Fan design

Abstract

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

Claims

1. A gas turbine engine for an aircraft 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; 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: at engine cruise conditions, a fan tip air angle is in the range: 57 degrees62 degrees, the fan tip air angle being defined as: $={\tan }^{-1}\left(\frac{V_{\mathrm{ThetaBladeTip}}}{{\mathrm{Vx}}_{\mathrm{air}}}\right)$ where: $V_{\mathrm{ThetaBladeTip}}=.Math..Math.\frac{D}{2}.Math.;$ is fan rotational speed in radians/second; D is the diameter of the fan in meters at its leading edge; and Vx.sub.air is the mean axial velocity of the flow into the fan over the leading edge.

2. A gas turbine engine according to claim 1, wherein the fan tip air angle is in the range: 57 degrees60 degrees.

3. A gas turbine engine according to claim 1, wherein a specific thrust is defined as net engine thrust divided by mass flow rate through the engine, and at engine cruise conditions the specific thrust is in the range of from 70 Nkg.sup.1 s to 100 Nkg.sup.1 s.

4. A gas turbine engine according to claim 3, wherein the specific thrust at cruise conditions is in the range of from 75 Nkg.sup.1 s to 95 Nkg.sup.1 s, optionally 75 Nkg.sup.1 s to 90 Nkg.sup.1 s.

5. A gas turbine engine according to claim 1, wherein: a fan blade tip angle is defined as the angle between the tangent to the leading edge of the camber line in a cross-section through the fan blade at 90% of the blade span from the root and a projection of the axial direction onto that cross-section; and the value of (fan tip air angle fan blade tip angle ) is in the range of from 0 degrees to 3 degrees.

6. A gas turbine engine according to claim 1, wherein: a fan blade tip angle is defined as the angle between the tangent to the leading edge of the camber line in a cross-section through the fan blade 90% of the blade span from the root and a projection of the axial direction onto that cross-section, the fan blade tip angle being in the range of from 57 to 65 degrees.

7. A gas turbine engine according to claim 6, wherein the blade tip angle is in the range of from 58 to 64 degrees.

8. A gas turbine engine according to claim 1, wherein the fan blades comprise a main body attached to a leading edge sheath, the main body and the leading edge sheath being formed using different materials.

9. A gas turbine engine according to claim 8, wherein the leading edge sheath material has better impact resistance than the main body material.

10. A gas turbine engine according to claim 8, wherein the leading edge sheath material comprises Titanium.

11. A gas turbine engine according to claim 8, wherein the main body material comprises carbon fibre or an aluminium alloy.

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

13. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades, wherein, at engine cruise conditions: a specific thrust, defined as net engine thrust divided by mass flow rate through the engine, is in the range of from 70 Nkg.sup.1 s to 100 Nkg.sup.1 s; and a fan tip air angle is in the range: 57 degrees62 degrees, the fan tip air angle being defined as: $={\tan }^{-1}\left(\frac{V_{\mathrm{ThetaBladeTip}}}{{\mathrm{Vx}}_{\mathrm{air}}}\right)$ where: $V_{\mathrm{ThetaBladeTip}}=.Math..Math.\frac{D}{2}.Math.;$ is fan rotational speed in radians/second; D is the diameter of the fan in meters at its leading edge; and Vx.sub.air is the mean axial velocity of the flow into the fan over the leading edge.

14. A gas turbine engine for an aircraft 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; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein: a fan blade tip angle is defined as the angle between the tangent to the leading edge of the camber line in a cross-section through the fan blade at 90% of the blade span from the root and a projection of the axial direction onto that cross-section, and the fan blade tip angle is in the range of from 57 to 65 degrees.

15. A gas turbine engine according to claim 1, wherein a quasi-non-dimensional mass flow rate Q is defined as: $Q=W\frac{\sqrt{T0}}{P0.Math.A_{\mathrm{fan}}}.$ where: W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of the air at the fan face in Kelvin; P0 is average stagnation pressure of the air at the fan face in Pa; A.sub.fan is the area of the fan face in m.sup.2, and at engine cruise conditions: 0.029 Kgs.sup.1 N.sup.1 K.sup.1/2Q0.036 Kgs.sup.1 N.sup.1 K.sup.1/2.

16. A gas turbine engine according to claim 1, wherein a fan tip loading is defined as dH/Utip.sup.2, where dH is the enthalpy rise across the fan and Utip is the translational velocity of the fan blades at the tip of the leading edge, and at cruise conditions, 0.28 Jkg.sup.1 K.sup.1/(ms.sup.1).sup.2<dH/Utip.sup.2<0.36 Jkg.sup.1 K.sup.1/(ms.sup.1).sup.2.

17. A gas turbine engine according to claim 1, wherein: a fan pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit to the mean total pressure of the flow at the fan inlet, is no greater than 1.5 at cruise conditions, optionally in the range of from 1.35 to 1.45; and/or a fan root pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core to the mean total pressure of the flow at the fan inlet, is in the range of from 1.18 to 1.25 at cruise conditions, wherein, optionally, the ratio between the fan root pressure ratio to a fan tip pressure ratio at cruise conditions is no greater than 0.95, where the fan tip pressure ratio is defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct to the mean total pressure of the flow at the fan inlet.

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

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

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

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

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. 4 is a cross-section through the tip region of a fan blade of a gas turbine engine in accordance with the present disclosure; and

(6) FIG. 5 is a cross-section through the tip region of a fan blade of a gas turbine engine in accordance with the present disclosure.

(7) 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.

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

(9) In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. A throttle 161 is provided to control the fuel supply to the combustor. The amount of fuel supplied is dependent on the throttle position. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

(10) 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.

(11) 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.

(12) 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.

(13) 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.

(14) 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.

(15) 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.

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

(17) Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 18 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. 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.

(18) 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.

(19) 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. 4. 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).

(20) The fan blade 230 has a tip 231, 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 as defined elsewhere herein.

(21) 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. 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. 4 as the blade tip angle . This angle may be in the ranges defined and/or claimed herein, for example in the range of from 57 to 65 degrees. In the FIG. 4 example, 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 23. 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 ranges 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, blade tip angle of the fan blade 230 shown in FIG. 4 is on the order of 60 degrees.

(22) 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:

(23) $V_{\mathrm{ThetaBladeTip}}=.Math..Math.\frac{D}{2}.Math.$

(24) 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. 4. 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.

(25) A fan tip air angle is shown in FIG. 4 and defined as:

(26) 0$={\tan }^{-1}\left(\frac{V_{\mathrm{ThetaBladeTip}}}{{\mathrm{Vx}}_{\mathrm{air}}}\right)$

(27) 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 231.

(28) 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. 4 is on the order of 60 degrees at cruise conditions of the gas turbine engine 10.

(29) The fan blades 230 may be manufactured using any suitable material or combination of materials, as described elsewhere herein. Purely by way of further example, FIG. 5 shows a fan blade 330 that is the same as the fan blade 230 described above (for example in relation to fan tip air angle and/or blade tip angles ), but has a main body 350 attached to a leading edge sheath 360. The main body 350 and the leading edge 360 in the FIG. 5 example are manufactured using different materials. Purely by way of example, the main body 350 may be manufactured using a carbon fibre composite material or an aluminium alloy material (such as an aluminium lithium alloy), and the leading edge sheath 360 may be manufactured from a material that is better able to withstand being struck by a foreign object (such as a bird). Again, purely by way of example, the leading edge sheath may be manufactured using a titanium alloy.

(30) As explained elsewhere herein, gas turbine engines having fan tip air angles 8 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. This may allow greater design freedom in other aspects of the fan system (including fan blades), such as weight, aerodynamic design, complexity and/or cost.

(31) A further example of a feature that may be better optimized for gas turbine engines 10 according to the present disclosure compared with conventional gas turbine engines is the intake region, for example the ratio between the intake length L and the fan diameter D. Referring to FIG. 1, the intake length L is defined as the axial distance between the leading edge of the intake and the leading edge of the tip of the fan blades, and the diameter D of the fan 23 is defined at the leading edge of the fan 23. Gas turbine engines 10 according to the present disclosure, such as that shown by way of example in FIG. 1, may have values of the ratio L/D as defined herein, for example less than or equal to 0.45. This may lead to further advantages, such as installation and/or aerodynamic benefits.

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

(33) 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.