionally facilitated by use of a gearbox 30) in order to achieve a low fan tip temperature rise; and/or A core compression system with high levels of efficiency and optimised loading that facilitate obtaining a high thermal efficiency.

(107) A low temperature rise across the fan root 69 may facilitate the fan achieving a high propulsive efficiency while being operable and mechanically feasible.

(108) A fan root temperature rise may be defined as:

(109) 0$\frac{\mathrm{the}\mathrm{core}\mathrm{entry}\mathrm{temperature}\left(T21\right)}{\mathrm{the}\mathrm{fan}\mathrm{rotor}\mathrm{entry}\mathrm{temperature}\left(T120\right)}$
The fan root temperature rise may be in the range from 1.03 to 1.09, and optionally in the range from 1.05 to 1.07. The fan root temperature rise may be, for example, equal to 1.07.

(110) A core compressor to fan root temperature rise ratio may be defined using the fan root temperature rise and the core compressor temperature rise:

(111) $\begin{array}{c}\frac{\mathrm{the}\mathrm{core}\mathrm{compressor}\mathrm{temperature}\mathrm{rise}}{\mathrm{the}\mathrm{fan}\mathrm{root}\mathrm{temperature}\mathrm{rise}}=\frac{T30/T21}{T21/T120}\\ =\frac{T30\times T120}{T{21}^2}\end{array}$
The core compressor to fan root temperature rise ratio may be in the range from 2.76 to 4.1, optionally in the range from 2.8 to 3.2, and may be, for example, equal to 2.9.

(112) The engine 10 may have a geared architecture comprising a gearbox 30. The engine 10 may have a high pressure ratio and an efficient level of core compression, for example achieved by having a core compressor aerodynamic design that is efficient at a high level of loading which typically can be achieved with 13 stages of compression or greater, as discussed above.

(113) The fan root 69 may be designed to have a low temperature rise and a low level of work to facilitate the operability of the fan and the obtaining of a high level of propulsive efficiency. The high level of propulsive efficiency may be provided by a relatively straight fan root 69, having a low level of curvature relative to the curvature of the fan tip. For example, the curvature of the fan root 69 may be less than 60% of the curvature of the fan tip. In the embodiment being described, the curvature of the root portion of the blade is between 40% and 60% less than the curvature across the tip portion of the blade, and optionally around 50% less. In alternative or additional embodiments, the curvature of the root portion may be less than that of the tip portion by an amount within a range having a lower bound of any of 5%, 10%, 20%, 30%, 40% and an upper bound of any of 40%, 50% or 60%. The listed percentages are percentages of the blade camber (i.e. difference between a line that is tangent to the camber line at the leading edge of the blade 68 and a line that is tangent to the camber line at the trailing edge of the blade 68). The skilled person would appreciate that the “root portion” of a fan blade is sometimes taken to mean the portion of a fan blade 64 within the hub 66 and used to connect the blade 64 to the hub 66; this is not the case as used herein—the root portion 69 refers to the radially inner portion of the blade as described elsewhere herein, extending from the hub 66 and across the entrance to the core 11. The radially inner portion of the blade as defined herein may also be known as the “hub section” of the blade by the skilled person.

(114) The Overall Pressure Ratio (OPR) of an engine 10 with a core compressor to fan root temperature rise ratio in the listed range may be greater than 40.

(115) In various embodiments, the engine core 11 comprises: a first, lower pressure, turbine 19 (sometimes referred to as the low pressure turbine, or LPT), a first compressor 14, and a first core shaft 26 connecting the first turbine 19 to the first compressor 14; and a second, higher pressure, turbine 17 (sometimes referred to as the high pressure turbine, or HPT), a second compressor 15, a second core shaft 27 connecting the second turbine 17 to the second compressor 15.

(116) In such embodiments, a second turbine exit temperature (T42) may be defined as an average temperature of airflow at the exit from the second turbine 17 at cruise conditions and a first turbine exit temperature (T50) may be defined as an average temperature of airflow at the exit from the first turbine 19 at cruise conditions. T42 may be referred to as the high pressure turbine exit temperature. T50 may be referred to as the low pressure turbine exit temperature.

(117) In embodiments wherein the engine 10 comprises more than two turbines 17, 19, and the highest pressure turbine 17 of the engine 10 may be selected as the second turbine 17 and the lowest pressure turbine 19 of the engine 10 may be selected as the first turbine 19.

(118) As shown in FIG. 5A, T42 may be measured at the position of the rearmost rotor of the second turbine 17 and T50 may be measured at the position of the rearmost rotor of the first turbine 19.

(119) In the embodiment being described, the first turbine 19 is located immediately downstream of the second turbine 17. The second turbine exit temperature T42 may therefore be measured anywhere between the second turbine 17 and the first turbine 19, and may also be similar or equivalent to an entrance temperature (T44) for the first, lower pressure, turbine 19. FIG. 5B illustrates the measurement region for T42. The skilled person would appreciate that there may be a change in temperature, for example of around 10-20 K, across this region, for example due to cooling air. However, this difference may be too small as to noticeably affect the claimed ratios.

(120) In embodiments with more than two turbines 17, 19, the highest pressure turbine exit temperature T42 may not be similar or equivalent to the lowest pressure turbine entry temperature T44.

(121) The low(est) pressure turbine 19 is located immediately upstream of a core exit nozzle in the embodiment being described. T50 may be measured at the position of the rearmost rotor of the first turbine 19, or anywhere within the core exit nozzle. The skilled person would appreciate that core gas stream temperature will gradually equilibrate with the surrounding atmosphere as it leaves/once it has left the engine 10. FIG. 5B illustrates the measurement region for T50.

(122) A low pressure turbine temperature change may be defined as:

(123) $\frac{\mathrm{the}\mathrm{second}\mathrm{turbine}\mathrm{exit}\mathrm{temperature}\left(T42\right)}{\mathrm{the}\mathrm{first}\mathrm{turbine}\mathrm{exit}\mathrm{temperature}\left(T50\right)}.$
The low pressure turbine 19 temperature change may alternatively be defined as follows, noting that the first turbine entrance temperature T44 is generally comparable to the second turbine exit temperature T42:

(124) $\frac{\mathrm{the}\mathrm{first}\mathrm{turbine}\mathrm{entrance}\mathrm{temperature}\left(T44\right)}{\mathrm{the}\mathrm{first}\mathrm{turbine}\mathrm{exit}\mathrm{temperature}\left(T50\right)}$

(125) The skilled person would appreciate that temperature falls across a turbine in use, such that the entrance temperature (T42) to the first (low pressure) turbine 19 is higher than the exit temperature (T50) from the first turbine 19. The temperature change may therefore be described as a temperature fall or temperature drop.

(126) In embodiments with more than two turbines 17, 19, the low pressure turbine (LPT) temperature change may instead be called the lowest pressure turbine temperature change—providing a measure of the temperature change over just the lowest pressure turbine.

(127) The low pressure turbine temperature change therefore provides a measure of the average temperature change between the leading edge of the LPT 19 nozzle guide vane 19a (the stator of the forward-most rotor-stator pair of the LPT 19) and the trailing edge of the final rotor stage 19c of the LPT 19, as indicated in FIG. 5A.

(128) The low pressure turbine temperature change (T42/T50) may be in the range from 1.6 to 1.85, and optionally in the range from 1.65 to 1.8. The low pressure turbine temperature change may be, for example, equal to 1.68.

(129) A turbine to fan tip temperature change ratio may be defined as shown below, using the fan tip temperature rise as defined above:

(130) $\frac{\mathrm{the}\mathrm{low}\mathrm{pressure}\mathrm{turbine}\mathrm{temperature}\mathrm{change}}{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{temperature}\mathrm{rise}}$
The turbine to fan tip temperature change ratio may be higher than in known engines, for example being in the range from 1.46 to 2.0 and optionally from 1.46 to 1.66. The turbine to fan tip temperature change ratio may be below 2.00, and optionally in the range from 1.5 to 1.8 in some embodiments.

(131) In engines 10 of various embodiments with this temperature relationship, one or more of the following features may be present: A gearbox that allows the LPT 19 to operate at a higher speed, for example having a maximum operating speed between 4500 and 8000 rpm, e.g. for a fan 23 with a fan diameter in the range from 330 cm to 380 cm or alternatively example having a maximum operating speed between 7000 and 12000 rpm, e.g. for a fan 23 with a fan diameter in the range from 240 cm to 280 cm; The skilled person would appreciate that, in embodiments with a gearbox 30, LPT speed is generally equal to the fan speed multiplied by the gear ratio of the gearbox 30. A LPT 19 with favourable stage loading, for example having three or more rotor stages; An efficient aerodynamic fan design, for example having relatively low fan root curvature as compared to fan tip curvature (as described in more detail elsewhere herein); A fan 23 arranged to rotate at a relatively slow speed at cruise, optionally enabled by a gearbox 30; The rotational speed of the fan at cruise conditions may be, for example, less than 2500 rpm, or less than 2300 rpm. For an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270 cm), the fan speed may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. For an engine having a fan diameter in the range of from 330 cm to 380 cm, the fan speed may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.
and/or An engine designed to be low specific thrust, high bypass ratio and to have a low fan pressure ratio (the ratio of the mean total pressure of the air flow at the exit of the fan 23 to the mean total pressure of the air flow at the inlet of the fan 23, at cruise conditions). “Low specific thrust” may mean, in various embodiments, a thrust at cruise conditions in the range of 60 to 100 NKg.sup.−1 s, and optionally in the range of 70-90 NKg.sup.−1 s. For example, the thrust at cruise may be equal to or below 90 NKg.sup.−1 s, and optionally equal to or below 88 or 85 NKg.sup.−1 s. “High bypass ratio” may mean, in various embodiments, a bypass ratio at cruise conditions in the range of, for example, 12.5 to 30; “Low fan pressure ratio” may mean, in various embodiments, a fan pressure ratio at cruise conditions in the range of 1.2 to 1.45; and optionally in the range 1.35 to 1.43. Further optionally, the fan pressure ratio at cruise may be in the range 1.37 to 1.41 or 1.38 to 1.40. For example, the fan pressure ratio at cruise may be equal to 1.39. In alternative or additional embodiments, the fan pressure ratio at cruise may be equal to or around 1.35, 1.36, 1.37, 1.38, 1.39, 1.40, 1.41, 1.42 or 1.43, and may for example be in the range from 1.39 to 1.43, from 1.35 to 1.40, and/or from 1.37 to 1.40.

(132) To facilitate the operation of the lower pressure turbine (LPT) 19 so as to provide a larger temperature change (a higher magnitude drop in temperature across the lower pressure turbine 19), a gearbox 30 may be provided to allow the LPT 19 to operate at a favourable, higher, speed.

(133) The LPT 19 may be designed to provide a favourable stage loading for the larger temperature change. Although dependent on engine thrust, such a design may typically be achieved with a LPT with three or more rotor stages, and optionally with four or more rotor stages.

(134) In an engine core 11 comprising a second, higher pressure, turbine 17, a second compressor 15, a second, core shaft 27 connecting the second turbine 17 to the second compressor 15, and a first, lower pressure, turbine 19, a first compressor 14, and a first core shaft 26 connecting the first turbine 19 to the first compressor 14, a second turbine entrance temperature (T40) may be defined as an average temperature of airflow at the entrance to the second turbine 17 at cruise conditions. T40 may be measured/defined at the upstream (leading) edge of an inlet nozzle guide vane 17a of the high pressure turbine 17, as shown in FIG. 5A. An inlet nozzle guide vane 17a may be thought of as the forward-most stator of the high pressure turbine 17. In the embodiment being described, the gas stream temperature increases between the outlet of the compressor 15 and the inlet of the turbine 17 due to the combustor equipment 16; T40 may therefore be measured, or determined for a point, anywhere between the exit from the combustor equipment 16 and the entrance to the second turbine 17, as marked in FIG. 5B, and may more specifically be measured or otherwise determined at a leading edge of the most upstream stator 17a of the second turbine 17.

(135) A high(est) pressure turbine temperature change may be defined as:

(136) $\frac{\mathrm{the}\mathrm{second}\mathrm{turbine}\mathrm{entrance}\mathrm{temperature}\left(T40\right)}{\mathrm{the}\mathrm{second}\mathrm{turbine}\mathrm{exit}\mathrm{temperature}\left(T42\right)}.$
The high pressure turbine (HPT) temperature change may therefore be defined as the average temperature change between upstream of the HPT 17 inlet nozzle guide vane 17a (the forward most stator of the HPT 17) and after the final rotor stage of the HPT 17 as indicated in FIG. 5A.

(137) In embodiments with more than two turbines 17, 19, the high pressure turbine temperature change may instead be referred to as the high pressure turbines temperature change and may be a measure of the temperature change across all turbines excluding the lowest pressure turbine 19. If the second turbine 17 is the highest pressure turbine, the second turbine exit temperature T42 may therefore be replaced with the lowest-but-one pressure turbine exit temperature in the calculation of the ratio, which may be at least substantially equal to the lowest pressure turbine entry temperature T44. T44 may not be similar to or equal to T42 in such embodiments. The high pressure turbine temperature change may be in the range from 1.40 to 1.55, and optionally in the range from 1.44 to 1.52. The high pressure turbine temperature change may be equal to 1.5, for example being 1.50 or 1.51.

(138) A low to high pressure turbine temperature change ratio may then be defined as below, noting that T42 is generally at least similar to T44 (e.g. within 10-20 K):

(139) $\frac{\mathrm{the}\mathrm{low}\mathrm{pressure}\mathrm{turbine}\mathrm{temperature}\mathrm{change}}{\mathrm{the}\mathrm{high}\mathrm{pressure}\mathrm{turbine}\mathrm{temperature}\mathrm{change}}=\frac{T44\text{/}T50}{T40\text{/}T42}=\frac{T42\times T44}{T50\times T40}\approx \frac{T{42}^2}{T50\times T40}$

(140) The low to high pressure turbine temperature change ratio, which may also be referred to as a temperature fall ratio, may be in the range from 1.09 to 1.30, and optionally in the range from 1.10 to 1.25.

(141) The low to high pressure turbine temperature change ratio provides a relationship between the temperature change across the low pressure turbine 19 and the temperature change across the high pressure turbine 17 at the cruise operating conditions.

(142) To reduce fuel burn, and optionally reduce or minimise core size, and/or maximise thermal efficiency across the high pressure turbine, the inventors appreciated that a relatively low temperature change across the higher pressure turbine (HPT) 17 (as compared to the temperature change across the lower pressure turbine 19) may be beneficial.

(143) In various embodiments, this relatively low HPT temperature change may be obtained by using an HPT 17 with an efficient design, for example having two rotor stages, or only a single rotor stage.

(144) In the embodiments being described, the first turbine 19 is arranged to receive airflow from the exit of the second turbine 17, such that the first turbine entrance temperature (T42) is generally similar to the second turbine exit temperature (T42), sometimes with a 10 to 20 K difference due to the introduction of cooling air.

(145) In the embodiments being described, the second turbine 17 is arranged to receive airflow from the exit of the (high pressure) compressor 15; this airflow passes via the combustion equipment 16 between the compressor and the turbine, such that T30 (the second turbine entrance temperature) is higher than the compressor exit temperature T40.

(146) In engines 10 of various embodiments with this temperature relationship, one or more of the following features may be present: A gearbox arranged to allow the LPT 19 to operate at a favourable, higher, speed; A LPT 19 with optimal stage loading, for example having three or more rotor stages 19a, 19b, 19c; A high pressure compressor 15 with an aerodynamic design and a low level of loading, for example having nine or more rotor stages; An efficient HPT 17, for example having two rotor stages or fewer.

(147) The present disclosure also relates to methods 1000 of operating a gas turbine engine 10 on an aircraft 50. The methods 1000 are illustrated in FIG. 8. The method 1000 comprises starting up 1002 the engine 10 (e.g. prior to taxiing on a runway), and operating 1004 the engine during taxiing, take-off, and climb of the aircraft 50, as suitable, so as to reach cruise conditions. Once cruise conditions have been reached, the method 1000 then comprises operating 1006 the gas turbine engine 10 described in embodiments elsewhere herein to provide propulsion under cruise conditions.

(148) The gas turbine engine 10 is operated such that any one or more of the parameters or ratios defined herein are within the specified ranges. For example, the method comprises operating 1006 the gas turbine engine 10 such that any one or more of:

(149) a) the fan hub to tip ratio of:

(150) $\frac{\mathrm{the}\mathrm{fan}\mathrm{hub}\mathrm{radius}\left(103\right)}{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{radius}\left(102\right)}$

(151) is in the range from 0.2 to 0.285; and

(152) the fan tip temperature rise of:

(153) $\frac{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{rotor}\mathrm{exit}\mathrm{temperature}\left(T125\right)\mathrm{in}\mathrm{Kelvin}}{\mathrm{the}\mathrm{fan}\mathrm{rotor}\mathrm{entry}\mathrm{temperature}\left(T120\right)\mathrm{in}\mathrm{Kelvin}}$

(154) is in the range from 1.11 to 1.05;

(155) b) the core to fan temperature rise ratio of:

(156) $\frac{\mathrm{the}\mathrm{core}\mathrm{temperature}\mathrm{rise}}{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{temperature}\mathrm{rise}}$

(157) is in the range from 2.845 to 3.8;

(158) c) the core compressor to fan tip temperature rise ratio of:

(159) 0$\frac{\mathrm{the}\mathrm{core}\mathrm{compressor}\mathrm{temperature}\mathrm{rise}}{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{temperature}\mathrm{rise}}$

(160) is in the range from 2.67 to 3.8, and optionally 2.67 to 3.7;

(161) d) the core compressor to fan root temperature rise ratio of:

(162) $\frac{\mathrm{the}\mathrm{core}\mathrm{compressor}\mathrm{temperature}\mathrm{rise}}{\mathrm{the}\mathrm{fan}\mathrm{root}\mathrm{temperature}\mathrm{rise}}$

(163) is in the range from 2.76 to 4.1;

(164) e) the turbine to fan tip temperature change ratio of:

(165) $\frac{\mathrm{the}\mathrm{low}\mathrm{pressure}\mathrm{turbine}\mathrm{temperature}\mathrm{change}}{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{temperature}\mathrm{rise}}$

(166) is in the range from 1.46 to 2.0; and/or

(167) f) a low to high pressure turbine temperature change ratio of:

(168) $\frac{\mathrm{the}\mathrm{low}\mathrm{pressure}\mathrm{turbine}\mathrm{temperature}\mathrm{change}}{\mathrm{the}\mathrm{high}\mathrm{pressure}\mathrm{turbine}\mathrm{temperature}\mathrm{change}}$

(169) is in the range from 1.09 to 1.30, and optionally from 1.10 to 1.25.

(170) FIG. 10 illustrates an example aircraft 50 having a gas turbine engine 10 attached to each wing 52a, 52b thereof. Each gas turbine engine 10 is attached via a respective pylon 54a, 54b. When the aircraft 50 is flying under cruise conditions, as defined herein, each gas turbine engine 10 operates according to the parameters defined herein. For example, the gas turbine engines 10 operate such that any one or more of the conditions (a) to (f) defined for the method 1000 above are obtained.

(171) 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.