GB2594712A

GB2594712A - Vane assembly for gas turbine engine

Info

Publication number: GB2594712A
Application number: GB2006526.4A
Authority: GB
Inventors: Namgoong Howoong
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2020-05-04
Filing date: 2020-05-04
Publication date: 2021-11-10
Also published as: GB202006526D0

Abstract

A vane assembly for a gas turbine engine comprising two rows of outlet guide vanes (Preferably fan outlet guide vanes in a bypass duct) wherein a first row of vanes 530 has a first sweep angle α, and a second row of vanes 540 has a second sweep angle β such that at least one first row vane and at least one second row vane overlap axially relative to the central longitudinal axis 515. The sweep of the vanes is intended to mitigate interactions between the vanes and the fan (Relative to straight radial vanes), while the crossed overlap provides greater structural strength than a single row of swept vanes. The vanes may overlap at about the vane midsections, at the vane tips or at the vane roots. The vanes may be angled at a range of ±0-70°. The second row of vanes may have alternately axially offset root positions, or have at least one second vane proximate a pylon.

Description

VANE ASSEMBLY FOR GAS TURBINE ENGINE

FIELD OF THE DISCLOSURE

The present disclosure relates to a vane assembly, and in particular to a vane 5 assembly for a gas turbine engine.

BACKGROUND

A gas turbine engine generally includes a fan mounted upstream of a core of the gas turbine engine. In order to improve fuel efficiency and reduce operational noise, current gas turbine engines are provided with a bypass duct. The bypass duct generally includes an outlet guide vane (OGV) assembly. Outlet guide vanes (OGVs) of the OGV assembly may guide fan airflow discharged from the fan to a desired direction prior to the fan airflow being exhausted from the bypass duct.

In addition to guiding the fan airflow, the OGV assembly may also provide structural stiffness to a fan casing. In some cases, a certain gap between the fan and the OGVs may be required to reduce fan and OGV interactions. It may be desirable to provide sufficient structural stiffness to the fan casing while reducing the fan and OGV interactions.

Straight (or radial) OGVs are often used in conventional gas turbine engines. Straight OGVs may provide adequate structural stiffness to the fan casing. However, straight OGVs may generate undesirable high fan and OGV 25 interactions.

To reduce the fan and OGV interactions, swept OGVs may be used. Swept OGVs may reduce fan and OGV interactions. However, structural integrity/stiffness of the swept OGVs may be weaker compared to the straight OGVs. Furthermore, a swept OGV configuration may increase stress concentration near the fan casing and a hub to which the OGVs are coupled to, compared to the straight OGVs.

Therefore, it may be desirable to design thin OGVs in order to reduce aerodynamic losses due to the fan and OGV interactions. However, structural integrity of a thin and swept OGV may not be adequate for an aircraft travelling at near transonic speeds.

Furthermore, as there is a trend to increase bypass ratio of gas turbine engines, there is a concomitant increase in fan diameter. The increase in fan diameter may present difficulty in designing a thin and swept OGV while maintaining structural integrity of the design.

SUMMARY

In a first aspect, there is provided a vane assembly for a gas turbine engine. The vane assembly includes a hub defining a central axis. The vane assembly further includes a plurality of vanes extending from the hub. Each vane has a leading edge and a trailing edge. Each vane extends from a root disposed at the hub to a tip. The leading edge of each vane defines a corresponding sweep angle relative to a corresponding baseline position. The plurality of vanes includes a set of first vanes circumferentially spaced from one another around the hub. The plurality of vanes further includes a set of second vanes circumferentially spaced from one another around the hub. The leading edge of each first vane is upstream of the leading edge of each second vane at the respective roots. The sweep angle of at least one first vane is different from the sweep angle of at least one second vane such that the at least one first vane at least partially and axially overlaps with the at least one second vane relative to the central axis.

The first vanes and the second vanes may form a crossed configuration. The crossed configuration may be advantageous from an aeromechanical point of view. The crossed configuration may allow thinner swept OGVs to provide similar structural strength as compared to traditional straight (or radial) OGVs. The crossed configuration may also generate lower fan and OGV interactions compared to the straight OGVs.

In some embodiments, the at least one first vane is swept backward such that the sweep angle of the at least one first vane is greater than -70 degrees and less than 0 degree.

In some embodiments, the at least one second vane is swept forward such that the sweep angle of the at least one second vane is greater than 0 degree and less than about 70 degrees.

In some embodiments, the at least one first vane axially overlaps with the at 10 least one second vane at around 50% of radial spans between the respective roots and the respective tips.

In some embodiments, the at least one first vane axially overlaps with the at least one second vane proximate the respective tips.

In some embodiments, the trailing edge of each first vane is upstream of the leading edge of each second vane at the respective roots.

In some embodiments, the at least one first vane is swept forward such that the 20 sweep angle of the at least one first vane is greater than 0 degree and less than about 70 degrees.

In some embodiments, the at least one second vane is swept backward such that the sweep angle of the at least one second vane is greater than -70 degrees 25 and less than 0 degree.

In some embodiments, the at least one first vane axially overlaps with the at least one second vane proximate the respective roots.

In some embodiments, the first vanes and the second vanes are swept forward and the sweep angle of the first vanes is different from the sweep angle of the second vanes.

In some embodiments, the first vanes and the second vanes are swept backward and the sweep angle of the first vanes is different from the sweep angle of the second vanes.

In some embodiments, at least two first vanes of the set of first vanes have different sweep angles.

In some embodiments, at least two second vanes of the set of second vanes have different sweep angles.

The first vanes and the second vanes having different sweep angles may allow a higher degree of freedom while designing the vane assembly and thereby may allow the vane assembly to be designed for a wider variety of engine types and models.

In some embodiments, at least two first vanes of the set of first vanes are axially offset from one another at the respective roots In some embodiments, at least two second vanes of the set of second vanes are 20 axially offset from one another at the respective roots.

In some embodiments, at least one second vane of the set of second vanes is circumferentially disposed proximate a pylon. The at least one second vane is axially positioned upstream of the other second vanes at the respective roots.

In some embodiments, adjacent second vanes of the set of second vanes are axially offset from one another at the respective roots. Alternating second vanes of the set of second vanes are arranged at an identical axial position at the respective roots.

Providing an axial offset to the first vanes and/or the second vanes may aid in arrangement of the vanes around certain structural components (for example, the pylon or bifurcations), thereby allowing a higher degree of freedom for mitigating potential effects from the bifurcations and the pylon.

In some embodiments, the roots of the set of first vanes are arranged at an identical first axial position.

In some embodiments, the roots of the set of second vanes are arranged at an identical second axial position.

In some embodiments, each of the plurality of vanes is a fan outlet guide vane.

In a second aspect, there is provided a gas turbine engine for an aircraft. The gas turbine engine includes the vane assembly of the first aspect.

In a third aspect, there is provide an aircraft including the gas turbine engine of the second aspect.

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise 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. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star gearbox having a ratio in the range of from 3.1 01 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, a combustor may 5 be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided The combustor may be 10 provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any plafform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further nonlim itative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/U2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied 5 by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values being dimensionless). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the 10 range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is 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 cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5,13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine core. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg-ls, 105 Nkg-ls, 100 Nkg-ls, 95 Nkg- 15, 90 Nkg-ls, 85 Nkg-ls or 80 Nkg-ls. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg-ls to 100 Nkg-1s, or 85 Nkg-ls to 95 Nkg-ls. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to 420 kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K. The maximum TEl may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc.

By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be 5 provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have any 10 desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this regard, m id-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint -in terms of time and/or distance -between top of climb and start of descent. Cruise conditions thus define an operating point of, the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide -in combination with any other engines on the aircraft -steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the m id-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the m id-cruise of the aircraft, as defined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine 5 engine; Figures 4, 4B and 4C are schematic side views of exemplary OGVs with different sweep angles; Figure 5A is a schematic side view of a vane assembly in accordance with an embodiment of the present disclosure; Figure 5B is a schematic perspective view of the vane assembly in Figure 5A in accordance with an embodiment of the present disclosure; Figures 6A, 6B, 6C, 6D, 6E, 6F, 6G, 6H, 61 are schematic side views of various embodiments of the vane assembly; Figure 7A is a schematic plan view taken at roots of vanes of an 15 embodiment of the vane assembly; Figure 7B is a schematic plan view taken at roots of vanes of another embodiment of the vane assembly, and Figure 7C is a schematic plan view taken at roots of vanes of yet another embodiment of the vane assembly.

DETAILED DESCRIPTION

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.

Figure 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 that flows through a core duct 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.

In use, the core airflow A is accelerated and compressed by the low pressure 5 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 10 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 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2. The low pressure turbine 19 (see Figure 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to process around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

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.

The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 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 Figure 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.

The epicyclic gearbox 30 illustrated by way of example in Figures 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.

It will be appreciated that the arrangement shown in Figures 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 Figure 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 Figure 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 Figure 2.

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.

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

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 Figure 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 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.

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 Figure 1), and a circumferential direction (perpendicular to the page in the Figure 1 view). The axial, radial and circumferential directions are mutually perpendicular.

In addition, the present invention is equally applicable to aero gas turbine 5 engines, marine gas turbine engines and land-based gas turbine engines.

Gas turbine engines include various vane assemblies. For example, a vane assembly may refer to an inlet guide vane (IGV) assembly. In other example, a vane assembly may refer to an outlet guide vane (OGV) assembly. In another example, a vane assembly may refer to an engine section stator vane (ESS) assembly.

Outlet guide vanes (OGVs) may be present in the bypass duct 22 and downstream of the fan 23. These vanes are generally referred to as Fan Outlet Guide Vanes (FOGVs) In the present disclosure, unless otherwise stated, references to outlet guide vanes are references to fan outlet guide vanes, The OGVs may turn and/or direct the bypass airflow B axially in the bypass duct 22 In addition to OGVs, many fan casing assemblies include one or more (frequently two, diametrically opposed) dividing structures, often called "bifurcations" (not shown). The bifurcations may divide the annular space defined by the bypass duct 22 into two semi-annular spaces. These dividing structures are typically hollow duct-like structures through which various mechanical, electrical, pneumatic, hydraulic, or other connections (including structural supports) can pass without causing disruption to the bypass airflow B through the bypass duct 22. The bifurcations "fair" or guide the bypass airflow B in aerodynamic fashion around these structures. In a typical installation of the gas turbine engine 10 under a wing of an aircraft (not shown), the upper bifurcation houses the engine mounts and various electrical, hydraulic, and pneumatic systems while the lower bifurcation houses oil drains and the like.

Due to structures such as a pylon (used to mount the gas turbine engine 10 to a wing of an aircraft) and the bifurcations, there may be a reduction in degree of freedom in the arrangement of OGVs. Therefore, there may be a need for a vane assembly which may allow a higher degree of freedom for mitigating potential effects from the bifurcations and the pylons.

Figures 4A, 4B and 4C illustrate exemplary OGV assemblies 400A, 400B and 400C. Each of the OGV assemblies 400A, 400B, 400C includes a hub 410 on which respective vanes 420A, 420B and 420C (only one shown in Figures 4A, 4B and 4C) are circumferentially disposed. Each of the vanes 420A, 420B, 420C includes a corresponding leading edge 421A, 421B. 421C and a corresponding trailing edge 422A, 422B, 422C.

In the OGV assembly 400A, the vanes 420A are radial vanes. The leading edge 421A of each of the vanes 420A of the OGV assembly 400A defines a corresponding baseline position 450. The leading edge 421A of the vanes 420A may be substantially perpendicular to a central axis 415 defined by the hub 410.

In the OGV assembly 400B, the vanes 420B of the vane assembly 400B are swept backward relative to the corresponding baseline position 450 (shown in dashed lines in Figure 4B). Further, the leading edge 421B of each of the vanes 420B defines a corresponding sweep angle SA1 relative to the corresponding 20 baseline position 450. The angle SA1 is defined between the leading edge 421A at the baseline position 450 (i.e., the radial position) and the leading edge 421B at the swept position. The angle SA1 is defined along the central axis 415.

In the OGV assembly 400C, the vanes 420C of the OGV assembly 400C are 25 swept forward relative to the corresponding baseline position 450 (shown in dashed lines in Figure 4C). Further, the leading edge 421C of each of the vanes 420C defines a corresponding sweep angle SA2 relative to the corresponding baseline position 450. The angle SA2 is defined between the leading edge 421A at the baseline position 450 (i.e., the radial position) and the leading edge 421C 30 at the swept position. The angle SA2 is defined along the central axis 415.

Furthermore, the vanes 420A, 420B, 420C may be leaned. More specifically, the leading edges 421A, 421B and 421C of the respective vanes 420A, 420B and 420C may be leaned. In some embodiments, the vanes 420A, 420B, 420C may also be staggered. It may be noted that due to the lean and a complex aerofoil structure of vanes, its leading edge generally may not lie on a straight line as illustrated in Figures 4A, 4B and 4C. The vanes 420A, 420B, 420C are for illustrative purpose only for showing a baseline position of a vane and forward and backward swept positions of the vane relative to the baseline position. Where the leading edges 421A, 421B, 421C do not lie on a straight line, the sweep angles SA1, SA2 may be defined as mean values between root and tip.

OGVs can be forwardly or backwardly swept relative to a baseline position. The baseline position corresponds to a radial or straight configuration. An OGV that is forwardly swept is inclined axially forward or upstream relative to the baseline position. An OGV that is backwardly swept is inclined axially backward or downstream relative to the baseline position. Hereinafter, a forward sweep is considered to be positive, i.e., the sweep angle SA2 formed by a vane swept forward forms a positive sweep angle +SA2. Similarly, a backward sweep is considered to be negative, i.e., the sweep angle SA1 formed by a vane swept backward forms a negative sweep angle -SA1.

Referring to Figures 5A and 5B, an embodiment of a vane assembly 500 for the gas turbine engine 10 (shown in Figure 1) is illustrated. In this embodiment, the vane assembly 500 is a fan outlet guide vane (FOGV) assembly. However, in some embodiments, the vane assembly 500 may be a vane assembly of a turbine section. The vane assembly 500 may be incorporated into other sections of the gas turbine engine 10, including but not limited to, a compressor section. For example, the vane assembly 500 may be an Inlet Guide Vane (IGV) assembly, or an engine section stator (ESS) vane assembly, arranged in the core dud upstream of the compressor. The vane assembly 500 may also be suitable for various other vane assemblies used in the gas turbine engine 10.

The vane assembly 500 includes a hub 510. The hub 510 may have a substantially circumferential structure. The hub 510 defines a central axis 515. In some embodiments, the central axis 515 may coincide with the principal rotational axis 9 of the gas turbine engine 10 (shown in Figure 1). In some embodiments, the hub 510 may be part of an engine core casing (in the case of an FOGV assembly) defining an inner wall of the bypass duct 22. In some embodiments, the hub 510 may be a stator hub (in the case of an Inlet Guide Vane (IGV) assembly, or an engine section stator (ESS) vane assembly) defining an inner wall of the core duct.

The vane assembly 500 further includes a plurality of vanes 520. Each of the vanes 520 extend from a root 523 disposed at the hub 510 to a tip 524. The tip 524 may be distal to the root 523 with a radial span R defined therebetween.

The radial span R may be defined perpendicular to the central axis 515 of the hub 510. In the case of an FOGV assembly, the tip 524 of the vanes 520 may be attached to a fan casing of the nacelle 21 (shown in Figure 1). Each of the vanes 520 has a leading edge 521 and a trailing edge 522. The leading edge 521 and the trailing edge 522 of each of the vanes 520 may extend from the root 523 to the tip 524. The trailing edge 522 of each of the vanes 520 may be axially downstream of its leading edge 521 along the central axis 515 of the hub 510. Each of the vanes 520 may be swept backward and/or forward relative to a corresponding baseline position 650 (shown by dashed lines in Figure 5A). Each vane 520 defines a corresponding sweep angle a or 13 relative to the corresponding baseline position 650.

The vanes 520 may be mounted on the hub 510 using any suitable mounting method. For example, the vanes 520 may be mounted on the hub 510 by clamping the vanes 520 on the hub 510. In other example, the vanes 520 may be mounted and secured to the hub 510 using dovetail or T-slot fixing, with an anti-rotation feature to lock the vanes 520 in a required position. In some embodiments, the vanes 520 may be formed integrally with the hub 510.

The vanes 520 include a set of first vanes 530. The first vanes 530 are circumferentially spaced from one another around the hub 510. The vanes 520 further include a set of second vanes 540. The second vanes 540 are circumferentially spaced from one another around the hub 510. The first vanes 530 and the second vanes 540 may be swept forward and/or backward relative to the corresponding baseline position 650. In some embodiments, each first vane 530 may be swept backward by a corresponding sweep angle a. Further, each second vane 540 may be swept forward by a corresponding sweep angle p. A magnitude of the sweep angle a of each first vane 530 and a magnitude of the sweep angle [3 of each second vane 540 may be equal or different. Forward and backward swept vanes and corresponding sweep angles have been discussed above with reference to Figures 4A, 4B and 4C.

The first vanes 530 and the second vanes 540 are disposed on the hub 510 such that the leading edge 521 of each first vane 530 is upstream of the leading edge 521 of each second vane 540 at the respective roots 523. Further, the sweep angle a of at least one first vane 530 is different from the sweep angle p of at least one second vane 540 such that the at least one first vane 530 at least partially and axially overlaps with the at least one second vane 540 relative to the central axis 515.

In some embodiments, the at least one first vane 530 is swept backward such that the sweep angle a of the at least one first vane 530 is greater than -70 degrees and less than 0 degree. Further, the at least one second vane 540 is swept forward such that the sweep angle p of the at least one second vane 540 is greater than 0 degree and less than about 70 degrees. Various configurations of the first vanes 530 and the second vanes 540 may be possible with different sweep angles a and [3 of the corresponding first vanes 530 and second vanes 540. In the illustrated embodiment, multiple first vanes 530 are swept backward by the sweep angle a and multiple second vanes 540 are swept forward by the sweep angle 13. Further, the first and second vanes 530, 540 are disposed in an alternating manner along a circumferential direction CD. Specifically, one second vane 540 is circumferentially disposed between two adjacent first vanes 530. Further, one first vane 530 is circumferentially disposed between two adjacent second vanes 540.

In some embodiments, the at least one first vane 530 axially overlaps with the at least one second vane 540 at around 50% of the radial spans R between the respective roots 523 and the respective tips 524. Furthermore, the trailing edge 522 of each first vane 530 is upstream of the leading edge 521 of each second vane 540 at the respective roots 523. The first vanes 530 and the second vanes 540 may form a crossed configuration. The crossed configuration may be advantageous from an aeromechanic point of view. The crossed configuration may allow thinner swept OGVs to provide similar structural strength compared to conventional straight OGVs. The crossed configuration may also generate lower fan and OGV interactions compared to conventional straight OGVs. The crossed configuration is shown in Figures 5A, 5B, 6B, 6E and 6H.

In some embodiments, at least two first vanes 530 of the set of first vanes 530 have different sweep angles a. In some embodiments, each of the first vanes 530 may have different sweep angles a. In some embodiments, at least two second vanes 540 of the set of second vanes 540 have different sweep angles I* In some embodiments, each of the second vanes 540 may have different sweep angles p. This may allow a higher degree of freedom while designing the vane assembly 500 and thereby may allow the vane assembly 500 to be designed for a wider variety of engine types and models.

Now referring to Figures 6A, 6B and 6C, various embodiments of the vane assembly 500 are illustrated. Two extreme configurations 600A, 600C of the vane assembly 500 are illustrated in Figures 6A and 6C. The sweep of each of the vanes 520 is defined with respect to the baseline position 650 (shown by dashed line). Various other intermediate configurations of the vane assembly 500 may be possible between these two extreme configurations 600A, 600C.

An intermediate configuration 600B is shown in Figure 6B. In some embodiments, the at least one first vane 530 is swept forward such that the sweep angle a of the at least first vane 530 is greater than 0 degree and less than about 70 degrees. Further, the at least one second vane 540 is swept backward such that the sweep angle p of the at least one second vane 540 is greater than -70 degrees and less than 0 degree.

As shown in Figure 6A, in some embodiments, the at least one first vane 530 axially overlaps with the at least one second vane 540 proximate the respective roots 523. An axial overlap region 660 is shown in FIG. 6A. In the configuration 600A, the axial offset between the leading edges 521 at the respective roots 523 of the first vanes 530 and the second vanes 540 may be substantially less than an axial offset at the respective tips 524 of the first vanes 530 and the second vanes 540. Further, the first vane 530 is swept forward such that the sweep angle a has a positive value. The second vane 540 is swept backward such that the sweep angle i3 has a negative value. In some embodiments, an axial offset at the respective roots 523 of the first vanes 530 and the second vanes 540 may be about half of an axial span of the vanes 520.

As shown in Figure 6B, in some embodiments, the at least one first vane 530 axially overlaps with the at least one second vane 540 at about 50% of the radial spans R between the respective roots 523 and the respective tips 524. An axial overlap region 670 is shown in FIG. 6B. In the configuration 600B, the trailing edge 522 of each first vane 530 is upstream of the leading edge 521 of each second vane 540 at the respective roots 523. The first vanes 530 and the second vanes 540 may form the crossed configuration. The first vane 530 is swept backward such that the sweep angle a has a negative value. The second vane 540 is swept forward such that the sweep angle i3 has a positive value.

As shown in Figure 6C, in some embodiments, the at least one first vane 530 is swept backward while the at least one second vane 540 is swept forward such that the at least one first vane 530 axially overlaps with the at least one second vane 540 proximate the respective tips 524. An axial overlap region 680 is shown in FIG. 6C. In the configuration 600C, the trailing edge 522 of each first vane 530 is upstream of the leading edge 521 of each second vane 540 at the respective roots 523. An axial offset between the leading edges 521 at the respective roots 523 of the first vanes 530 and the second vanes 540 may be substantially greater than an axial offset between the respective tips 524 of the first vanes 530 and the second vanes 540. In some embodiments, an axial offset at the respective tip 524 of the first vanes 530 and the second vanes 540 may be about half of an axial span of the vanes 520.

Now referring to Figures 6D, 6E and 6F, various embodiments of the vane assembly 500 are illustrated. Two extreme configurations 600D, 600F of the vane assembly 500 are illustrated in Figures 6D and 6F. The sweep of each of the vanes 520 is defined with respect to the baseline position 650 (shown by dashed line). Various other intermediate configurations of the vane assembly 500 may be possible between these two extreme configurations 600D, 600F. An intermediate configuration 600E is shown in Figure 6E. In some embodiments, 5 the at least one first vane 530 is swept backward such that the sweep angle a of the at least first vane 530 is less than 0 degree and greater than -70 degrees. Further, the at least one second vane 540 is swept backward such that the sweep angle p of the at least one second vane 540 is less than 0 degree and greater than -70 degrees. As shown in Figures 6D, 6E and 6F, in some 10 embodiments, the sweep angle a is different from the sweep angle 13.

As shown in Figure 6D, in some embodiments, the at least one first vane 530 axially overlaps with the at least one second vane 540 proximate the respective roots 523. An axial overlap region 661 is shown in FIG. 6D. In the configuration 600D, the axial offset between the leading edges 521 at the respective roots 523 of the first vanes 530 and the second vanes 540 may be substantially less than an axial offset at the respective tips 524 of the first vanes 530 and the second vanes 540. Further, the first vane 530 is swept backward such that the sweep angle a has a negative value. The second vane 540 is swept backward such that the sweep angle 13 has a negative value. In the illustrated embodiment, the sweep angle a is smaller in absolute value than the sweep angle 13. In some embodiments, an axial offset at the respective roots 523 of the first vanes 530 and the second vanes 540 may be about half of an axial span of the vanes 520.

As shown in Figure 6E, in some embodiments, the at least one first vane 530 axially overlaps with the at least one second vane 540 at about 50% of the radial spans R between the respective roots 523 and the respective tips 524. An axial overlap region 671 is shown in FIG. 6E. In the configuration 600E, the trailing edge 522 of each first vane 530 is upstream of the leading edge 521 of each second vane 540 at the respective roots 523. The first vanes 530 and the second vanes 540 may form the crossed configuration. The first vane 530 is swept backward such that the sweep angle a has a negative value. The second vane 540 is swept backward such that the sweep angle 13 has a negative value as well. In the illustrated embodiment, the sweep angle a is greater in absolute value than the sweep angle 13.

As shown in Figure 6F, in some embodiments, the at least one first vane 530 5 and the at least one second vane 540 are swept backward such that the at least one first vane 530 axially overlaps with the at least one second vane 540 proximate the respective tips 524. An axial overlap region 681 is shown in FIG. 6F. In the configuration 600F, the trailing edge 522 of each first vane 530 is upstream of the leading edge 521 of each second vane 540 at the respective 10 roots 523. An axial offset between the leading edges 521 at the respective roots 523 of the first vanes 530 and the second vanes 540 may be substantially greater than an axial offset between the respective tips 524 of the first vanes 530 and the second vanes 540. In the illustrated embodiment, the sweep angle a is greater in absolute value than the sweep angle p. Now referring to Figures 6G, 6H and 61, various embodiments of the vane assembly 500 are illustrated. Two extreme configurations 600G, 6001 of the vane assembly 500 are illustrated in Figures 6G and 61. The sweep of each of the vanes 520 is defined with respect to the baseline position 650 (shown by dashed line). Various other intermediate configurations of the vane assembly 500 may be possible between these two extreme configurations 600G, 6001. An intermediate configuration 600H is shown in Figure 6H. In some embodiments, the at least one first vane 530 is swept forward such that the sweep angle a of the at least first vane 530 is greater than 0 degree and less than 70 degrees.

Further, the at least one second vane 540 is swept forward such that the sweep angle p of the at least one second vane 540 is greater than 0 degree and less than 70 degrees. As shown in Figures 6G, 6H and 61, in some embodiments, the sweep angle a is different from the sweep angle p. As shown in Figure 6G, in some embodiments, the at least one first vane 530 axially overlaps with the at least one second vane 540 proximate the respective roots 523. An axial overlap region 662 is shown in FIG. 6G. In the configuration 600G, the axial offset between the leading edges 521 at the respective roots 523 of the first vanes 530 and the second vanes 540 may be substantially less than an axial offset at the respective tips 524 of the first vanes 530 and the second vanes 540. Further, the first vane 530 is swept forward such that the sweep angle a has a positive value. The second vane 540 is swept forward such that the sweep angle [3 has a positive value. In the illustrated embodiment, the sweep angle a is greater than the sweep angle [3. In some embodiments, an axial offset at the respective roots 523 of the first vanes 530 and the second vanes 540 may be about half of an axial span of the vanes 520.

As shown in Figure 6H, in some embodiments, the at least one first vane 530 axially overlaps with the at least one second vane 540 at about 50% of the radial spans R between the respective roots 523 and the respective tips 524. An axial overlap region 672 is shown in FIG. 6H. In the configuration 600H, the trailing edge 522 of each first vane 530 is upstream of the leading edge 521 of each second vane 540 at the respective roots 523. The first vanes 530 and the second vanes 540 may form the crossed configuration. The first vane 530 is swept forward such that the sweep angle a has a positive value. The second vane 540 is swept forward such that the sweep angle p has a positive value as well. In the illustrated embodiment, the sweep angle a is greater than the sweep angle [3.

As shown in Figure 61, in some embodiments, the at least one first vane 530 and the at least one second vane 540 are swept forward such that the at least one first vane 530 axially overlaps with the at least one second vane 540 proximate the respective tips 524. An axial overlap region 682 is shown in FIG. 61. In the configuration 6001, the trailing edge 522 of each first vane 530 is upstream of the leading edge 521 of each second vane 540 at the respective roots 523. An axial offset between the leading edges 521 at the respective roots 523 of the first vanes 530 and the second vanes 540 may be substantially greater than an axial offset between the respective tips 524 of the first vanes 530 and the second vanes 540. In the illustrated embodiment, the sweep angle a is less than the sweep angle p. In some embodiments, an axial offset at the respective tip 524 of the first vanes 530 and the second vanes 540 may be about half of an axial span of the vanes 520.

Figures 7A, 7B and 7C illustrate plan views of the roots 523 of the first vanes 530 and the second vanes 540 disposed on the hub 510 in accordance with various embodiments of the present disclosure.

Referring to Figure 7A, the roots 523 of the set of first vanes 530 are arranged at an identical first axial position 710A. In other words, the set of first vanes 530 may be disposed circumferentially around the hub 510 with no axial offset between the first vanes 530 at the respective roots 523. Further, the roots 523 of the set of second vanes 540 are arranged at an identical second axial position 710B. In other words, the second vanes 540 may also be disposed circumferentially around the hub 510 with no axial offset between the respective roots 523. The first vanes 530 are spaced apart from each other along the circumferential direction CD. The second vanes 540 are spaced apart from each other along the circumferential direction CD. The roots 523 of first and second vanes 530, 540 are also alternatively arranged along the circumferential direction CD. The root 523 of each first vane 530 is circumferentially disposed between the roots 523 of two adjacent second vanes 540. Similarly, the root 523 of each second vane 540 is circumferentially disposed between the roots 523 of two adjacent first vanes 530.

Now referring to Figure 7B, the roots 523 of the set of first vanes 530 are arranged at an identical first axial position 720A. In other words, the set of first vanes 530 may be disposed circumferentially around the hub 510 with no axial offset between the respective roots 523. Further, adjacent second vanes of the set of second vanes 540 are axially offset from one another at the respective roots 523. Further, alternating second vanes of the set of second vanes 540 are arranged at an identical axial position at the respective roots 523. In other words, the set of second vanes 540 may form two subsets, namely, a first subset 721 and a second subset 722. The first subset 721 may be disposed at an identical second axial position 720B. Further, the second subset 722 may be disposed at an identical third axial position 720C. The first subset 721 and the second subset 722 may include alternating second vanes 540 of the set of second vanes 540. Therefore, the second vanes 540 of the first subset 721 may be arranged at the identical second axial position 720B at the respective roots 523. The second vanes 540 of the second subset 722 may be arranged at the identical third axial position 720C at the respective roots 523.

Referring to Figure 7C, the roots 523 of the set of first vanes 530 are arranged at an identical first axial position 730A. In other words, the first vanes 530 may be disposed circumferentially around the hub 510 with no axial offset at the respective roots 523. However, in some other embodiments, the first vanes 530 may be disposed circumferentially around the hub 510 with an axial offset at the respective roots 523. At least one second vane 540' of the set of second vanes 540 is circumferentially disposed proximate a pylon 700. The at least one second vane 540' is axially positioned upstream of the other second vanes 540 at the respective roots 523. As illustrated, some of the second vanes 540' are circumferentially disposed proximate the pylon 700. These second vanes 540' may be axially positioned upstream of the other second vanes 540 at the respective roots 523 at a distance from the pylon 700.

Furthermore, in some embodiments, the first vanes 530 may be disposed circumferentially around the hub 510 with an axial offset at the respective roots 523. Specifically, at least two first vanes 530 of the set of first vanes 530 are axially offset from one another at the respective roots 523. The second vanes 540 may also be disposed circumferentially around the hub 510 with an axial offset at the respective roots 523.

Providing an axial offset to the first vanes 530 and/or the second vanes 540 may aid in an arrangement of the vanes 520 (shown in Figure 5A) around certain structural components (such as the pylon 700 of Figure 7C or the bifurcations), thereby allowing a higher degree of freedom for mitigating potential effect from the bifurcations and the pylon 700. For example, as illustrated in Figure 7C, in the case of an OGV assembly, a set of OGVs near the pylon 700 attached to the gas turbine engine 10 may be arranged at a different axial position than other OGVs of the set such that the vanes are arranged at a distance from the pylon 700, thereby allowing a higher degree of freedom for mitigating potential effect from the pylon 700.

In some embodiments, the gas turbine engine 10 for an aircraft includes the vane assembly 500 in accordance with the present disclosure. In some embodiments, the aircraft includes the gas turbine engine 10 including the vane assembly 500 in accordance with the present disclosure.

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.

Claims

CLAIMS: 1. A vane assembly (500) for a gas turbine engine (10), the vane assembly (500) comprising: a hub (510) defining a central axis (515); and a plurality of vanes (520) extending from the hub (510), each vane (520) having a leading edge (521) and a trailing edge (522), wherein each vane (520) extends from a root (523) disposed at the hub (510) to a tip (524), wherein the leading edge (521) of each vane (520) defines a corresponding sweep angle (a, 13) relative to a corresponding baseline position (650), the plurality of vanes (520) comprising a set of first vanes (530) circumferentially spaced from one another around the hub (510) and a set of second vanes (540) circumferentially spaced from one another around the hub (510); wherein the leading edge (521) of each first vane (530) is upstream of the leading edge (521) of each second vane (540) at the respective roots (523); and wherein the sweep angle (a) of at least one first vane (530) is different from the sweep angle (13) of at least one second vane (540) such that the at least one first vane (530) at least partially and axially overlaps with the at least one second vane (540) relative to the central axis (515).
2. The vane assembly (500) of claim 1, wherein the at least one first vane (530) is swept backward such that the sweep angle (a) of the at least one first vane (530) is greater than -70 degrees and less than 0 degree
3 The vane assembly (500) of claim 2, wherein the at least one second vane (540) is swept forward such that the sweep angle (13) of the at least one second vane (540) is greater than 0 degree and less than about 70 degrees.
4. The vane assembly (500) of claim 3, wherein the at least one first vane (530) axially overlaps with the at least one second vane (540) at around 50% of radial spans (R) between the respective roots (523) and the respective tips (524).
The vane assembly (500) of claim 3, wherein the at least one first vane (530) axially overlaps with the at least one second vane (540) proximate the respective tips (524).
6. The vane assembly (500) of any one of claims 1 to 5, wherein the trailing edge (522) of each first vane (530) is upstream of the leading edge (521) of each second vane (540) at the respective roots (523).
7. The vane assembly (500) of claim 1, wherein the at least one first vane (530) is swept forward such that the sweep angle (a) of the at least one first vane (530) is greater than 0 degree and less than about 70 degrees.
8. The vane assembly (500) of claim 7, wherein the at least one second vane (540) is swept backward such that the sweep angle (8) of the at least one second vane (540) is greater than -70 degrees and less than 0 degree.
9. The vane assembly (500) of claim 8, wherein the at least one first vane (530) axially overlaps with the at least one second vane (540) proximate the respective roots (523).
10. The vane assembly (500) of any one of claims 1 to 9, wherein at least two first vanes (530) of the set of first vanes (530) have different sweep angles (a).
11 The vane assembly (500) of any one of claims 1 to 10, wherein at least two second vanes (540) of the set of second vanes (540) have different sweep angles (6).
12. The vane assembly (500) of any one of claims 1 to 11, wherein at least two first vanes (530) of the set of first vanes (530) are axially offset from one another at the respective roots (523).
13.
14.
15.
16.
17. 18. 19. 20.The vane assembly (500) of any one of claims 1 to 12, wherein at least two second vanes (540) of the set of second vanes (540) are axially offset from one another at the respective roots (523).The vane assembly (500) of any one of claims 1 to 12, wherein at least one second vane (540') of the set of second vanes (540) is circumferentially disposed proximate a pylon (700), and wherein the at least one second vane (540') is axially positioned upstream of the other second vanes (540) at the respective roots (523).The vane assembly (500) of any one of claims 1 to 12, wherein adjacent second vanes (540) of the set of second vanes (540) are axially offset from one another at the respective roots (523), and wherein alternating second vanes (540) of the set of second vanes (540) are arranged at an identical axial position (720B, 720C) at the respective roots (523).The vane assembly (500) of any one of claims 1 to 11, wherein the roots (523) of the set of first vanes (530) are arranged at an identical first axial position (710A, 720A, 730A).The vane assembly (500) of any one of claims 1 to 12 and 16, wherein the roots (523) of the set of second vanes (540) are arranged at an identical second axial position (710B).The vane assembly (500) of any one of claims 1 to 17, wherein each of the plurality of vanes (520) is a fan outlet guide vane.A gas turbine engine (10) for an aircraft, the gas turbine engine (10) comprising the vane assembly (500) of any one of claims 1 to
18.An aircraft comprising the gas turbine (10) engine of claim 19.