GAS TURBINE ENGINE

Abstract

Gas turbine aircraft engine comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor; and a fan upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of tandem fan blades. Each fan blade comprises a main blade and an auxiliary blade. Over substantially all of the auxiliary blade's radial span, the leading edge of the auxiliary blade is rearwards of the closest point on the trailing edge of the main fan blade, and on a given aerofoil chordal section of the main fan blade, the leading edge position of an aerofoil chordal section of the auxiliary fan blade lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate in tandem.

Claims

1. A gas turbine engine for an aircraft, the gas turbine engine 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 and driven by the turbine, the fan comprising a circumferential row of tandem fan blades; wherein each of the tandem fan blades comprises a main fan blade and an auxiliary fan blade positioned at the rear of the main fan blade, such that, over substantially all of the auxiliary fan blade's radial span, the leading edge of the auxiliary fan blade is rearwards of the closest point on the trailing edge of the main fan blade, and on a given aerofoil chordal section of the main fan blade, the leading edge position of an aerofoil chordal section of the auxiliary fan blade lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate in tandem; and wherein the auxiliary fan blade is movable within a range of pitch angles relative to the main fan blade.

2. The gas turbine engine of claim 1, wherein the auxiliary fan blade is positioned such that, over substantially all of its radial span, the leading edge of the corresponding aerofoil chordal section of the auxiliary fan blade lies on a rearwards extension of the camber line of a respective aerofoil chordal section of the main fan blade.

3. The gas turbine engine of claim 1, wherein the auxiliary fan blade is positioned such that, over substantially all of its radial span, the leading edge of the auxiliary fan blade is rearwards of the closest part of the trailing edge of the main fan blade.

4. The gas turbine engine of claim 1, further comprising an annular splitter positioned behind the fan and configured to split an air flow generated by the fan into a core airflow that flows through the engine core and a bypass airflow that flows through a bypass duct surrounding the engine core, wherein the auxiliary fan blades have radially outer tips which are inside or level with the radial position of the annular splitter.

5. The gas turbine engine of claim 1, wherein, in side elevational view, the projected area of the auxiliary fan blade of each tandem fan blade is less than the projected area of the respective main fan blade.

6. The gas turbine engine of claim 1, wherein the main fan blade and the auxiliary fan blade are configured so that, in side elevational view, the trailing edge of the main fan blade has the appearance of a cut-out region at the rear of the base of the main fan blade, which cut-out region embraces the auxiliary fan blade.

7. The gas turbine engine of claim 1, wherein the auxiliary fan blade is movable within a range of pitch angles of at least 10°, whereby increasing the pitch angle increases the overall camber angle of the given aerofoil chordal sections of the main fan blade and the auxiliary fan blade by a corresponding amount.

8. The gas turbine engine of claim 1, wherein moving the auxiliary fan blades across the total range of pitch angles varies the total work done by the fan by at least 20%.

9. The gas turbine engine of claim 1, wherein the auxiliary fan blade is configured to contribute at least 20% of the total work done by the tandem fan blades at cruise operating condition of the gas turbine engine.

10. The gas turbine engine of claim 1, wherein the main fan blade has a radial span extending from a hub at a 0% span position to a tip at a 100% span position, and wherein the average camber angle of the radially innermost 10% of the radial span of the main fan blade is less than 75% of the average camber angle of the 10% portion of the radial span of the main fan blade that has the maximum average camber angle.

11. The gas turbine engine of claim 1, further comprising: 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.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[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 of FIG. 1;

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

[0065] FIG. 4 is a schematic side elevational view of a tandem fan blade for use with the gas turbine engine of FIGS. 1 and 2;

[0066] FIG. 5A is a schematic chordal section of the tandem fan blade of FIG. 4 on line P.sub.A-P.sub.A in a first configuration;

[0067] FIG. 5B is a schematic chordal section of the tandem fan blade of FIG. 4 on line P.sub.B-P.sub.B;

[0068] FIG. 6 is a schematic chordal section of the tandem fan blade of FIG. 4 on line P.sub.A-P.sub.A in a second configuration; and

[0069] FIG. 7 is a schematic side elevational view of a variant tandem fan blade.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0070] Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

[0071] 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 having a circumferential row of tandem fan blades 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 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0072] 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 core exhaust 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 blades 23 generally provide the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0073] 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 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.

[0074] 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) 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). 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 may be referred to as a first, or lowest pressure, compression stage.

[0075] 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.

[0076] 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 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.

[0077] 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.

[0078] 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.

[0079] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0080] 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 that is separate to and radially outside the core exhaust 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.

[0081] 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.

[0082] The fan, as shown in FIGS. 1 and 2, comprises a plurality of tandem fan blades 23 attached to a fan rotor disc of a fan hub 50, and arranged in a circumferential row around the fan hub 50. Each of the tandem fan blades comprises a main fan blade 23a and an auxiliary fan blade 23b positioned rearwards of the main fan blades 23a, such that on a given aerofoil chordal section of the main fan blade, the leading edge position of an aerofoil chordal section of the auxiliary fan blade 23b lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade 23a. Moreover, the main fan blade 23a and the auxiliary fan blade 23b in each of the tandem fan blades 23 are arranged to rotate in tandem so that they retain their angular positions relative to each other.

[0083] Each auxiliary fan blade 23b is pivotable about its leading edge, such that it is movable within a range of pitch angles relative to the main fan blade 23a. The pivoting movement of the auxiliary fan blade is actuated by a gear-based variable pitch mechanism 52 provided in the fan hub 50.

[0084] FIG. 4 shows a schematic side elevational view of the tandem fan blade 23 with respect to the principal rotational axis 9. In this view, the main fan blade 23a projects a larger area than the auxiliary fan blade 23b. The main fan blade 23a also has a greater span in the radial direction R of the engine than the auxiliary fan blade 23b. Specifically, the main fan blade 23a extends from a hub 62 to a tip 64 adjacent the inner surface of the nacelle 21, while the auxiliary fan blade 23b extends from a hub 72 to a tip 74 which is at or inwards of the radial position of the leading edge of an annular splitter 51 located behind the fan, the splitter splitting the air flow generated by the fan into the core airflow A the bypass airflow B. Thus the auxiliary fan blades 23b have little or no effect on the bypass airflow B.

[0085] The main fan blade 23a and the auxiliary fan blade 23b have respective leading edges 66, 76 and trailing edges 68, 78. The main fan blade 23a has a root portion, such as a dovetail (not shown), which engages with a complimentary formation formed in the fan rotor disc to removably mount the blade to the disc.

[0086] As mentioned above, the auxiliary fan blade 23b is pivotable about its leading edge 76 by the variable pitch mechanism 52. More specifically, the auxiliary fan blade 23b is pivotable about rotational axis 56, shown as a dotted line in FIG. 4. There is a small clearance between the trailing edge 68 of the main fan blade 23a and the leading edge 76 of the auxiliary fan blade 23b. This ensures the leading edge 76 of the auxiliary fan blade 23b does not interfere with the trailing edge 68 of the main fan blade 23a when the auxiliary fan blade pivots. The clearance may provide a gap between the leading edge of the auxiliary fan blade and the trailing edge of the main fan blade which, measured in the axial direction of the engine, has a width which is no more than 10%, and preferably no more than 5%, of the maximum axial spacing between the leading and trailing edges of the main fan blade.

[0087] Schematic aerofoil chordal sections of the tandem fan blade 23 of FIG. 4 on line P.sub.A-P.sub.A are shown in FIG. 5A. The line P.sub.A-P.sub.A is within the first 10% of the span of the main fan blade 23a as measured from its hub 62. In contrast, a schematic aerofoil chordal section of the main fan blade 23a of FIG. 4 on line P.sub.B-P.sub.B is shown in FIG. 5B. The line P.sub.B-P.sub.B is within the first 40% of the span of the main fan blade 23a as measured from its hub 62, but is above the tip 74 of the auxiliary fan blade 23b and hence no aerofoil chordal section for the auxiliary fan blade is displayed in FIG. 5B.

[0088] A combined camber line 80 of the tandem fan blade 23 on line P.sub.A-P.sub.A is shown in FIG. 5A, and a camber line 82 of just the main fan blade 23a on line P.sub.B-P.sub.B is shown in FIG. 5B. The combined camber line 80 is a combination of the camber lines of the main fan blade 23a and the auxiliary fan blade 23b on line P.sub.A-P.sub.A. More specifically, in FIG. 5A the camber line of the aerofoil chordal section of the auxiliary fan blade 23b is a rearwards continuation of the camber line of the aerofoil chordal section of the main fan blade 23a. As such the air flow over the main fan blade 23a continues with little disturbance over the auxiliary fan blade 23b. A combined camber line of the type shown in FIG. 5A, i.e. such that the leading edge position of the aerofoil chordal section of the auxiliary fan blade 23b lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade 23a, is provided over substantially the entire radial span of the auxiliary fan blade 23b.

[0089] In FIG. 5A the angle between the direction of the combined camber line 80 at the leading edge position of the main fan blade aerofoil chordal section and the axial direction is α.sub.1, and the angle between the direction of the combined camber line at the trailing edge position of the auxiliary fan blade aerofoil chordal section and the axial direction is α.sub.2. The overall camber angle of the aerofoil chordal sections, therefore, is (α.sub.1−α.sub.2).

[0090] In FIG. 5B the angle between the direction of the camber line 82 at the leading edge position of the main fan blade aerofoil chordal section and the axial direction is β.sub.1, and the angle between the direction of the camber line at the trailing edge position of the main fan blade aerofoil chordal section and the axial direction is β.sub.2. Thus, the camber angle is (β.sub.1−β.sub.2).

[0091] The main fan blade 23a as shown in FIGS. 5A and 5B is a decambered fan blade. For example, the average camber angle of the radially innermost 10% of the radial span of the main fan blade is less than 75% of the average camber angle of the 10% portion of the radial span of the main fan blade that has the maximum average camber angle. The fan hub pressure ratio at the foot of the main fan blade may thus be less than 1.2.

[0092] During aircraft take-off, the gas turbine engine provides the maximum amount of thrust during a flight. This is an operating condition which is typically limited by the T30 delivery temperature of the high-pressure compressor 15 to the combustor 16. Therefore it is desirable to limit the fan hub pressure ratio in order to reduce this contribution to the T30 temperature. For example, as illustrated in FIG. 5A, the auxiliary fan blade 23b can be pivoted to its lowest pitch angle, which results in a low overall camber angle (α.sub.1−α.sub.2). As a result, the auxiliary fan blade 23b does not significantly increase the fan hub pressure ratio.

[0093] However, during e.g. cruise operating condition, which is not limited by the T30 delivery temperature, the auxiliary fan blade 23b can be pivoted to a pitch angle which results in a higher overall camber angle (α.sub.1−α.sub.2), as shown in FIG. 6 which are the same schematic chordal sections as FIG. 5A but with the auxiliary fan blade 23b pivoted about its leading edge 76. In this way, the auxiliary fan blade 23b supplements the flow turning capability at the foot of the main fan blade 23a, and thereby increases the fan hub pressure ratio from 1.2 to 1.3 or more. This leads to an increase in T30 temperature and overall pressure ratio, and accordingly improves the efficiency of the gas turbine engine. The auxiliary fan blades 23b can be configured to contribute at least 20% of the total work done by the tandem fan blades 23 at cruise operating condition of the engine. Calculations suggest that use of the auxiliary fan blades during cruise can lead to about a 20° C. increase in T30 temperature and a significant improvement in specific fuel consumption (SFC)

[0094] Typically, the auxiliary fan blade 23b can be pivoted through a range of pitch angles of about 15°, with a pitch angle of 0° corresponding to the lowest overall camber angle (e.g. (α.sub.1−α.sub.2) of about 10°) and a pitch angle of 15° corresponding to the highest overall camber angle (e.g. (α.sub.1−α.sub.2) of about 25°). At a pitch angle of 15° the auxiliary fan blade 23b essentially acts as a booster for the low pressure compressor 14. Moving the auxiliary fan blades across the total range of pitch angles can vary the work done by the fan by about 20%.

[0095] In a variant tandem fan blade (not shown), the pitch angle of the auxiliary fan blade is fixed, e.g. non-adjustable. In this variant tandem fan blade, the pitch angle of the auxiliary fan blade can be set to provide an overall camber angle which balances the desire for an increased fan foot pressure ratio at cruise operating condition with a suitably limited T30 temperature at take-off. As the variant fan blade does not require a variable pitch mechanism, it can help to reduce the overall weight of the fan.

[0096] FIG. 7 shows a schematic side elevational view of a variant tandem fan blade 123, comprising a main fan blade 123a and an auxiliary fan blade 123b. In this view, the trailing edge of the main fan blade 123a has the appearance of a cut-out region at the rear of the base of the main fan blade 123a, which cut-out region embraces the auxiliary fan blade. For example, the cut-out region is sized so that the trailing edge 168 of the main fan blade radially outward of the auxiliary blade forms an extension of the line of the trailing edge 178 of the auxiliary fan blade. This can help to reduce unwanted flow disturbances at the tip 174 of the auxiliary blade. In the side elevational view, the tandem fan blade 123 exhibits a similar profile as a decambered mono fan blade.

[0097] 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.