GB2586269A

GB2586269A - A variable guide vane system

Info

Publication number: GB2586269A
Application number: GB1911750.6A
Authority: GB
Inventors: Troughton William
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2019-08-16
Filing date: 2019-08-16
Publication date: 2021-02-17
Also published as: GB201911750D0

Abstract

A variable guide vane system 50 for a gas turbine engine comprises a variable guide vane 48 which is rotatably positioned within a housing (41, 42, fig.4) and caused to rotate by a force applied to a linkage element 54 acting on a lever arm 412, eg fixed to the outer end of the vane 48. The linkage element 54 deforms plastically (fig.7) in the event of a sufficient impact force being applied to the variable guide vane 48, eg by foreign objects such as ice or birds, so that the linkage element is sacrificed to reduce stress on the remainder of the system. The linkage element 54 may be mounted in a threaded hole in the radially inner surface 56 of a through-hole 55 in a unison ring 51. The linkage element may be pin-shaped or rod-shaped and be made of ductile material, eg 17-4PH or A286 steel. A visual indication of deformation and an alarm may be provided. The variable guide vane may be a variable inlet guide vane (VIGV) or a variable stator guide vane.

Description

A VARIABLE GUIDE VANE SYSTEM

Field of the disclosure

The present disclosure relates to a variable guide vane system of a gas turbine engine, and in particular to a linkage element for attaching a variable guide vane to a unison ring of the system.

Background

Control of fluid gas flows through a gas turbine engine is important to achieve efficiency and performance. For that reason, existing gas turbine engines comprise variable guide vane systems, which include a plurality of variable guide vanes (e.g. variable inlet guide vanes (VIGVs) or variable stator (guide) vanes (VSV)) to direct and present working fluid flows to and from the compressor and turbine stages of the engine, for example. Each guide vane typically comprises an aerofoil section and an integral spindle about which the guide vane can be rotated to modify the incidence of fluid flowing past the plurality of guide vanes. The plurality of variable guide vanes are typically arranged in a row about the circumference of the engine and are connected to a common unison ring via suitable linkages so that, when the unison ring is rotated, so do the vanes about their spindle axis.

During operation of the gas turbine engine, the plurality of variable guide vanes are vulnerable to impact events, such as ice impacts or bird impacts, and the resulting forces can damage the variable guide vanes and their respective linkages. Accordingly, the plurality of variable guide vanes and their corresponding linkages are usually strengthened to cope with such impact forces. However, this comes at a significant increased weight penalty.

The present disclosure seeks to alleviate impact forces on the variable guide vane system.

Summary of the disclosure

According to a first aspect of the disclosure, there is provided a variable guide vane system for a gas turbine engine, the system comprising a variable guide vane, a lever arm and a linkage element. The vane is rotatably positioned within a housing and the lever arm is configured to effect rotation of the vane within the housing by a force applied to the linkage element. The linkage element is configured to deform plastically in the event of an impact force being applied to the vane.

The linkage element may be in the form of a pin or rod shaped element.

The linkage element may be formed of a ductile material.

The linkage element may have a first end that is connected to a unison ring at a connection point on the unison ring. The linkage element may have a free end opposite the first end. The linkage element may be configured to deform plastically in the event of an impact force being applied to the vane by plastically deflecting about the connection point from a first inclination to a second inclination which is angularly offset from the first inclination.

At least a part of the linkage element may be located within a radially extending slot of the unison ring. The slot may be shaped to restrict movement of the linkage element in an axial direction of the engine. The slot may be shaped to allow movement of the linkage element in a circumferential direction of the engine.

The radially extending slot may comprise a first end wall and a second end wall that together delimit a circumferential extent of the slot. The circumferential extent of the slot may be set to limit angular deflection of the linkage element from the first inclination to less than 5 degrees about the connection point.

The slot may be shaped to define a tapered section that is to engage the linkage element when the linkage element plastically deflects in response to an impact force being applied to the vane.

A marker may be located on the unison ring at the same circumferential position as that of the linkage element when the linkage element is disposed at the first inclination, to allow a user to determine whether the linkage element is oriented at the first inclination.

The vane may be one of a set of plural variable guide vanes that are connected to a common unison ring by respective lever arms and linkage elements. Each linkage element may be configured to deform plastically in the event of an impact force being applied to its respective vane.

The variable guide vane system may further comprise an alarm system that includes an alarm which is configured to be turned on in response to plastic deformation of any one of the linkage elements.

The alarm system may comprise a primary circuit that includes the alarm, a primary power source and an electric relay. The electric relay may be configured to open the primary circuit in response to receiving a current at a control contact of the relay and to close the primary circuit in the absence of a current at the control contact. The primary power source may be configured to power the alarm when the primary circuit is closed.

The alarm system may comprise a secondary circuit that includes a secondary power source to supply the current to the control contact of the relay via an electrically conductive wire. The electrically conductive wire may be connected to each of the plurality of linkage elements and may be held in a state of tension such that it will break in response to plastic deformation of one of the linkage elements, to prevent the supply of current to the control contact of the relay and thereby close the primary circuit.

The or each linkage element may be configured to plastically deform in the event of an impact force being applied to the vane in that it has a yield stress that is less than a predetermined, e.g. minimum, impact stress on the linkage element.

The yield stress of the linkage element may be greater than a predetermined, e.g. maximum, non-impact stress on the linkage element. The yield stress of the linkage element may be at least ten times greater than the predetermined, e.g. maximum, non-impact stress.

The yield stress of the linkage element may be less than a yield stress of the lever arm of the variable guide vane system. The yield stress of the linkage element may be less than the yield stress of each other component of the variable guide vane system that connects the guide vane to a unison ring.

The variable guide vane may be a variable inlet guide vane or a variable stator guide vane According to a second aspect of the disclosure, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, the fan comprising a plurality of fan blades, 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, and a variable guide vane system according to the first aspect.

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

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

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

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

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

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). 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: 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 or 390 cm (around 155 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).

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 250 cm to 300 cm (for example 250 cm to 280 cm) 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 320 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 1600 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/Ut1p2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Ufip 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: 0.3, 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).

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, or 17. 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). The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. 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).

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg-ls, 105 Nkg-ls, 100 Nkg-ls, 95 Nkg-ls, 90 Nkg-ls, 85 Nkg-ls or 80 Nkg-ls. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). 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). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 deg C (ambient pressure 101.3kPa, temperature 30 deg 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 TEl 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). 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 blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. 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 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of decent.

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 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 at an altitude that is in the range of from 10000m to 15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m. 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: a forward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of -55 deg C As used anywhere herein, "cruise" or "cruise conditions" may mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach Number, environmental conditions and thrust requirement) for which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to have optimum efficiency.

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.

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 drawinqs

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 the turbine engine; Figure 3 is a partially cut-away view of a gearbox for the gas turbine engine, Figure 4 is a close up sectional side view of a compressor section and a variable guide vane system of the gas turbine engine; Figure 5 is a perspective view of a part of a variable guide vane system, in accordance with an embodiment of the present disclosure; Figure 6 is a sectional side view of a unison ring and a unison ring linkage element of the variable guide vane system of Figure 5, which illustrates the function of the unison ring linkage element when operating under normal operating loads; Figure 7 is a sectional side view of the unison ring and the unison ring linkage element of Figures 5 and 6, which illustrates the function of the unison ring linkage element when operating under impact loads; Figure 8 is a partially cut-away perspective view of the unison ring and the unison ring linkage element of Figures 5 to 7, after plastic deformation in response to an impact load; Figure 9 is a partially cut-away perspective view of a unison ring of a variable guide vane system, in accordance with a further embodiment of the present disclosure; and Figure 10 is a partially cut-away perspective view of a unison ring and a unison ring linkage element of a variable guide vane system, in accordance with a further embodiment of the present disclosure.

Detailed description

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 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 disclosure. 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 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

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 close up sectional side view of a compressor section of a gas turbine engine, such as the low pressure compressor 14 described above with respect to Figures 1 and 2.

The low pressure compressor 14 comprises a radially inner casing 41 and a radially outer casing 42 that are stationary and held apart in the radial direction by a stationary supporting structure 24. The radially inner and outer casings 41, 42 define a flowpath 44 therebetween for receiving the core airflow A to be accelerated and compressed by the low pressure compressor 14, as described with respect to Figure 1.

The compressor 14 also comprises a rotor structure comprising a rotor shaft (not shown) which is centred about the principal rotational axis 9, and a disc 46 mounted thereon (by conventional means) for rotation about the principal rotational axis 9. Attached along the radially outer circumference of the disc 46 is a plurality of compressor blades 47 extending from the disc 46 in the radial direction, although only one such blade 47 is shown for clarity. During operation, the rotor structure is driven at high speed, e.g. by the low pressure turbine (not shown), such that the compressor blades 47 rotate and impart energy to the airflow A passing through the compressor 14.

The low pressure compressor section 14 comprises a variable guide vane 48 disposed between, and housed by, the inner and outer casing 41, 42 such that an aerofoil portion thereof is disposed within the casing flowpath 44. The variable guide vane 48 is a variable inlet guide vane in that it is located at the entrance to the compressor, i.e. upstream of the compressor blades (and, e.g., downstream of the stationary supporting structure 24). Although only a single variable inlet guide vane 48 is shown in Figure 4, it will be appreciated that the compressor 14 will include a plurality of such variable inlet guide vanes 48 arranged circumferentially about the principal rotational axis 9 between the inner and outer casings 41, 42. Furthermore, although the disclosure is described herein with reference to a variable inlet guide vane 48, the disclosure is applicable more widely to any type of variable guide vane, such as a variable stator vane which is arranged as part of a row of variable stator vanes in the compressor 14 itself.

The variable inlet guide vane 48 is rotatable about a rotational axis for controlling the airflow within the flowpath 44 to achieve efficient engine and compressor operation. For example, at low engine speed, the variable inlet guide vane 48 may be rotated to reduce the incidence of airflow onto the compressor blades 47 to tolerable angles. In order to achieve such rotation, each of the variable inlet guide vanes 48 are rotationally mounted or housed by the inner and outer casing 41, 42 by radially inner and outer radially extending spigots 43 and 45 which extend at the respective radially inner and outer spanwise ends of the vanes 48. The inner spigot 43 is rotatably mounted within a stationary mechanical bush 49 which is fixed to or formed by the inner casing 41 and surrounding engine support structure. The radially outer spigot 45 is rotatably mounted within another stationary mechanical bush 410 provided within a radially extending boss 411 of the radially outer casing 42. The longitudinal axis of each of the radially inner spigot 43 and the radially outer spigot 45 are aligned so as to define a spindle about which the guide vane can be rotated to modify the incidence of fluid flowing past the plurality of guide vanes.

Fitted to the radially outer end of the radially outer spigot 45 is a first end 414 of a rigid lever arm 412 which is fixed relative to the spigot and which extends substantially perpendicular to the spindle axis which defines the rotational axis of the variable guide vane 48. The first end 414 of the lever arm 412 is fixedly secured to the spigot 45 by means of a fastening bolt 413. A second end 415 of the lever arm 412 opposite the first end 414, however, is pivotably connected to a linkage element of a common unison ring (see reference 510 of Figure 5). The unison ring itself is rotatably mounted about the principal rotational axis 9 so that when the unison ring is rotated by suitable actuation means, a torque is applied to the lever arm 412 at the second end 415 via the unison ring linkage element so that the plurality of vanes 48 are caused to rotate within their respective bushes 48, 49 about their spindle axis.

In previously considered arrangements, the components of the variable guide vane system are designed and manufactured with a sufficiently high yield strength to elastically absorb the loads that they are expected to face during operation of the gas turbine engine. For example, the unison ring linkage element must be strong enough to withstand not only the forces acting upon it during normal operation, but also forces that occur as a result of a torque applied to the aerofoil portion of the guide vane as a result of bird or ice impact events, without plastically deforming or breaking. In that regard, the forces acting on the unison ring linkage element as a result of an impact force on the aerofoil section of the guide vane can have a magnitude that is as much as one-hundred times greater than the forces acting on the linkage element during normal operation. While strengthening the various elements of the variable guide vane system ensures that they are able to withstand applied forces without failure or plastic deformation, i.e. by elastically deforming, this comes with a significant increased weight and size penalty.

In contrast to these previously considered arrangements, the present disclosure ensures that the variable guide vane system can absorb the energy imparted on the system by impact events without breaking by connecting the lever arm to the unison ring via a unison ring linkage element which is configured to plastically deform in the event of an impact force being applied to the vane, as will now be described with respect to Figures 5 to 8.

Figure 5 is a perspective view of a part of a variable guide vane system 50, in accordance with an embodiment of the present disclosure. The variable guide vane system 50 comprises a unison ring 51, a variable guide vane 48 and a lever arm 412 as described above with respect to Figure 4 (and for that reason like features in Figures 4 and 5 are labelled with like reference numerals).

It will be appreciated that the housing structure within which the variable guide vane 48 is mounted, i.e. the inner and outer casing 41, 42 and bushes 49, 410 etc. as described above, is not shown for ease of illustration. Furthermore, although only one variable guide vane 48 is shown in Figure 5, it will be appreciated that in practice there may be a plurality of such vanes arranged in a row extending about the circumference of the engine and each such vane will be connected to the common unison ring via respective lever arms and unison ring linkage elements.

The lever arm 412 is in the form of a plate which is elongate between the first end 414 and the second end 415 of the lever arm 412. The second end 415 of the lever arm 412 has a radially extending through hole (not shown) that receives a linkage element 54 of the unison ring 51 therethrough. The linkage element 54 of the unison ring 51 is correspondingly shaped as a pin or rod, i.e. an elongate, straight bar of substantially circular cross-section, so that it is snugly received by the through hole of the lever arm 412. The through hole has a spherical bearing 59 to aid rotation of the lever arm 412 about the linkage element 54.

The unison ring 51 has a plurality of axially extending through-holes 55, one for each variable guide vane 48 that is to be attached to the unison ring 51, wherein each through-hole 55 is defined by a radially inner surface 56 and a radially outer surface 57 that extend in an arc along the circumferential direction 59 of the engine. The linkage element 54 is removably fastened to a connection point on, and extends from, the radially inner surface 56 of the through-hole 55. The linkage element may be removably fastened to the unison ring at the connection point using any known fastening means. However, in this embodiment the linkage element 54 has a threaded portion (not shown) that engages and mates with a correspondingly threaded hole (not shown) on the radially inner surface 56 of the through-hole 55. In this way, the linkage element 54 may be fastened and unfastened by rotation of the linkage element 54 when the threaded portion is received in the threaded hole.

When fastened to the unison ring 51, the linkage element 54 has a free end 511 at its radially outermost side opposite the threaded end. In this embodiment, the free end 511 extends through a radially extending slot 58 through the unison ring 51 between the radially outer surface 57 of the through-hole 55 and the radially outermost surface 510 of the unison ring 51 itself. In this example, the free end 511 extends radially outwards of a radially outermost surface 510 of the unison ring 51.

In use, the variable guide vane 48 will experience loads acting upon the aerofoil portion disposed within the casing flowpath (reference 44 in Figure 4). These may be loads that are applied during normal operating conditions of the engine, where the forces acting on the aerofoil portion are applied by the core airflow A received within the casing flowpath, or may be loads that arise as a result of impact forces imparted by foreign objects impacting the aerofoil portion. In either case, as illustrated in Figure 5, a force 512 acting upon the aerofoil portion will in turn generate a torque 513 that acts upon the spigot 45 and thus the lever arm 412 of the variable guide vane 48. The torque 513 on the lever arm 412 will correspondingly apply a force 514 on the unison ring linkage element 54, where the force 514 on the unison ring linkage element 54 is directed in substantially the circumferential direction 59.

Although the unison ring 51 is configured to be rotated in the circumferential direction 59 by actuation means, an impact force 512 on the aerofoil portion of the vane 48 is usually unable to generate a force 514 on the unison ring 51 that is of sufficient magnitude to overcome the inertia of the unison ring 51 (and the plurality of other guide vanes attached thereto). As a result, the variable guide vane 48, lever arm 412 and linkage element 54 are subjected to most, if not all, of the impact loads.

In order for the variable guide vane system to adequately deal with the stresses caused by such impact loads, the unison ring linkage element 54 is configured to plastically deform in response to an impact force. To achieve this, the linkage element 54 has a yield stress (also referred to in the art as yield strength), i.e. the stress at which an object stops deforming elastically and instead begins to deform plastically, which is equal to or less than a predetermined impact stress for the engine, which is the stress that the linkage element 54 will be subjected to as a result of an impact force being applied to the variable guide vane 48 during an impact event. The predetermined impact stress may be the minimum stress to which the linkage element 54 may be subjected during an impact event. The yield stress of the linkage element 54 is also less than the yield stress of the other components of the variable guide vane system that connects a variable guide vane to the unison ring 51, such as the lever arm 412. This arrangement is in contrast to previously considered arrangements wherein each linkage element and its associated lever arm have yield stresses that are significantly above the impact stresses to be faced by them during an impact event.

In this way, the linkage element 54 is effectively sacrificed during an impact event to reduce the loads and thus stresses on the remainder of the system, such as the lever arm 412 and variable guide vane 48. This may be advantageous in that it allows one to increase the impact force that the variable guide vane system can withstand, but without having to increase the strength (and thus weight etc.) of the other elements.

The yield stress is also set at a value that is greater than a predetermined non-impact stress to be faced by the linkage element 54 during normal (i.e. non-impact) operation of the variable guide vane system, to ensure that the linkage element 54 remains functionally operable (without plastic deformation) at normal operating load conditions. The predetermined non-impact stress may be the maximum stress on the linkage element 54 during normal operation of the vane 48.

The stresses on the linkage element 54 (and other components of the variable guide vane system) during normal operation and during impact events will vary from engine-to-engine. This is because the stresses depend on a number of factors that are unique to each engine configuration, such as the aerodynamic profile of the guide vane, the geometry of the variable guide vane lever arm and the architecture of the engine itself (e.g. the extent to which the vane is exposed within the flowpath). However, the stresses can be readily (pre-)determined using conventional methods that are known in the art, such as elastic/plastic finite element analysis, foreign object trajectory analysis etc. That is, for any given engine configuration, the impact stress and the non-impact stress on the linkage element 54 may be pre-determined using conventional means, and the linkage element 54 is manufactured to have a yield stress that is set to a value between those two pre-determined values, accordingly.

The impact stresses acting on the linkage element 54 during an impact event are known to be multiple orders of magnitude greater than the non-impact stresses acting on the linkage element 54 during normal operation of the engine. Therefore, an appropriate value for the yield stress of the linkage element 54 may be a value that is at least ten, and in embodiments one-hundred, times greater than the, e.g. maximum, non-impact stress on the linkage element 54 during normal (i.e. non-impact) operation Any number of conventional methods known in the art may be used to engineer the linkage element 54 to a desired yield stress. The linkage element 54 may be formed of 17-4PH steel or A286 steel, which are known types of steel, for example. Further, the yield stress of those materials could be modified by a number of processing procedures, e.g. by altering dislocation density, impurity levels etc. of the materials. Additionally or alternatively, the yield stress of the linkage element 54 may be set by optimising the geometry of the linkage element. For example, the length and thickness of the linkage element 54 could be set to achieve a specific percentage of plasticity under an impact force, for a specific material type.

The function of the unison ring linkage element will now be described with respect to Figures 6 to 8 (in which like features are labelled with like reference numerals). In all of those Figures, the unison ring 51 and linkage element 54 are shown without the guide vane 48 or lever arm 412, for ease of illustration, but it will be appreciated that in practice the linkage element 54 is attached to the lever arm 412 in the manner described above.

Figure 6 illustrates the function of the linkage element 54 of the unison ring when under normal operating loads. In this arrangement the linkage element 54 has a first inclination 61 that is aligned with a radial direction 53 of the engine. The linkage element 54 is configured to remain at the first inclination 61 when the aerofoil portion of the guide vane is subjected to normal load conditions.

In particular, under normal load conditions the stress on the linkage element 54 as a result of force 60 will correspond to non-impact stresses that, as mentioned above, are orders of magnitude less than the yield strength of the linkage element 54. Therefore, the linkage element 54 will operate significantly below its yield point and therefore absorb the force 60 with no plastic deformation. At most, the linkage element 54 will elastically deflect from the first inclination 61 in response to the force 60, before returning back to the first inclination 61 after the force 60 has been removed.

Figure 7 illustrates the function of the linkage element 54 when subjected to impact loads.

During an impact event, when an object comes forcibly into contact with the aerofoil portion of the variable guide vane within the casing flowpath, a resultant force 70 on the unison ring linkage element 54 will cause it to be stressed beyond the yield strength of the linkage element 54. The linkage element 54 will therefore operate at or above its yield stress point and cause the free end 411 of the linkage element 54 to deflect and yield (i.e. plastically deform) in a circumferential direction 59 of the engine corresponding to the direction of the force 70, from the first inclination 61 to a second, different inclination 71. The linkage element 54 will deflect about its connection point to the unison ring 51 to define an angle relative to the first inclination 61.

The material of the linkage element 54 is ductile to ensure that the linkage element 54 will undergo significant plastic deformation before rupture. In particular, the linkage element 54 has an area-reduction-to-rupture value of at least 20%.

As best shown in Figure 8, which illustrates the unison ring and linkage element 54 after plastic deformation, the radially extending slot 58 through which the free end 511 of the linkage element 54 extends is shaped to define a race or channel which restricts movement of the linkage element 54 in an axial direction 52 of the engine but allows for the deflection of the linkage element 54 along the slot 58 in the circumferential direction 59.

In that regard, the slot 58 comprises a first side wall 81 and a second side wall 82 that extend parallel along an arc in the circumferential direction 59. The first side wall 81 and the second side wall 82 are separated in the axial direction 52 by an axial extent 83 that substantially corresponds to the diameter 84 of the linkage element 54. The axial extent 83 may substantially correspond to the diameter of the linkage element 54 in that the slot 58 and the linkage element have an engineering fit, e.g. a clearance fit, in the axial direction 59.

At circumferentially opposite ends of the slot 58 is a first end wall 85 and a second end wall 86. The end walls 85, 86 serve to delimit the circumferential extent of the slot 58 and thus the extent by which the linkage element 54 is allowed to angularly deform away from the first inclination 61. The linkage element 54, when caused to deform under the force 70 of an impact event, may abut one of the end walls (the first end wall 85 in this example) to prevent excessive deformation. Each end wall 85, 86 is shaped to conform to the shape of the linkage element 54, for uniform engagement with the linkage element 54. In particular, the rounded walls 85, 86 have a radius of curvature that matches the radius of the linkage element 54.

It will be appreciated that the optimum circumferential extent for the slot 58 to prevent excessive deformation will vary from system to system, and will depend on the length of the linkage element 54 etc. However, in embodiments the circumferential extent of the slot 58 will be set to limit the angular displacement of the linkage element from the first inclination 61 to less than 5 degrees, preferably 2 or 3 degrees. By limiting the extent of plastic deformation in this manner, it may be possible to prevent the linkage element 54 from deflecting beyond a predetermined point of deflection that would cause the linkage element 54 to break. That is, the end walls 85, 86 of the slot 58 will limit the extent of strain that is induced in the linkage element 54 as a result of the force 70 acting on the linkage element 54 to levels that are below the level of strain that would lead to fracture of the linkage element 54.

Figure 9 is a partially cut-away view of the unison ring 51 in accordance with an embodiment of the present disclosure. It will be appreciated that the unison ring 51 is shown without the linkage element for ease of illustration only.

In the embodiment of Figure 9, the slot 58 includes a tapered section configured to receive the linkage element when the linkage element plastically deforms in response to an impact force, in order to increase the impact force required to bend the linkage element throughout its motion. In the embodiment of Figure 9, the tapered section is at or proximate a distal end of the slot 58, particularly towards the first end wall 85 of the slot 58. That is, the axial extent of the slot 58 tapers in the circumferential direction 59 away from a central circumferential point of the slot 58 at which the axial extent is at its maximum (i.e. the circumferential position within the slot 58 at which the linkage element is located when at the first inclination). The axial narrowing or taper is defined, in this example, by a first tapered wall 91, which extends between the first side wall 81 and the first end wall 85, and a second tapered wall 92, which extends between the second side wall 82 and the second end wall 85. The first tapered wall 91 and the second tapered wall 92 are planar walls that converge to restrict the slot in the axial direction. The axial extent within the tapered section, i.e. between the tapered walls 91, 92, is less than the diameter of the linkage element in order to resist plastic deformation of the linkage element 54.

In the manner described above, the slot 58 and the linkage element 54 may have an engineering clearance fit in the axial direction 59, when the linkage element is disposed at the first inclination 61 described above with respect to Figure 6, but an engineering transition fit in the axial direction, when the linkage element 54 is within the tapered section of the slot 58.

Although the tapered section has been described above as being formed by two tapered walls that together reduce the axial extent of the slot in the circumferential direction, it will be appreciated that the tapered section may take any form or shape that is suitable for reducing the axial extent of the slot in the circumferential direction. For example, the side walls 81 and 82 of the slot may be non-planar or otherwise shaped so that the axial extent of the slot varies in the circumferential direction, or the slot may be provided with only one planar tapered wall, e.g. the first tapered wall 91 described above, for this purpose.

Figure 10 is a partially cut-away view of a unison ring 51 and linkage element 54 in accordance with an embodiment of the present disclosure.

In the embodiment of Figure 10, the unison ring 51 is provided with an indicator of the current state of the linkage element 54. The indicator includes a first visual marker 101 which indicates to a user the correct or desired position of the linkage element 53. To this effect, the first marker 101 is located on the unison ring 51 at the same circumferential position as the linkage element 54 when the linkage element 54 is disposed at the first inclination 61 during normal load conditions, as described above with respect to Figure 6. The first marker 101 is in the form of a straight line which is aligned with the axial direction of the engine, however the first marker may take any shape, e.g. a dot, which is suitable for indicating the correct position of the linkage element 54.

The first marker 101 enables a user, upon inspection of the unison ring 51 during routine maintenance work, for example, to quickly identify whether the free end of the linkage element 54 is aligned with, i.e. at the same circumferential position as, the first marker 101, or whether the free end of the linkage element is out of alignment with the first marker 101, i.e. at a different circumferential position to, the first marker 101. If the linkage element 54 is out of alignment with the first marker 101, then this may be taken as an indication that the linkage element 54 is functionally inoperable due to plastic deformation of the linkage element. In that case, a user will know to conduct maintenance of the variable guide vane system, e.g. by replacing the deformed linkage element with a new one.

A second visual marker 102, which in this example takes the form of an axially aligned straight line, is provided on the radially outer surface of the free end 511 of the linkage element 54, to further aid inspection. In particular, the first marker 101 and the second marker 102 will be in alignment when the linkage element 54 is correctly located at the first inclination, but out of alignment when the linkage element has plastically deformed as a result of an applied impact force.

In the manner described above, the visual indicator provides the user with information regarding the technical condition of the linkage element and wider variable guide vane system and enables the user to take appropriate action to ensure the correct and proper operation of the system.

Additionally or alternatively to physical markers, the variable guide vane system may be provided with electrical means of indicating to a user the current state or condition of the linkage element. In particular, the variable guide vane system may be provided with an alarm system in which an alarm is turned on in response to plastic deformation of one of the linkage elements.

In an embodiment, the alarm system includes a primary alarm circuit that includes a primary power source, an electric relay and an alarm that is powered by the primary power source. The alarm may be in the form of a Light Emitting Diode, for example, which provides a visual indication to the user. Additionally or alternatively the alarm may output sound to the user.

The relay is configured to be switched between an ON position, at which the alarm circuit is closed to initiate the alarm, and an OFF position, at which the alarm circuit is open to stop the alarm. The position of the relay is controlled by the application of a current to a control contact of the relay from a secondary power source that is connected to the relay via an electrically conductive wire (and in that sense the relay forms part of a secondary circuit in addition to the alarm circuit). The relay may be an electromagnet or transistor based relay which is configured to remain at the OFF position when an electric current is received by the relay from a secondary power source, but to remain at the ON position when an electric current is not received by the relay from the secondary power source.

The electrically conductive wire that forms part of the secondary circuit extends between, and is connected to, each of the plurality of linkage elements of the unison ring. The wire is connected to each linkage element by mechanical means, such as a bolt in the free end of the linkage element. The wire is held in a state of tension between each respective pair of circumferentially adjacent linkage elements, so that if one of the linkage elements were to plastically deform as a result of an impact force in the manner described above, the conductive wire will break.

When the plurality of linkage elements are oriented at their respective first inclinations, the secondary circuit is closed and the current through the secondary circuit powers an electromagnet of the relay to an OFF position so that the alarm circuit remains open. However, when one or more of the plurality of linkage elements plastically deform from their respective first inclinations in response to an impact force, the force applied on the electrically conductive wire of the secondary circuit will cause it to break, thereby opening the secondary circuit. When the secondary circuit is open, the supply of current to the relay will be cut off so as to turn the relay to the ON position so that the alarm circuit is closed, thereby triggering the alarm.

It will be understood that the disclosure is not limited to the embodiments described above 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

CLAIMS1. A variable guide vane system (50) for a gas turbine engine (10), the system (50) comprising a variable guide vane (48), a lever arm (412) and a linkage element (54); wherein: the vane (48) is rotatably positioned within a housing (41, 42) and the lever arm (412) is configured to effect rotation of the vane (48) within the housing (41, 42) by a force applied to the linkage element (54); and the linkage element (54) is configured to deform plastically in the event of an impact force being applied to the vane (48).
2. The variable guide vane system (50) of Claim 1, wherein the linkage element (54) is in the form of a pin or rod shaped element.
3. The variable guide vane system (50) of Claim 1 or Claim 2, wherein the linkage element (54) is formed of a ductile material.
4. The variable guide vane system (50) of any one of Claims 1 to 3, wherein: the linkage element (54) has a first end that is connected to a unison ring (51) at a connection point on the unison ring (51); the linkage element (54) has a free end (511) opposite the first end; and the linkage element (54) is configured to deform plastically in the event of an impact force being applied to the vane (48) by plastically deflecting about the connection point from a first inclination (61) to a second inclination (71) which is angularly offset from the first inclination (61).
5. The variable guide vane system (50) of Claim 4, wherein: at least a part of the linkage element (54) is located within a radially extending slot (58) of the unison ring (51); and the slot (58) is shaped to restrict movement of the linkage element (54) in an axial direction (59) of the engine (10), but to allow movement of the linkage element (54) in a circumferential direction (55) of the engine (10).
6. The variable guide vane system (50) of Claim 5, wherein: the radially extending slot (58) comprises a first end wall (85) and a second end wall (86) that together delimit a circumferential extent of the slot (58); and the circumferential extent of the slot (58) is set to limit angular deflection of the linkage element (54) from the first inclination (61) to less than 5 degrees about the connection point.
7. The variable guide vane system (50) of Claim 5 or Claim 6, wherein the slot (58) is shaped to define a tapered section that is to engage the linkage element (54) when the linkage element (54) plastically deflects in response to an impact force being applied to the vane (48).
8. The variable guide vane system (50) of any one of Claims 4 to 7, wherein a marker (101) is located on the unison ring (51) at the same circumferential position as that of the linkage element (54) when the linkage element (54) is disposed at the first inclination (61), to allow a user to determine whether the linkage element (54) is oriented at the first inclination (61).
9. The variable guide vane system (50) of any one of the preceding claims, wherein: the vane (48) is one of a set of plural variable guide vanes (48) that are connected to a common unison ring (51) by respective lever arms (412) and linkage elements (54); each linkage element (54) is configured to deform plastically in the event of an impact force being applied to its respective vane (48); and the variable guide vane system (50) further comprises an alarm system that includes an alarm which is configured to be turned on in response to plastic deformation of any one of the linkage elements (54).
10. The variable guide vane system (50) of Claim 9, wherein the alarm system comprises: a primary circuit that includes the alarm, a primary power source and an electric relay, wherein: the electric relay is configured to open the primary circuit in response to receiving a current at a control contact of the relay and to close the primary circuit in the absence of a current at the control contact; and the primary power source is configured to power the alarm when the primary circuit is closed; and a secondary circuit that includes a secondary power source to supply the current to the control contact of the relay via an electrically conductive wire, wherein: the electrically conductive wire is connected to each of the plurality of linkage elements (54) and is held in a state of tension such that it will break in response to plastic deformation of one of the linkage elements (54), to prevent the supply of current to the control contact of the relay and thereby close the primary circuit.
11. The variable guide vane system (50) of any one of the preceding claims, wherein the or each linkage element (54) is configured to plastically deform in the event of an impact force being applied to the vane (48) in that it has a yield stress that is less than a predetermined impact stress on the linkage element (54).
12. The variable guide vane system (50) of Claim 11, wherein the yield stress of the linkage element (54) is greater than a predetermined non-impact stress on the linkage element (54).
13. The variable guide vane system (50) of any one of the preceding claims, wherein the yield stress of the linkage element (54) is less than a yield stress of the lever arm (412) of the variable guide vane system (50).
14. The variable guide vane system (50) of any one of the preceding claims, wherein the variable guide vane (48) is a variable inlet guide vane or a variable stator guide vane.
15. A gas turbine engine (10) for an aircraft 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 guide vane system (50) as claimed in any one of the preceding claims.