GEARED GAS TURBINE ENGINE

Abstract

A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

Claims

1. A method of operating a gas turbine engine on an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein the method comprises operating the gas turbine engine to provide propulsion such that a jet velocity ratio, R.sub.J, of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core is defined as: $R_J=\frac{V_B{C}_B}{V_C{C}_C{\eta }_{\mathrm{LPT}}{\eta }_F}$ where V.sub.B is a fully expanded first jet velocity, C.sub.B is a thrust coefficient of the bypass exhaust nozzle, V.sub.C is a fully expanded second jet velocity, C.sub.C is a thrust coefficient of the engine core exhaust nozzle, η.sub.LPT is an isentropic efficiency of a lowest pressure turbine of the engine core and η.sub.F is an isentropic efficiency of a fan tip; and the jet velocity ratio, R.sub.J, is between around 0.75 and 1.3 at maximum take-off conditions.

2. The method of claim 1, wherein the jet velocity ratio R.sub.J, is between around 0.8 and 1.0 at the maximum take-off conditions.

3. The method of claim 1, wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1800K at the maximum take-off conditions.

4. The method of claim 1, wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1850K at the maximum take-off conditions.

5. The method of claim 1, wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1900K at the maximum take-off conditions.

6. The method of claim 1, wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1950K at the maximum take-off conditions.

7. The method of claim 4, wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is less than 2000K at the maximum take-off conditions.

8. The method of claim 1, wherein the jet velocity ratio, R.sub.J, is between around 2 and 3 at idle conditions.

9. The method of claim 1, wherein the jet velocity ratio, R.sub.J, is between around 0.75 and 1.3 at cruise conditions.

10. The method of claim 1, wherein a gear ratio of the gearbox is between 3.1 and 3.3; a fan tip loading defined as dH/U.sub.tip.sup.2 is between 0.28 and 0.31 at cruise conditions, where dH is the enthalpy rise across the fan and U.sub.tip is the translational velocity of the leading edge of the fan tip; and a bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at the cruise conditions, is in a range of from 11.5 to 13.5.

11. The method of claim 10, wherein an overall pressure ratio defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at an exit of a highest pressure compressor is between 40 and 55 at the cruise conditions.

12. The method of claim 1, wherein the fan has an outer diameter of around 220 cm.

13. A method of operating a gas turbine engine on an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein the method comprises operating the gas turbine engine to provide propulsion such that a jet velocity ratio, R.sub.J, of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core is defined as: $R_J=\frac{V_B{C}_B}{V_C{C}_C{\eta }_{\mathrm{LPT}}{\eta }_F}$ where V.sub.B is a fully expanded first jet velocity, C.sub.B is a thrust coefficient of the bypass exhaust nozzle, V.sub.C is a fully expanded second jet velocity, C.sub.C is a thrust coefficient of the engine core exhaust nozzle, η.sub.LPT is an isentropic efficiency of a lowest pressure turbine of the engine core and η.sub.F is an isentropic efficiency of a fan tip; and the jet velocity ratio, R.sub.J, is between around 2 and 3 at idle conditions.

14. The method of claim 13, wherein the jet velocity ratio, R.sub.J, is between around 0.75 and 1.3 at cruise conditions.

15. The method of claim 13, wherein a fan tip loading defined as dH/U.sub.tip.sup.2 is between 0.28 and 0.35 at cruise conditions, where dH is the enthalpy rise across the fan and U.sub.tip is the translational velocity of the leading edge of the fan tip.

16. The method of claim 13, wherein a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is between 80 Nkg.sup.−1s and 95 Nkg.sup.−1s at cruise conditions.

17. The method of claim 13, wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1600K at cruise conditions.

18. The method of claim 13, wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is between 1600K and 1650K at cruise conditions.

19. The method of claim 13, wherein the fan comprises 22 or 24 fan blades.

20. The method of claim 13, wherein a gear ratio of the gearbox is between 3.1 and 3.4; a fan tip loading defined as dH/U.sub.tip.sup.2 is between 0.28 and 0.32 at cruise conditions, where dH is the enthalpy rise across the fan and U.sub.tip is the translational velocity of the leading edge of the fan tip; a bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core the at cruise conditions, is in a range of from 11.5 to 14; an overall pressure ratio defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at an exit of a highest pressure compressor is between 40 and 55 at the cruise conditions; the ratio of the radius of a fan blade of the fan at the hub to the radius of the fan blade at the tip is in a range of 0.27 to 0.31; and a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is between 80 Nkg.sup.−1s and 95 Nkg.sup.−1s at cruise conditions.

Description

[0061] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0062] FIG. 1 is a sectional side view of a gas turbine engine;

[0063] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0064] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0065] FIG. 4 is an example plot of change in fuel burn as a function of jet velocity ratio;

[0066] FIG. 5 is a schematic drawing of an aircraft having a gas turbine engine mounted thereon; and

[0067] FIG. 6 is a schematic drawing illustrating the concept of a fully expanded jet velocity.

[0068] 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.

[0069] 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.

[0070] 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.

[0071] 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.

[0072] 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.

[0073] 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.

[0074] 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.

[0075] 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.

[0076] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0077] 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.

[0078] 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.

[0079] FIG. 4 illustrates an example plot of change in fuel burn, ΔFB, as a function of jet velocity ratio, R.sub.J., other factors being constant. The change in fuel burn contribution from propulsive efficiency 401 is determined relative to an optimum value for the jet velocity ratio of around 1.0, with an increase in fuel burn both above and below this value. Factors that may affect the jet velocity ratio include the relative rotational speeds of the fan and turbine and the areas of the exhaust nozzles for the bypass and core exhausts. A lower gear ratio of the gearbox, i.e. a gear ratio of around 3.4 or lower, will tend to result in values for the jet velocity ratio of 1.0 or greater. To keep the fuel burn loss to within around 0.5% or less, it can be seen from FIG. 4 that the jet velocity ratio should be between around 1.0 and around 1.3. As the jet velocity ratio increases, the increase in fuel burn becomes greater. A further preferred upper limit for the jet velocity ratio is around 1.2, which keeps the increase in fuel burn to around 0.25-0.3%.

[0080] For a higher gearing ratio, i.e. around 3.3 to 3.4 and above, for example up to around 3.8 or in some cases even higher, the jet velocity ratio tends to be around 1.0 or below. As the jet velocity ratio decreases, the fuel burn contribution from propulsive efficiency 401 increases, and at a higher rate than for the portion above 1.0. To maintain this loss to within around 0.5%, it can be seen from FIG. 4 that the jet velocity ratio should be kept within around 0.8 to around 1.0, and for a ratio of around 0.75 and below, the fuel burn contribution from propulsive efficiency becomes dominant, rising to around 0.7% and above. A further preferred lower limit for the jet velocity ratio of around 0.85 or 0.90 may be used to keep the fuel burn contribution from propulsive efficiency to around 0.25% or below. However, further decreasing the jet velocity ratio enables a higher gear ratio to be used and/or decreases the pressure ratio across the IP turbine, thereby allowing for a smaller, faster and/or lighter IP turbine, reflected in a lower contribution to fuel burn loss 402 by the IP turbine. A range of around 0.75 to around 0.82 for the jet velocity ratio is thereby advantageous.

[0081] For a given set of gears making up an epicyclic gearbox, a planetary driving arrangement will produce a higher gearing ratio than a star driving arrangement. A star arrangement is therefore generally preferred in combination with a jet velocity ratio of around 1.0 and above, and a planetary arrangement for a jet velocity ratio of around 1.0 and below.

[0082] In a general aspect therefore, the gas turbine engine may be configured such that the jet velocity ratio is within a range from around 0.75 to around 1.3 at cruise conditions.

[0083] FIG. 5 illustrates an example aircraft 50 having a gas turbine engine 10 attached to each wing 51a, 51b thereof. When the aircraft is flying under cruise conditions, as defined herein, each gas turbine engine 10 operates such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine 10 and a second jet velocity exiting from an exhaust nozzle 20 of the engine core is within a range from around 0.75 to around 1.3.

[0084] FIG. 6 illustrates an example exhaust nozzle 60 of a gas turbine engine. The pressure Pj at the exit or throat 61 of the exhaust nozzle 60 is greater than the ambient pressure Pa around the engine. At some distance away from the nozzle exit 61 the jet pressure will equalise with the ambient pressure, i.e. Pj=Pa. The fully expanded jet velocity is defined as the jet velocity 62 at this point, i.e. the jet velocity along the axis of the engine at a minimum distance from the exhaust nozzle where the pressure is equal to ambient pressure.

[0085] Parameters that may be adjusted to achieve a jet velocity ratio within the desired range may include the LPT blade exit angle, LPT exit area, and the LPT rotation speed.

[0086] The following table illustrates example parameters for two engine examples, example 1 being for a relatively small, or lower power, engine and example 2 for a relatively large, or higher power, engine. A small engine may for example have a fan diameter of between around 200 and 280 cm and/or a maximum net thrust of between around 160 and 250 kN or as defined elsewhere herein. A large engine may for example have a fan diameter of between around 310 and 380 cm and/or a maximum net thrust of between around 310 and 450 kN or as defined elsewhere herein.

TABLE-US-00001 Example 1 Example 2 Parameter (small engine) (large engine) Fan diameter (cm) 215 320 LPT Exit Total Pressure at 130 130 maximum flow (kPa) Maximum LPT Exit 50 100 Mass Flow (kg/s) LPT Final Rotor Area (m.sup.2) 0.38 or less, for 0.75 or less, for example example 0.25 to 0.38 0.5 to 0.75 ESS Inlet Total Pressure at 140 140 maximum flow (kPa) ESS Inlet Mass Flow (kg/s) 50 100 ESS Inlet Rotor Area (m.sup.2) 0.275 or greater, 0.55 or greater, for example for example 0.27-0.3 0.55-0.6

[0087] The above parameters relating to LPT exit total pressure at maximum flow, maximum LPT exit mass flow and LPT final rotor area together determine the exit flow velocity of the LPT, i.e. the flow velocity at an exit of the engine core. The ESS inlet total pressure at maximum flow, maximum ESS inlet mass flow and ESS inlet rotor area together determine the velocity at the inlet of the engine core. The axial exhaust flow velocity from the bypass exhaust nozzle may be determined, at least in part, by the area of the bypass exhaust nozzle outlet.

[0088] In order to achieve a jet velocity ratio within the desired range, the fan may be provided with features such as a straighter fan root. The compressors, in particular the high pressure compressor, may be provided with features to manage their operability to allow the compressors to operate at a low power required to meet the defined ratios, which may for example include devices such as variable guide vanes. This changes the flow incidence onto the blades and helps to maintain an operability margin preventing the compressor from surging or stalling when operating at lower speeds.

[0089] 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.