GB2603148A - Gas turbine engine - Google Patents

Abstract

A gas turbine engine for an aircraft has an engine core 11 comprising a compressor 14 driven via a core shaft by a turbine 19, a bypass duct 22 surrounding the engine core 11 and a splitter 12 to split the airflow received by the engine into a core airflow A and a bypass airflow B. A fan 23 is disposed axially between the compressor 14 and the turbine 19. The fan 23 is driven by the turbine 19 to move the bypass airflow through the bypass duct 22. Each fan blade 41 may have an inner portion 42 extending across the core airflow path, an outer portion 43 extending across the bypass duct 22, and an intermediate portion 45 separating the core airflow A from the bypass airflow B. The inner and outer blade portions 42, 43 may be pivotable about a longitudinal axis of the blade 41. Inlet guide vanes 46 and outlet guide vanes 50 may be provided upstream and downstream, respectively, of the fan 41.

Description

GAS TURBINE ENGINE

Field of the disclosure

The present disclosure relates to a gas turbine engine for an aircraft. Background Modern gas turbine engines typically comprise a fan positioned upstream of an engine core, having a circumferential row of fan blades mounted onto a fan hub. The fan is driven by a turbine in the engine core to generate an air flow. A radially outer portion of the generated air flow enters a bypass duct to provide propulsive thrust. The remainder of the air flow, towards the fan hub, flows through to the engine core to power a Brayton cycle, and to provide additional propulsive thrust. In typical turbofan engines the fan is directly driven by a shaft that extends from a low-pressure turbine of the engine to the fan, and as such the fan rotates at the same rotational speed as the low-pressure turbine.

In known engines, the bypass duct is defined between the engine core and a nacelle that surrounds the engine core (but that is spaced therefrom). The nacelle circumscribes the fan and projects forward of the fan such that the fan is spaced from an inlet of the turbine engine defined by a leading edge of the nacelle. The distance between the leading edge and the fan helps to remove turbulence from the air prior to the air entering the fan.

This portion of the nacelle, however, can be particular susceptible to movement, because it is commonly cantilevered from a support structure that is located rearward of the fan. This can result in relative movement between the nacelle and the fan, which can in turn cause increased wear of the fan.

This extension of the nacelle beyond the fan also increases the length of the engine. Such engines are usually mounted so as to be somewhat forward of the wing. Thus, the longer the engine is, the greater the moment applied to the wing by the engine, which in turn results in more demanding structural requirements with respect to the design of the wing.

A further issue with known arrangements is that, in some cases, when the thrust of the engine is reversed by a thrust reverser of the engine (e.g. during landing), air may flow to the engine core via the bypass duct from a rear end of the bypass duct. This may occur, for example, when the thrust reversing is such that airflow is significantly restricted through the fan of the engine. Such an airflow path results in inefficient operation of the engine core.

Thus, it may be desirable to provide a gas turbine engine with an improved structure and/or a lower susceptibility to inefficient airflow paths being formed during reverse thrust operations.

Summary

According to a first aspect there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a compressor, a turbine fluidly connected to, and downstream of, the compressor, and a core shaft connecting the turbine to the compressor; a bypass duct circumferentially surrounding the engine core; a splitter arranged to split an airflow received by the engine into a core airflow that flows along an engine core airflow path through the compressor and the turbine, and a bypass airflow that flows through the bypass duct; and a fan disposed axially between the compressor and the turbine, the fan driven by the turbine to move the bypass airflow through the bypass duct.

Positioning the fan between the compressor and the turbine (i.e. so as to be downstream of the compressor) may result in the turbine engine having a more compact design. The position of a leading edge of the housing (or nacelle) of a turbine engine is somewhat dependent on the location of the fan in the engine. This is because the nacelle scoops air into the engine over a wide range of incoming angles and a sufficient distance must be provided between the fan and the leading edge of the nacelle to ensure that turbulence is reduced by the time the air reaches the fan from the leading edge. Thus, positioning the fan downstream of the compressor (i.e. so as be in the bypass duct) may allow for a more compact (i.e. shorter) engine.

The provision of a more compact engine may result in the centre of gravity of the engine being closer to the wing of an aircraft when the engine is mounted thereto. This may result in a reduced moment applied to the wing of the aircraft by the turbine engine. In this way, the structural requirements of the turbine engine, the wing, and any mounting structure mounting the engine to the wing, may be eased.

The terms "downstream" and "upstream" used herein are used with reference to the direction of airflow through the turbine engine in normal use (e.g. during forward movement of an aircraft to which the turbine engine may be mounted). Thus, the upstream end is a forward end of the turbine engine (i.e. closer to the front of an aircraft when mounted thereto) and the downstream end is a rearward end of the turbine engine. The term "axial" as used herein is a reference to a direction that is generally parallel to a rotational axis of the turbine engine (e.g. the rotational axis of the core shaft of the turbine engine).

Optional features of the present disclosure will now be set out. These are applicable singly or in any combination with any aspect of the present disclosure.

The fan may extend radially across the engine core airflow path, and at least partly across the bypass duct. The fan may comprise plurality of fan blades. The fan blades may extend radially at least partway across (e.g. fully across) the bypass duct.

The fan blades of the fan may extend across the engine core airflow path. In this respect, each fan blade may comprise an inner blade portion that extends across the engine core airflow path and an outer blade portion that extends at least partway across the bypass duct.

The fan may comprise a fan hub (e.g. disk) from which the fan blades extend radially. The fan hub may form part of a central structure of the engine core (i.e. at least partly defining a radially inner surface of the core airflow path).

The inner blade portion of each fan blade may represent less than 25% of the total length of the fan blade (i.e. in the radial direction). The inner blade portion of each fan blade may represent less than e.g. 20% of the total length of the fan blade.

At least a portion of each fan blade may be adjustable so as to alter the airflow passing across the fan blade. The adjustable portion may be rotatable about a longitudinal axis of the fan blade. The adjustable portion may be in the form of a trailing flap. The inner blade portion and/or the outer blade portion of each fan blade may be adjustable. Alternatively, only a portion of the inner and/or outer blade portion may be adjustable (e.g. for example a trailing flap of the inner blade portion).

Adjustment of the outer blade portions of the fan blades may allow adjustment of the airflow through the bypass duct. In some cases, such adjustment may, for example, be used in lieu of thrust reversers. In such embodiments, the turbine engine may not comprise any thrust reversers (other than the adjustable fan blades), which may allow further reduction of the size of the housing/nacelle. In particular, this may allow a thinner (i.e. radial dimension) nacelle, such that the nacelle provides less air resistance to oncoming air in use.

The inner blade portion and outer blade portion of each fan blade may have different chord lengths. The outer blade portion of each fan blade may have a greater chord length than the corresponding inner blade portion of the fan blade. Thus, the outer blade portions may be wider than the inner blade portions. Each fan blade may extend substantially fully across the core airflow path and bypass duct. It is noted that, in general, it may be desirable to have a small clearance between the tips (i.e. distal ends) of the fan blades and the housing to ensure rotation of the fan blades is not impeded by the housing.

An intermediate blade portion may be interposed between the inner and outer portions of each fan blade. The intermediate portion may be configured to separate the core airflow from the bypass airflow (i.e. so as to seal between the bypass duct and the core airflow path). For example, the intermediate portion may be an enlarged portion of the blade. The intermediate portion may extend circumferentially so as to connect with the intermediate portions of adjacent blades. In this way, the intermediate portions may connect blades so as to form a circumferentially extending ring that provides a seal between the bypass duct and the core airflow path. In this way, the intermediate portions may be referred to as "snubbers" and the fan blades may be referred to as "snubbered" blades.

In other embodiments, the fan blades may only extend across the bypass duct (i.e. so as not to have inner blade portions). In such embodiments, the fan may comprise a blade support structure from which the fan blades extend radially. In this respect, the blade support structure may be interposed (radially) between the fan hub and the fan blades.

The blade support structure may extend across the engine core airflow path. The blade support structure may comprise a plurality of radially extending (circumferentially spaced) vanes or support members that extend radially across the engine core airflow path (i.e. so as to only partially obstruct the engine core airflow path). In such embodiments, the blade support structure may substantially seal between the bypass duct and the engine core airflow path.

The bypass duct may be defined between a housing (e.g. nacelle) of the turbine engine and the engine core. The housing may extend circumferentially about, and be radially spaced from, the engine core (so as to define the bypass duct). The bypass duct may be annular. In this respect, the bypass duct and core airflow path may be concentrically arranged. The core airflow path may have a smaller cross-sectional area than the bypass duct.

In some embodiments a leading edge of the splitter may be substantially aligned with (i.e. aligned along the radial axis) a leading edge of the housing. In other embodiments, the leading edge of the splitter may be forward of a leading edge of the housing. Alternatively, the leading edge of the splitter may be rearward of the leading edge of the housing. The splitter may define an inlet to the engine core. The splitter may be an edge or lip of the engine core defining an outer edge of the engine core inlet.

The splitter may form a portion of a casing structure (e.g. a nacelle) of the engine core. The casing structure may define a radially outer boundary of the core airflow path and a radially inner boundary of the bypass duct. In this respect, the casing structure may maintain separation of the core airflow and the bypass airflow. The engine core inlet may be substantially annular, or in some cases, may be circular.

The turbine engine may comprise a plurality of guide vanes extending radially from the engine core to the housing (i.e. across the bypass duct). The guide vanes may support the engine core within the housing (i.e. so as to maintain the position of the engine core with respect to the housing). The guide vanes may be upstream of the fan and, for example, the axial distance from the guide vanes to the fan may be less than a radius of the fan (or less than a fan blade length). The guide vanes may be referred to as inlet guide vanes of the turbine engine. The guide vanes may be axially rearward of the compressor of the turbine engine. The guide vanes may each have an aerofoil profile.

At least a portion of each inlet guide vane may be adjustable so as to alter the airflow passing across the guide vane. For example, each inlet guide vane may be rotatable about a longitudinal (i.e. radially extending) axis of the guide vane. Alternatively, only a portion of the guide vane may be adjustable. For example, each inlet guide vane may comprise a flap that is rotatable (e.g. piyotable) about a longitudinal axis of the inlet guide vane. The flap may be at (and may define) a trailing edge of the guide vane.

Such adjustment of the inlet guide vanes may allow control of the bypass and core airflows passing from the inlet guide vanes to the fan. For example, the inlet guide vanes may be configured to introduce swirl into the airflow flowing across the inlet guide vanes. This may reduce the incidence and Mach number of the airflow that is received by the fan, which may help to reduce the noise emitted by the engine.

The provision of inlet guide vanes could, in some cases, result in increased noise due to interaction between the wake downstream of the inlet guide vanes and the fan. This may, however, be somewhat mitigated by way of selection of the number of inlet guide vanes, and the distance of those guide vanes upstream of the fan. In addition, the acceleration of the flow by the inlet guide vanes (resulting in turning of the flow) may homogenise the flow (i.e. reducing circumferential variation in the flow) that enters the fan, which may help to provide a reduction in noise caused by distortion in the flow.

Each inlet guide vane may extend radially across the bypass duct. Each inlet guide vane may additionally extend radially across the core airflow path. The portion of the guide vane in the core airflow path may be downstream of the compressor. Thus, the guide vanes may be rearward/downstream of the splitter, such that the core and bypass airflows are physically separated (by the splitter) at the point at which they are received by the guide vanes. A portion of each guide vane may pass radially through the engine core casing structure.

Each inlet guide vane may thus comprise an inner portion (in the core airflow path) and an outer portion On the bypass duct). The inner and outer portions of each inlet guide vane may be independently variable. In some embodiments the outer portion of each inlet guide vane may be variable, whilst the inner portion of each inlet guide vane may be static (or vice-versa).

In other embodiments, the turbine engine may comprise a plurality of outer inlet guide vanes extending radially across the bypass duct (between the engine core and the housing), and a plurality of separate (i.e. to the outer inlet guide vanes) inner guide vanes extending radially across the core airflow path (within the engine core). The inner and outer guide vanes may each be mounted to the engine core casing structure, so as to be connected via the casing structure. Similar to that described above, the inner and outer inlet guide vanes may be independently variable. In some embodiments, the outer support structures may be variable whilst the inner inlet guide vanes are static (or vice-versa).

The turbine engine may comprise a plurality of radially extending guide vanes located downstream of the fan. Such guide vanes may be referred to as outlet guide vanes. The axial distance between the outlet guide vanes may be less than a radius of the fan (or less than a fan blade length of the fan). The outlet guide vanes (like the inlet guide vanes) may extend radially between the engine core and the housing, and may support the engine core within the housing. The inlet guide vanes may extend from (an outer surface of) the casing structure of the engine core across the bypass duct (e.g. only across the bypass duct and not across the core airflow path).

As should be apparent, when both inlet and outlet guide vanes are provided, the fan is interposed (axially) between the two sets of the guide vanes. In such embodiments, the inlet and outlet guide vanes provide a "box" structure that increases the rigidity of the engine. Such rigidity may reduce relative movement between the engine core, the fan, and the housing (e.g. nacelle). Such movement may be undesirable. For example, relative movement between the fan and the housing may cause the tips of the fan blades to contact the housing (so as to cause wear of the fan blades).

Additionally, such an arrangement (i.e. having inlet and outlet guide vanes) may provide improved containment of the fan blades should one or more of the fan blades fail.

In some embodiments, the turbine engine may comprise a mounting structure for mounting the turbine engine to an aircraft (e.g. the wing of an aircraft). The mounting structure may be referred to as a pylon of the turbine engine. The mounting structure may be connected to (e.g. coupled to or integral with) an outlet guide vane of the plurality of outlet guide vanes of the turbine engine. The mounting structure (pylon) may extend rearwardly from the outlet guide vane to which it coupled. The rearward end of the mounting structure may be configured for mounting to an aircraft wing (e.g. the underside of an aircraft wing).

The engine core may further comprise a particle separator/filter upstream of the compressor (e.g. proximal or at the engine core inlet). The filter may prevent unwanted particles from entering the engine core and causing wear/damage of the components of the engine core.

As is provided above, the turbine and compressor are connected (e.g. directly connected) by a core shaft, such that the turbine and compressor co-rotate. The engine core airflow path may extend in a generally axial direction from the compressor to the turbine. The engine core may further comprise a combustor downstream of the compressor and upstream of the turbine. The combustor may be disposed axially between the compressor and the turbine.

Thus, the combustor and turbine may be radially inward of the bypass dud.

Each of bypass duct and engine core airflow path may comprise a respective outlet (for discharge of their respective airflows). The outlet of the bypass duct may be radially outward of the outlet of the engine core airflow path.

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 or 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 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 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 platform.

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), 350cm, 360cm (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-limitafive 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 non-limitative 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 Ufip.

The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/U1ip2, where dH is the enthalpy rise (for example the 1D 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 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 units in this paragraph being Jkg-1K-1/(ms-1)2). 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 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 form 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-limitafive 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-is, 105 Nkg-is, 100 Nkg-is, 95 Nkg-is, 90 Nkg-ls, 85 Nkg-ls or 80 Nkg-is. 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-ls, 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 TET 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 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 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, mid-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 mid-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 mid-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; and Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine. 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 an exemplary gas turbine engine 10 having a principal rotational axis 9. 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 On 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.

The engine 10 comprises a splitter 12 that splits an airflow received by the engine 10 into two airflows: a core airflow A and a bypass airflow B. In particular, the splitter 12 is at a forward (or leading) end of the turbine engine 10 such that the airflow is split prior to interacting with any of the components within the core airflow A, such as a low-pressure compressor 14 or optional particle filter 52, and prior to interacting with any of the components within the bypass airflow B, such as a fan 23 or optional inlet guide vanes 46. In this example, once the air coming into the turbine engine 10 has been split into the core airflow A and bypass airflow B by the splitter 12, the two airflows remain independent of each other until they are ejected from the turbine engine 10 at a core exhaust nozzle 20 and a bypass exhaust nozzle 18 respectively.

The gas turbine engine 10 comprises an engine core 11 (from which the splitter 12 projects forward) that defines a core airflow path 44, which receives the core airflow A. In particular, the engine core 11 comprises an annular engine core inlet 53 that receives the core airflow A. The outer circumference of the annular engine core inlet 53 is defined by the splitter 12.

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 housing in the form of a nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The engine core 11 may optionally include a particle filter 52 upstream of the low-pressure compressor 14 to prevent particles from travelling further down into the engine core. The bypass airflow B flows through the bypass duct 22. A fan 23, which is downstream of the splitter 12 and the low-pressure (or intermediate-pressure) compressor 14, is attached to and driven by the low-pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30 (as will be described further below with reference to Figures 2 and 3).

The fan 23 comprises a plurality of fan blades 41 that each extend from the engine core 11 to the nacelle 21. In particular, each fan blade 41 comprises an inner blade portion 42 and an outer blade portion 43 that is radially outward of the inner blade portion 43. In other embodiments, the fan 23 may instead comprise a blade support structure that supports the fan blades 41.

The inner blade portions 42 extend within (and across) the core airflow path 44 so as to move the core airflow A through the engine core 11 (when rotated by the low-pressure turbine 19). The outer blade portions 43, on the other hand, extend radially across the bypass duct 22, such that when the fan 23 is rotated, the outer blade portions 43 move the bypass airflow B through the bypass duct 22. Whilst not apparent from the figures, each inner blade portion 42 and outer blade portion 43 is independently pivotable about their respective longitudinal axes (i.e. in the radial direction). In this way, the bypass B and core A airflows may be controlled by adjustment of the inner blade portions 42 and outer blade portions 43.

Each blade 41 of the fan 23 also comprises an intermediate portion in the form of a snubber interposed between the inner 42 and outer 43 blade portions of the fan blade 41. Whilst not apparent from the figures, each snubber 45 extends circumferentially so as to join with adjacent snubbers 45 so as to maintain the separation between the core A and bypass B airflows even as the fan 23 rotates.

The turbine engine 10 further comprises a plurality of radially extending inlet guide vanes 46 extending from the engine core 11 to the nacelle 21. Like the fan blades 41, each inlet guide vane 46 comprises an inner portion 47 and an outer portion 48. Each inner portion 47 is located in (and extends across) the core airflow path 44, and each outer portion 48 is located in (and extends across) the bypass duct 22. The inlet guide vanes 46 are downstream of the splitter 12 and the low-pressure compressor 14, but upstream of the fan 23. In this way, the inlet guide vanes 46 may redirect the airflows A, B so as to be in a desired condition prior to receipt by the fan 23. In order to facilitate this, the outer portions 48 each comprise a flap 49 at a trailing (downstream) edge. Each flap 49 is pivotable about a longitudinal axis of the respective fan blade 41 such that the airflow passing through the inlet guide vanes 46 can be adjusted.

The turbine engine 10 also includes outlet guide vanes 50 that are downstream of the fan 23 and that extend radially between the engine core 11 and the nacelle 21. Unlike the inlet guide vanes 46 and the fan blades 41, the outlet guide vanes 50 extend only within the bypass duct 22 and are static (i.e. they are not movable). A mounting structure in the form of a pylon 51 (shown in dashed lines for clarity) is coupled to, and extends rearwardly from, the uppermost outlet guide vane 50. This pylon 51 is configured to mount the turbine engine 10 to the wing of an aircraft. The integration of the pylon 51 with the outlet guide vane 50 reduces the obstruction caused by the pylon 51 to the airflow in the bypass duct 22.

As is discussed previously, the provision of inlet guide vanes 46 and outlet guide vanes 50 upstream and downstream of the fan 23 provides a box structure that increases the rigidity of the turbine engine 10. This structure, for example, reduces relative movement between the nacelle 21 and the engine core 11.

In use, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15 (via the inner portions 47 of the inlet guide vanes 46 and the inner blade portions 42 of the fan 23) 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 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

The arrangement of Figure 1 is shown in more detail 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 precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to the inlet guide vanes 46.

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. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

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 (14)

1. CLAIMSA gas turbine engine (10) for an aircraft, the gas turbine engine (10) comprising: an engine core (11) comprising a compressor (14), a turbine (19) fluidly connected to, and downstream of, the compressor (14), and a core shaft connecting the turbine to the compressor (14); a bypass duct (22) circumferentially surrounding the engine core (11); a splitter (12) arranged to split an airflow received by the engine (10) into a core airflow (A) that flows along an engine core airflow path (44) through the compressor and the turbine (19), and a bypass airflow (B) that flows through the bypass duct (22); and a fan (23) disposed axially between the compressor (14) and the turbine (19), the fan driven by the turbine (19) to move the bypass airflow through the bypass duct (22).
2. 2. The gas turbine engine (10) of claim 1 wherein the fan comprises a plurality of radially extending fan blades, each fan blade comprising an inner blade portion (42) extending across the engine core airflow path (44) and an outer blade portion (41) extending at least partway across the bypass duct (22).
3. The gas turbine engine (10) of claim 2 wherein each fan blade (41) comprises an intermediate blade portion (45) interposed between the inner (42) and outer (43) blade portions, each intermediate portion (45) configured to separate the core airflow (A) from the bypass airflow (B)
4. 4. The gas turbine engine (10) according to claim 2 or 3 wherein the inner blade portion (42) of each fan blade (41) is pivotable about a longitudinal axis of the fan blade (41).
5. 5. The gas turbine engine (10) according to any one of claims 2 to 4 wherein the outer blade portion (43) of each fan blade (41) is pivotable about a longitudinal axis of the fan blade (41)
6. 6. The gas turbine engine (10) according to any one of the preceding claims further comprising a plurality of inlet guide vanes (46) upstream of the fan (23) and downstream of the splitter (12), the plurality of guide vanes (46) extending radially across the bypass duct (22).
7. 7. The gas turbine engine (10) of claim 6 wherein at least a portion of each inlet guide vane (46) is pivotable about a longitudinal axis of the inlet guide vane (46).
8. 8. The gas turbine engine (10) according to claim 6 or 7 wherein each inlet guide vane (46) further comprises an inner portion (47) extending radially across the core airflow path (44) and an outer portion (48) extending radially across the bypass duct (22).
9. 9. The gas turbine engine (10) of claim 8 wherein the outer portion (48) of each inlet guide vane (46) is rotatable about a longitudinal axis of the inlet guide vane (46) to adjust airflow across the guide vane (46).
10. The gas turbine engine (10) of claim 8 wherein the outer portion (48) of each inlet guide vane (46) further comprises a flap (49) at a trailing edge thereof, the flap (49) being pivotable about a longitudinal axis of the guide vane (46) to adjust airflow across the guide vane (46).
11. 11. The gas turbine engine (10) according to any one of claims 8 to 10 wherein the inner portion (47) of each inlet guide vane (46) is downstream of the compressor (14).
12. 12 The gas turbine engine (10) according to any one of the preceding claims further comprising a plurality of outlet guide vanes (50) downstream of the fan (23) and extending radially across the bypass duct (22).
13. 13 The gas turbine engine (10) of claim 12 further comprising a mounting structure (51) for mounting the engine (10) to the wing of an aircraft, the mounting structure (51) coupled to an outlet guide vane (50) of the plurality of outlet guide vanes (50).
14. 14 The gas turbine engine (10) according to any one of the preceding claims further comprising a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan (23) at a lower rotational speed than the core shaft (26) The gas turbine engine (10) according to any one of the preceding claims further comprising a combustor (16) disposed axially between, and fluidly connected to, the compressor (14) and the turbine (19).16 The gas turbine engine (10) of claim 15 wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the core shaft is a first core shaft (26); the engine core (11) further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) connecting the second turbine to the second compressor; and wherein the second turbine (17), second compressor (15), and second core shaft (27) are arranged to rotate at a higher rotational speed than the first core shaft (26).