GB2588956A - A variable vane assembly - Google Patents

Abstract

A variable vane assembly 42 comprises a spindle 52 coupled to a lever 66, the spindle at least partially defining a slot 88 within which a seal 90 is positioned. The lever at least partially covers the slot to prevent removal of the seal from the slot. The lever also defines a sealing surface 94 against which the seal abuts. The spindle and lever may be distinct and seperable components. The slot may be circular and the seal annular. The spindle may be received within a bush 64 which also at least partially defines the slot. The bush may comprise axially and radially extending portions 86, 84, which may each provide a further sealing surface. The lever or a casing part may comprise an axially extending flange which defines a further sealing surface. The spindle may comprise a recess adjacent to the lever, the recess at least partially defining the slot. The variable vane assembly may be in a high-pressure compressor of a gas turbine engine such as a turbofan engine.

Description

A VARIABLE VANE ASSEMBLY

Field of the Disclosure

The present disclosure relates to a variable vane assembly.

Background of the Disclosure

Variable vane assemblies such as variable inlet guide vanes (VGVs) and variable stator vanes (VSV) comprise aerofoils that are located in a core airflow and rotatably mounted to a case. During operation, high temperature gas can leak from the core airflow via the joint between the aerofoil and the case. This leakage increases the temperature of components adjacent the joint, which can detrimentally affect their performance. In particular, the variable vane comprises a number of components made from low-friction materials that slide or rotate relative to each other, and which are prone to wear at an accelerated rate or deteriorate to complete loss of function when exposed to such elevated temperatures. Leakage of this type has also been shown to increase specific fuel consumption (SFC).

Existing gas turbine engines prevent leakage in part by allowing an outer penny of the variable vane to seal against a bush located between the penny and the case. A sealing effect is provided as a result of the high pressure within the core airflow exerting a radially outward force on the variable vane. However, it has been found that the pressure within the core airflow is often insufficient to create the seal, resulting in the leakage of gas. An alternative approach is to locate an 0-ring in a groove in a spindle of the variable vane, which seals against a bush. However, such a design requires the 0-ring to be stretched around the spindle during assembly, and materials that are both sufficiently flexible to allow this whilst also being able to operate under high temperatures are not readily available.

It is therefore desirable to provide an improved variable vane assembly that overcomes these issues.

Summary of the Disclosure

According to a first aspect there is provided a variable vane assembly comprising a variable vane assembly, the variable vane comprising a spindle coupled to a lever, the spindle at least partially defining a slot within which a seal is positioned, the lever at least partially occluding the slot so as to prevent removal of the seal from the slot, the lever defining a sealing surface against which the seal abuts.

The spindle and the lever may be distinct components.

The spindle and the lever may be separable from each other.

The slot may be circular. The seal may be annular.

The variable vane assembly may be a high pressure compressor variable vane assembly.

The variable vane assembly may further comprise a bush within which the spindle is received for rotation about a rotational axis. The bush may at least partially define the slot.

The variable vane assembly may further comprise a case having an opening for receiving the bush. The bush may comprise a radially extending portion extending in a radial direction away from the rotational axis.

The bush may define a further sealing surface against which the seal abuts.

The bush may comprise an axially extending portion extending in an axial direction away from the radially extending portion. The axially extending portion may define the further sealing surface.

The further sealing surface may be disposed radially between the spindle and an outer surface of the lever.

The outer surface of the lever may be disposed radially between the spindle and the further sealing surface.

The radially extending portion may define the further sealing surface.

The lever may comprise an axially extending flange. The flange may at least partially occlude the slot so as to prevent removal of the seal from the slot. The flange may define the sealing surface against which the seal abuts.

The case may comprise an axially extending flange. The flange may define a further sealing surface against which the seal abuts.

The spindle may comprise a recess adjacent the lever. The recess may at least partially define the slot.

The recess may have a rectangular profile.

A gas turbine engine may further comprise an engine core, a fan, a gearbox and a variable vane assembly as described in any preceding statement. The engine core may comprise a turbine, a compressor and a core shaft connecting the turbine to the compressor. The fan may be located upstream of the engine core. The fan may comprise a plurality of fan blades. The gearbox may receive an input from the core shaft and output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The turbine may be a first turbine. The compressor may be a first compressor. 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.

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 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-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 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 Utip. 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/Utip2, 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 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-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-'s, 105 Nkg-'s, 100 Nkg-'s, 95 Nkg-'s, 90 Nkals, 85 Nkg-ls or 80 Nkg-'s. 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-is to 100 Nkals, or 85 Nkals to 95 Nkals. 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 15 engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; Figure 4 is a cross-sectional side view of a variable vane assembly; Figure 5 is a close-up cross-sectional view of the variable vane assembly; Figure 6 is a close-up cross-sectional view of a first alternative variable vane assembly; Figure 7 is a close-up cross-sectional view of a second alternative variable vane assembly; Figure 8 is a close-up cross-sectional view of a third alternative variable vane 30 assembly; Figure 9 is a close-up cross-sectional view of a fourth alternative variable vane assembly; and Figure 10 is a close-up cross-sectional view of a fifth alternative variable 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 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 compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the core exhaust nozzle 20 to provide some propulsive thrust.

The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 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 precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

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.

Figure 4 is a cross-sectional side view of a high pressure compressor variable vane assembly 42. The high pressure compressor variable vane assembly 42 comprises a vane 44 having an aerofoil portion 50 and a spindle 52. A case 54 of the high pressure compressor 15 separates the core 56 of the high pressure compressor 15 from an exterior 58 of the core flow of the high pressure compressor 15. The case 54 is provided with an opening 80 that extends between the core 56 and the exterior 58. A turret or cylindrical flange 62 of the case 54 extends away from the core 56 and partly defines the opening 80. A bush 64 extends part way into the opening 80. The aerofoil portion 50 is disposed in the core 56 of the high pressure compressor 15 and the spindle 52 extends through the opening 80. A first end of the lever 66 is fixedly attached to the end of the spindle 52. The lever 66 and the spindle 52 are distinct components that are connected during assembly, and may be separable from each other. A second end of the lever 66 is attached to a pin 72. The pin 72 is attached to a unison ring 74. The unison ring 74 is connected to an actuator (not shown) that is configured to rotate the unison ring 74 about the principal rotational axis 9. This in turn cause the lever 66 and the vane 44 to rotate about a rotational axis 76 that is perpendicular to the principal rotational axis 9.

Figure 5 shows a close-up view of the interface between the spindle 52, the case 54, the bush 64 and the lever 66. The area of the close-up view is denoted using reference numeral 78 in Figure 4. It will be appreciated that the geometry shown in the following figures extends around the rotational axis 76.

The bush 64 comprises a first axially extending portion 82, a radially extending portion 84 and a second axially extending portion 86 arranged in series. The first axially extending portion 82 is tubular, is aligned with the rotational axis 76 and extends along the rotational axis 76. The first axially extending portion 82 extends part way into the opening 80. The radially extending portion 84 is planar and in the form of a disc. The radially extending portion 84 extends in a radial direction away from the first axially extending portion 82 at an outer end thereof and away from the rotational axis 76. The second axially extending portion 86 is tubular, is aligned with the rotational axis 76 and extends along the rotational axis 76. The diameter of the second axially extending portion 86 is greater than the diameter of the first axially extending portion 82.

The spindle 52 and the bush 64 define a slot 88. The slot 88 is circular and extends around the rotational axis 76. A seal 90 is positioned within the slot 88. The seal 90 is annular and has a rectangular cross-sectional profile. When the lever 66 is not attached to the spindle 52, the slot 88 is open at its outer end (i.e. the upper end shown in Figure 5). When the lever 66 is attached to the spindle 52, the lever 66 partially occludes the slot 88. The lever 66 is spaced from the bush 64 in an axial direction (i.e. along the rotational axis 76). A gap 92 is thus formed between the second axially extending portion 86 and the lever 66. The gap 92 extends in an axial direction (i.e. it is formed of two faces that are spaced from each other along the rotational axis 76). The width of the gap 92 is smaller than the width of the seal 90. Accordingly, the seal 90 is unable to pass through the gap 92 and the lever 66 prevents removal of the seal from the slot 88.

During operation, the core 56 of the high pressure compressor 16 has a higher pressure than the exterior 58. Air enters the slot 88 from the core 56 and the pressure differential between the air in the slot 88 and the air in the exterior 58 forces the seal 90 against the lever 66 and the second axially extending portion 86. The lever 66 thus defines a first sealing surface 94 against which the seal 90 abuts and the second axially extending portion 86 thus defines a second sealing surface 96 against which the seal 90 abuts. The second sealing surface 96 is disposed radially (with reference to the rotational axis 76) between the spindle 52 and an outer surface of the lever 66 (i.e. the radially outer surface of the hub of the lever 66). That is, the distance between the second sealing surface 96 and the rotational axis 76 is less than the distance between the outer surface of the lever 66 and the rotational axis 76. A seal is created between the seal 90 and both the lever 66 and the second axially extending portion 86. Air from the core 56 is therefore prevented from venting from the core 56 to the exterior 58.

Figures 6, 7, 8, 9 and 10 show close-up views of interfaces of a first alternative high pressure compressor variable vane assembly 142, a second alternative high pressure compressor variable vane assembly 242, a third alternative high pressure compressor variable vane assembly 342, a fourth alternative high pressure compressor variable vane assembly 442 and a fifth alternative high pressure compressor variable vane assembly 542, respectfully. The areas shown in the close-up views of Figures 6 to 10 correspond to that shown in the close-up view of Figure 5. Features not shown in the close-up view of Figures 6 to 10 correspond to those shown in Figure 4. Features that correspond to those already described with reference to Figures 4 and 5 are denoted using the same reference numerals, are structurally the same as those already described with reference to Figures 4 and 5 and function in the same manner as those already described with reference to Figures 4 and 5.

As shown in Figure 6, the outer surface of the lever 66 of the first alternative high pressure compressor variable vane assembly 142 is disposed radially between the spindle 52 and the second sealing surface 196. That is, the distance between the outer surface of the lever 66 and the rotational axis 76 is less than the distance between the second sealing surface 196 and the rotational axis 76. The gap 192 thus extends in both an axial direction and a radial direction (i.e. it is formed of two edges that are spaced from each other in both an axial direction and a radial direction).

As shown in Figure 7, the outer surface of the lever 66 of the second alternative high pressure compressor variable vane assembly 242 is disposed radially between the spindle 52 and the second sealing surface 296. That is, the distance between the outer surface of the lever 66 and the rotational axis 76 is less than the distance between the second sealing surface 296 and the rotational axis 76. The bush 264 overlaps the lever 66 in an axial direction. The gap 292 thus extends in a radial direction (i.e. it is formed of two faces that are spaced from each other in a radial direction with respect to the rotational axis 76).

As shown in Figure 8, the bush 364 of the third alternative high pressure compressor variable vane assembly 342 does not comprise a second axially extending portion 86. The lever 366 of the third alternative high pressure compressor variable vane assembly 342 comprises an annular flange 900 that extends towards the bush 364 and the case 54. The bush 364 is spaced from the annular flange 900 in an axial direction. A gap 392 is thus formed between the radially extending portion 384 of the bush 364 and the annular flange 900. The gap 392 extends in an axial direction (i.e. it is formed of two faces that are spaced from each other in an axial direction). An inner edge of the annular flange 900 defines the first sealing surface 394 against which the seal 90 abuts and the radially extending portion 384 defines the second sealing surface 396 against which the seal 90 abuts.

As shown in Figure 9, the bush 464 of the fourth alternative high pressure compressor variable vane assembly 442 does not comprise a second axially extending portion 86. The case 454 of the fourth alternative high pressure compressor variable vane assembly 442 comprises an annular flange 902. The lever 66 defines a first sealing surface 494 against which the seal 90 abuts and the inner edge of the annular flange 900 defines the second sealing surface 496 against which the seal 90 abuts. The outer surface of the lever 66 of the fourth alternative high pressure compressor variable vane assembly 442 is disposed radially between the spindle 52 and the second sealing surface 496. The annular flange 902 extends towards and overlaps the lever 66 in an axial direction. The gap 492 thus extends in a radial direction.

As shown in Figure 10, the bush 564 of the fifth alternative high pressure compressor variable vane assembly 542 does not comprise a second axially extending portion 86.

The edge of the spindle 552 adjacent the lever 66 is recessed or stepped. In the arrangement shown in Figure 10, the recess 806 is rectangular in cross-section, such that it comprises a first planar surface that is perpendicular to the rotational axis 76 and a second cylindrical surface that extends around the rotational axis 76. However, the recess 806 could have a different cross-sectional profile such as a triangular cross-sectional profile, thus forming a chamfer. The slot 588 of the fifth alternative high pressure compressor variable vane assembly 542 is defined by the recess 806 of the spindle 552, the bush 64 and the lever 66. The bush 564 is spaced from the lever 66 in an axial direction. A gap 592 is thus formed between the radially extending portion 584 of the bush 564 and the lever 66. The gap 592 extends in an axial direction. The lever 66 defines the first sealing surface 594 against which the seal 90 abuts and the bush 564 defines the second sealing surface 596 against which the seal 90 abuts. An additional sealing surface is provided by the recess 806.

In the arrangement shown in Figure 10, the diameter of the lever 66 is greater than the inner diameter of the bush 564 such that a lower surface of the lever 66 rests on or is adjacent to an upper surface of the bush 564. However, in alternative arrangements, the diameter of the lever 66 is less than or equal to the inner diameter of the bush 564 such that a lower end of the lever 66 can be inserted into the bush 564. In such arrangements, the seal 90 and the interface between the lever 66 and the spindle 552 are positioned within the bush 564.

To assemble the abovementioned high pressure compressor variable vanes, the bush is inserted into the opening 80. The spindle is then inserted into the bush. The seal 90 is then positioned within the slot defined by the spindle. The lever is then attached to the spindle so as to at least partially occlude the slot. Alternatively, the spindle is inserted into the opening 80 before the bush is inserted into the opening 80. The reverse processes may be carried out to disassemble the abovementioned high pressure compressor variable vanes (e.g. for inspection, repair or replacement).

The abovementioned arrangements are beneficial in that they provide good sealing performance without the need for a groove in the spindle in which the seal 90 is located. Accordingly, it is not necessary to stretch the seal 90 around the spindle during assembly and the seal 90 does not need to be as flexible. This results in improved sealing and high temperature performance, lower cost, lower mechanism loads and better reliability in service. Since the seal is located at a position that is external relative to the core of the engine, it is easier to inspect, repair and replace any damaged or worn components without the need for extensive disassembly of the engine, leading to reduced operating cost and disruption risks.

In the abovementioned arrangements, the bush 64 extends into the portion of the opening 80 defined by the flange 62 of the case 54 and the radially extending portion 82 abuts the end of the flange 62. However, in alternative arrangements the case 54 may not comprise a flange 62 and the bush 64 (e.g. the first axially extending portion thereof) may extend into an opening 80 defined by the main body of the case 54. In such arrangements, the radially extending portion 82 may abut the main body of the case 54.

In the high pressure compressor variable vane assemblies 42, 142, 242, 342, 442 described with reference to Figures 4 to 9, the slot 88 is shown as only being partially filled with the seal 90. In the high pressure compressor variable vane assembly 542 described with reference to Figure 10, the slot 588 is shown as being completely filled with the seal 90. However, in any of the arrangements, the slot may be either partially or completely filled with the seal 90.

The outer diameter of the outer surface of the seals 90 of the abovementioned arrangements may be sized so as to form an interference fit with their respective mating components. Alternatively, there may be a small clearance between the outer surface and the mating component. In such arrangements, the pressure differential is sufficiently large to cause the seal to dilate into place.

In the abovementioned arrangements, the variable vane is a high pressure compressor variable vane. However, the variable vane could be different type of variable vane such as a variable inlet guide vane (VIGV).

It has been described that the seal 90 is annular and has a rectangular cross-sectional profile. However, the seal 90 may have any other suitable cross-sectional profile.

For example, the radially outer surface of the seal 90 may be outwardly curved (i.e. barrel-shaped). Alternatively, the seal 90 may be an 0-ring, and, thus, have a circular cross-sectional profile. The seal 90 may be continuous (i.e. be formed of a single annular body) or split (i.e. formed of multiple segments of an annular body).

The vane 44 of the variable vane assembly 42 shown in Figure 4 does not comprise an outer penny. However, the vanes any of the variable vane assemblies 42, 142, 242, 342, 442, 542 described herein may comprise an outer penny (i.e. a disc-shaped element adjacent the spindle). In such arrangements the case may be recessed to receive the outer penny.

It has been described that the lever of each of the abovementioned arrangements partially occludes its respective slot 88. However, in alternative arrangements the lever may fully occlude the slot 88.

It has been described that the vane interfaces with a case. However, in alternative arrangements the vane may interface with any other component of the gas turbine engine, depending on where the vane is located within the gas turbine engine. For example, the vane may instead interface with a shroud ring of the gas turbine engine.

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 (18)

1. CLAIMS1. A variable vane assembly (42, 142, 242, 342, 442, 542), the variable vane assembly (42, 142, 242, 342, 442, 542) comprising a spindle (52, 552) coupled to a lever (66, 366), the spindle (52, 552) at least partially defining a slot (88, 588) within which a seal (90) is positioned, the lever (66, 366) at least partially occluding the slot (88, 588) so as to prevent removal of the seal (90) from the slot (88, 588), the lever defining a sealing surface (94, 394, 494, 594) against which the seal (90) abuts.
2. 2. The variable vane assembly (42, 142, 242, 342, 442, 542) as claimed in claim 1, wherein the spindle (52, 552) and the lever (66) are distinct components.
3. 3. The variable vane assembly (42, 142, 242, 342, 442, 542) as claimed in claim 1 or 2, wherein the spindle (52, 552) and the lever (66) are separable from each other.
4. 4. The variable vane assembly (42, 142, 242, 342, 442, 542) as claimed in any preceding claim, wherein the slot is circular (88, 588) and the seal (90) is annular.
5. 5. The variable vane assembly (42, 142, 242, 342, 442, 542) as claimed in any preceding claim, wherein the variable vane assembly (42, 142, 242, 342, 442, 542) is a high pressure compressor variable vane assembly.
6. 6. The variable vane assembly (42, 142, 242, 342, 442, 542) as claimed in any preceding claim, further comprising a bush (64, 164, 264, 364, 464, 564) within which the spindle (52, 552) is received for rotation about a rotational axis (76), the bush (64, 164, 264, 364, 464, 564) at least partially defining the slot (88, 588).
7. 7. The variable vane assembly (42, 142, 242, 342, 442, 542) as claimed in claim 6, further comprising a case (54, 454) having an opening (80) for receiving the bush (64, 164, 264, 364, 464, 564), the bush (64, 164, 264, 364, 464, 564) comprising a radially extending portion (84, 184, 284, 384, 484, 584) extending in a radial direction away from the rotational axis (76).
8. 8. The variable vane assembly (42, 142, 242, 342) as claimed in claim 6 or 7, wherein the bush (64, 164, 264, 364, 564) defines a further sealing surface (96, 196, 296, 396, 596) against which the seal (90) abuts.
9. 9. The variable vane assembly (42, 142, 242) as claimed in claim 8 when appended to claim 7, wherein the bush (64, 164, 264) comprises an axially extending portion (86, 186, 286) extending in an axial direction away from the radially extending portion (84, 184, 284), the axially extending portion (86, 186, 286) defining the further sealing surface (96, 196, 296).
10. 10. The variable vane assembly (42, 142, 242) as claimed in claim 9, wherein the further sealing surface (96, 196, 296) is disposed radially between the spindle (52) and an outer surface of the lever (66).
11. 11. The variable vane assembly (42, 142, 242) as claimed in claim 9, wherein the outer surface of the lever (66) is disposed radially between the spindle (52) and the further sealing surface (96, 196, 296).
12. 12. The variable vane assembly (342) as claimed in claim 8 when appended to claim 7, wherein the radially extending portion (384) defines the further sealing surface (396).
13. 13. The variable vane assembly (342) as claimed in claim 12, wherein the lever (366) comprises an axially extending flange (900), wherein the flange (900) at least partially occludes the slot (88) so as to prevent removal of the seal (90) from the slot (88) and defines the sealing surface (394) against which the seal (90) abuts.
14. 14. The variable vane assembly (442) as claimed in claim 7, wherein the case (454) comprises an axially extending flange (902), wherein the flange (902) defines a further sealing surface (496) against which the seal (90) abuts.
15. 15. The variable vane assembly (542) as claimed in any of claims 1 to 8, wherein the spindle (552) comprises a recess (806) adjacent the lever (66), wherein the recess (806) at least partially defines the slot (588).
16. 16. The variable vane assembly (542) as claimed in claim 15, wherein the recess (806) has a rectangular profile.
17. 17. A gas turbine engine (10), the gas turbine engine (10) further comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; and a variable vane assembly (42, 142, 242, 342, 442, 542) as claimed in any preceding claim.
18. 18. The gas turbine engine (10) as claimed in claim 17, 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 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 the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.