COLD NOZZLE OPTIMISATION

Abstract

The present disclosure relates to optimisation of a cold nozzle, or bypass exit area, for a gas turbine engine, in particular for a geared turbofan gas turbine engine. Example embodiments include a method of optimising a geared turbofan gas turbine engine for an aircraft, the method comprising: determining expected service parameters for the aircraft, the expected service parameters including an expected range of travel for the aircraft; selecting components for the geared turbofan gas turbine engine to define a first smaller cold nozzle area if the range of travel is within a first smaller range and to define a second larger cold nozzle area if the range of travel is within a second larger range.

Claims

1. A method of optimising a geared turbofan gas turbine engine for an aircraft, the method comprising: determining expected service parameters for the aircraft, the expected service parameters including an expected range of travel for the aircraft; selecting components for the geared turbofan gas turbine engine to define a first smaller cold nozzle area if the range of travel is within a first smaller range and to define a second larger cold nozzle area if the range of travel is within a second larger range.

2. The method of claim 1, wherein the geared turbofan gas turbine engine comprises an engine core and a nacelle defining a bypass duct and the cold nozzle area.

3. The method of claim 2, wherein the components of the geared turbofan gas turbine engine include: the nacelle, or a portion thereof; and/or an outer casing of the engine core, or a portion thereof.

4. The method of claim 3, wherein the step of selecting components for the geared turbofan gas turbine engine from two alternative selections defining respective first and second cold nozzle areas.

5. The method of claim 1, wherein a specific thrust of the geared turbofan gas turbine engine with the first smaller cold nozzle area is at least 1% greater than a specific thrust of the geared turbofan gas turbine with the second larger cold nozzle area.

6. The method of claim 5, wherein the specific thrust of the geared turbofan gas turbine engine with the first smaller cold nozzle is between 1% and 3% greater than the specific thrust of the geared turbofan gas turbine with the second larger cold nozzle.

7. The method of claim 5, wherein the specific thrust of the geared turbofan gas turbine engine is within the range of 80 to 110 Nkg.sup.1s.

8. The method of claim 1, wherein a fan inlet area of the geared turbofan gas turbine engine is between 4.1 and 11.2 m.sup.2.

9. The method of claim 1, wherein the geared turbofan gas turbine engine is configured to have an fan tip air angle of between 57 and 62 degrees at cruise, the fan tip air angle being defined as: $={\tan }^{-1}\left(\frac{V_{\mathrm{ThetaBladeTip}}}{{\mathrm{Vx}}_{\mathrm{air}}}\right)$ $\mathrm{where}.Math..Math.V_{\mathrm{ThtaBladeTip}}=.Math..Math.\frac{D}{2}.Math.;$ is the fan rotational speed in radians/second, D is the outer tip diameter of the fan in metres at its leading edge and Vx.sub.air is the mean axial velocity of the flow into the fan over the leading edge of the fan blades.

10. The method of claim 1, further comprising assembling the geared turbofan gas turbine engine.

11. The method of claim 1, wherein the first smaller range of travel is between 556 and 1296 km (300 and 700 nautical miles) and the second larger range of travel is greater than 1296 km (700 nautical miles).

12. A plurality of gas turbine engines for a plurality of aircraft, each gas turbine engine comprising: an engine core comprising a turbine, 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 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; and an outer housing defining a bypass duct between the outer housing and the engine core and further defining a cold nozzle, wherein each of a first subset of the plurality of gas turbine engines comprises components defining a first smaller cold nozzle area and each of a second subset of the plurality of gas turbine engines comprises components defining a second larger cold nozzle.

13. The plurality of gas turbine engines of claim 12, wherein the components defining the first and second cold nozzle areas are the outer housing, or portion thereof, and/or an outer casing, or portion thereof, of the engine core.

14. The plurality of gas turbine engines of claim 12, wherein the first plurality of gas turbine engines are for aircraft having a first smaller range of travel and the second plurality of gas turbine engines are for aircraft having a second larger range of travel.

15. The plurality of gas turbine engines of claim 14, wherein the first smaller range of travel is between 556 and 1296 km (300 and 700 nautical miles) and the second larger range of travel is greater than 1296 km (700 nautical miles).

16. The plurality of gas turbine engines of claim 12, wherein each gas turbine engine (10) has a nominally identical engine core, fan and gearbox.

17. The plurality of gas turbine engines according to claim 12, wherein for each gas turbine engine: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

Description

DESCRIPTION OF THE DRAWINGS

[0052] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0053] FIG. 1 is a sectional side view of a gas turbine engine;

[0054] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0055] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0056] FIG. 4 is a schematic plot of fan bypass efficiency as a function of bypass flow rate for two types of gas turbine engine;

[0057] FIG. 5 is a sectional side view of a gas turbine engine, in which alternative bypass exits, or cold nozzles, are illustrated;

[0058] FIG. 6 is a cross-section A-A shown in FIG. 2 taken through a tip of one of the fan blades 230; and

[0059] FIG. 7 is a schematic flow chart illustrating a method of optimising a geared turbofan gas turbine engine.

DETAILED DESCRIPTION

[0060] 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, or cold 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.

[0061] 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.

[0062] 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.

[0063] 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.

[0064] 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.

[0065] 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.

[0066] 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.

[0067] 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.

[0068] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0069] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle.

[0070] 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.

[0071] FIG. 4 illustrates two plots of fan bypass efficiency as a function of fan inlet bypass flow rate at a constant fan speed, in which bypass flow rate is expressed as a percentage of maximum flow rate. A typical relation between these parameters for a conventional non-geared gas turbine engine is shown in a first curve 41, while a typical relation for a geared gas turbine engine is shown in a second curve 42. In each case, the position on the curve can be changed by changing the cold nozzle area. The two curves show quite different relationships, with the first curve 41 showing a sharper drop off in efficiency either side of an operating point 43. For the second curve 42, the drop off in efficiency is shallower, allowing for a greater degree of flexibility in designing the cold nozzle area. This realisation enables more than one design of cold nozzle area to be selected while retaining acceptable overall efficiency. Two possible operating points 44 and 45 are shown on the second curve, representing possible alternative operating points for the geared turbofan gas turbine engine.

[0072] As an example, for a geared turbofan gas turbine engine, altering the cold nozzle area to increase the specific thrust by 2% can result in the changes to efficiency and fuel burn, as shown in Table 1 below.

TABLE-US-00001 TABLE 1 Changes in efficiency and fuel burn for an example geared turbofan gas turbine engine for a 2% change in specific thrust. % % improvement improvement in Fuel burn in Fuel burn Fan Propulsive for 1000 nm for 500 nm efficiency efficiency mission mission % % (estimated) (estimated) Cruise +0.3 0.3 0.1% 0.1% Top of Climb +1.8 0.3 +0.25% 1.1% Maximum 0.4 0.4 N/A N/A Take-off

[0073] For a long range mission, this higher specific thrust cycle is unlikely to be better for overall fuel burn, but on short range missions (with possibly significant derate), the climb efficiency improvement may be beneficial overall. As shown in the above table, for a 1000 nm (1852 km) mission, the higher specific thrust variant is predicted to have a net 0.15% fuel burn improvement, but at 500 nm (926 km) the benefit is predicted to be approximately 1%. A higher specific thrust can therefore be beneficial overall if the engine is to be used primarily for short range missions.

[0074] Another way at looking at this is that the baseline cycle has similar efficiency as the +2% specific thrust cycle, but at a reduced stability risk (the static working line for the +2% specific thrust cycle is 2.5% higher).

[0075] More generally, this trade of cruise efficiency vs climb efficiency vs take off efficiency through changing nozzle areas for different missions will likely change for different engine designs. The advantage, however, is that the same engine and fan design can be used for more than one design by varying only those components needed to change the cold nozzle area.

[0076] FIG. 5 illustrates a modified version of the gas turbine engine 10 of FIG. 1, in which alternative cold nozzles are illustrated. The other components of the gas turbine engine 10 are the same as those in FIG. 1. A first option for changing the cold nozzle area is to change the dimensions of the outer nacelle 21, such that a rear portion 51a of the nacelle 21 in a first option defines a first smaller cold nozzle area and a rear portion 51b of the nacelle 21 in a second option defines a second larger cold nozzle area. An alternative way of defining the cold nozzle area is to alter the dimensions of the outer casing for the engine core 11, such that a rear portion 52a of the outer casing defines a first smaller cold nozzle area and a rear portion 52b of the outer casing defines a second larger cold nozzle area. In some cases adjustments to both the outer casing of the engine core 11 and the nacelle 21 may be made to change the cold nozzle area. An alternative way of achieving different sized cold nozzles would be to use a variable area nozzle (VAN), but this would add substantial weight and complexity to the engine, potentially outweighing any benefits of using different size cold nozzles to achieve different overall engine performance. The engine therefore preferably does not incorporate a VAN.

[0077] The fan 23 comprises individual fan blades 230. A cross-section A-A (indicated in FIG. 2) through a tip of one of the fan blades 230 is shown in FIG. 6. The cross-section may be at 90% of the blade span from the root (i.e. from the radially innermost gas-washed part of the fan blade 230).

[0078] The fan blade 230 has a leading edge 232, a trailing edge 234, a pressure surface 236 and a suction surface 238. The cross-section A-A also has a camber line 240. The camber line 240 is defined as the line formed by the points in the cross-section that are equidistant from the pressure surface 236 and the suction surface 238 for that cross-section. The cross-section A-A may be generated using a plane.

[0079] A line 90 is a projection into the cross-section A-A of a line that is parallel to the rotational axis 9 of the engine 10 (see FIG. 5). The line 90 passes through the leading edge 232 of the cross-section A-A. The angle between this line 90 and the tangent to the camber line 240 is shown in FIG. 6 as the blade tip angle . This angle may for example be within the range of 57 to 65 degrees. In the example shown in FIG. 6, the tangent to the camber line 240 that is used to define the angle is taken at the very leading edge 232 of the fan blade 230. However, in other arrangements, the tangent to the leading edge of the camber line 240 may be taken at any point within 5% of the total length of the camber line 240 from the leading edge 232. This means that blades having unusual leading edge curvature affecting the forwardmost 5% portion of the blade may still be within the defined range of blade tip angle , even if the tangent taken at the very leading edge 232 would not result in an angle falling within such a range. Purely by way of example, the blade tip angle of the fan blade 230 shown in FIG. 6 is around 60 degrees.

[0080] As noted elsewhere herein, in use the fan 23, and thus the fan blades 230, rotate about the rotational axis 9. At cruise conditions (as defined elsewhere herein), the fan rotates at a rotational speed , resulting in a linear velocity V.sub.ThetaBladeTip at the leading edge 232 of the blade tip 231 given by:

[00002]$V_{\mathrm{ThtaBladeTip}}=.Math..Math.\frac{D}{2}.Math.$

[0081] At least in part due to the rotation of the fan 230, air is ingested into the fan, resulting in a flow over the leading edge 232. The mean axial velocity of the flow at the leading edge 232 of the fan blade is shown as Vx.sub.air in FIG. 6. The vector sum of Vx.sub.air and (V.sub.ThetaBladeTip) gives the relative velocity V.sub.rel of the air at the leading edge 232 of the blade tip 231.

[0082] A fan tip air angle is shown in FIG. 6 and defined as:

[00003]$={\tan }^{-1}\left(\frac{V_{\mathrm{ThetaBladeTip}}}{{\mathrm{Vx}}_{\mathrm{air}}}\right)$

[0083] This fan tip air angle may be thought of as the angle between the vector representing Vx.sub.air (which is in an axial direction) and the vector representing the relative velocity V.sub.rel of the air at the leading edge 232 of the blade tip.

[0084] Gas turbine engines in accordance with some aspects of the present disclosure may have a fan tip air angle in the ranges described and/or claimed herein, for example in the range of from 57 degrees to 62 degrees. Purely by way of example, the fan tip air angle of the fan blade 230 shown in FIG. 6 is on the order of 60 degrees at cruise conditions of the gas turbine engine 10.

[0085] Gas turbine engines having fan tip air angles and/or blade tip angles in the ranges outlined herein may provide various advantages, such as improving the bird strike capability whilst retaining the efficiency advantages associated with geared and/or low specific thrust gas turbine engines, and may allow greater design freedom in other aspects of the fan system (including fan blades), such as weight, aerodynamic design, complexity and/or cost. Optimising the fan tip air angle may therefore further optimise the ability to design an engine for which more than one cold nozzle area may be selected depending on the expected range of travel.

[0086] FIG. 7 is a schematic flow diagram for a method 70 of optimising a geared turbofan gas turbine engine for an aircraft. In a first step 71, the expected service parameters for the aircraft on which the engine is to be used are determined. In a second step 72, a decision is made depending on whether the service parameters indicate a short range of travel or a long range of travel. If the range of travel is within a first smaller range, components are selected (step 73) to define a first smaller cold nozzle. If the range of travel is within a second larger range, components are selected (step 74) to define a second larger cold nozzle. The components selected may include a nacelle, or a portion thereof, an outer casing for the engine core, or a portion thereof, or a combination of the two. In a final step 75, the geared turbofan gas turbine engine is assembled, and may then be installed on an aircraft.

[0087] 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.