GB2591104A

GB2591104A - Component for mounting arrangement, mounting arrangement and gas turbine engine comprising mounting arrangement

Info

Publication number: GB2591104A
Application number: GB2000581.5A
Authority: GB
Inventors: Cousins Daniel; Gupta Alok; Howe Samuel
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2020-01-15
Filing date: 2020-01-15
Publication date: 2021-07-21
Also published as: GB202000581D0

Abstract

A component 100 has a first abutment 110 spaced apart from a second abutment 120 with a surface (fig.6,122) facing the first abutment and an opposing surface (fig.6,124). Between abutments lies a mount 130, and between each abutment and the mount a respective spring 140,150. A conduit (fig.6,160) is connected to the second abutment, with a through-hole (fig.6,170) extending through the first abutment, mount, and springs. When used in a mounting arrangement 300 of a gas turbine engine a fastener shaft 332 passes through the through-hole, with a head 330 contacting the first abutment, a first part 310 couples to the mount and a second part 320, coupled to the fastener 330, contacts the second abutment, so the parts are isolated from on another. The mount may have a number of through-holes 132. The conduit is preferably cylindrical, extending from the first abutment; the abutments and/or mount may be disc-shaped. Preferably, each spring connects a respective abutment to the mount, and at least partially surrounds the conduit; the mount preferably surrounds the conduit. A unit (fig.7,340) may be mounted on the first part. The component may be integrally formed as one element by an additive manufacturing process.

Description

COMPONENT FOR MOUNTING ARRANGEMENT, MOUNTING

ARRANGEMENT AND GAS TURBINE ENGINE COMPRISING MOUNTING

ARRANGEMENT

FIELD OF THE DISCLOSURE

The present disclosure relates to a component for mounting arrangements, and in particular to mounting arrangements within a gas turbine engine for an aircraft.

BACKGROUND

Gas turbine engines may include regions which operate at high temperatures, such as an engine core and a casing of the engine core. Parts that are hard mounted to the casing are affected by vibrations imparted onto them from the casing and subsequently, their movement causes reaction loads to be imparted back onto the casing.

In order to protect the parts from vibrations, an anti-vibration mount (AVM) is used to damp or isolate vibrations experienced by the parts. The AVMs reduce the vibrations imparted to the parts from the casing thereby minimizing damage to the parts due to vibrations. However, the AVMs are typically made from rubber, which cannot be used on the engine core and the casing of the engine core, as the high temperatures may deform or melt the rubber in the AVM. As the rubber AVMs cannot be used in this high temperature environment, the parts and/or their mounting brackets are generally made larger and stronger (and therefore heavier) in order to survive vibrations. This may increase an overall engine mass.

It is an aim of the present disclosure to at least partially address the above problems.

SUMMARY

According to a first aspect there is provided a component for a mounting arrangement of a gas turbine engine. The component includes a first abutment member and a second abutment member spaced apart from the first abutment member. The second abutment member includes a first surface facing the first abutment member and a second surface opposite to the first surface. A mounting member is disposed between the first abutment member and the second abutment member. A first spring member is disposed between the first abutment member and the mounting member and a second spring member is disposed between the mounting member and the second abutment member. A conduit includes a through-hole extending at least through the first abutment member, the mounting member, the first spring member and the second spring member. The conduit is connected to the second abutment member. When the component is in the mounting arrangement, the component is configured such that a shaft of a fastener passes through the through-hole, a head of the fastener abuts the first abutment member, a first part to be mounted by the mounting arrangement is coupled to the mounting member and a second part to be coupled to the fastener abuts the second abutment member.

In some cases, the abutment of the second abutment member with the second part may allow transmission of loads from the fastener through the conduit to the second part while allowing the mounting member to move freely along a length of the component. This may isolate the first part from vibrations imparted from the second part.

The vibrations imparted from the first part may load the first and second spring members in opposite directions. The loading of the first and second spring members in opposite directions may cause the first part to remain substantially at its datum position.

In some cases, the component may provide vibration isolation in high temperature environments without requiring the first part and/or the second part to have larger size and weight. This may decrease an overall mass of a system including the first part and the second part.

In some embodiments, the mounting member may include a plurality of mounting holes extending therethrough.

In some embodiments, the conduit may be substantially cylindrical in shape.

In some embodiments, the first abutment member, the second abutment member and/or the mounting member may be substantially disc-shaped.

In some embodiments, the conduit may extend from the first abutment member. In some other embodiments, the conduit may extend from the second abutment member. Alternatively, the conduit may extend beyond the first abutment member.

In some embodiments, the conduit may be integrally formed with the first abutment member and the second abutment member.

In some embodiments, the first spring member may connect the first abutment member to the mounting member. Alternatively, the first spring member may be 15 connected to intermediate members disposed between the first abutment member and the mounting member.

In some embodiments, the second spring member may connect the mounting member to the second abutment member. Alternatively, the second spring 20 member may be connected to intermediate members disposed between the mounting member and the second abutment member.

In some embodiments, each of the first spring member and the second spring member may at least partially surround the conduit.

In some embodiments, the mounting member may surround the conduit.

In some embodiments, the first abutment member, the second abutment member, the mounting member, the first spring member and the second spring 30 member may be integrally formed.

In some embodiments, the component may be formed as one element by an additive manufacturing process.

In an example, the additive manufacturing process may include additive layer manufacturing (ALM). The overall size and shape of the component may be varied conveniently using ALM as per application requirements.

According to a second aspect there is provided a mounting arrangement for vibrationally isolating a first part from a second part of the gas turbine engine. The mounting arrangement includes the component of the first aspect. The first part is coupled to the mounting member and the second part abuts the second abutment member. The mounting arrangement further includes a fastener including a shaft extending through the through-hole of the conduit and received in the second part and a head abutting the first abutment member.

In some embodiments, the mounting arrangement may include a unit mounted on the first part.

According to a third aspect there is provided a gas turbine engine for an aircraft. The gas turbine engine includes the mounting arrangement of the second aspect.

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

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

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for 5 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 10 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 15 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 On 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 01 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- 25!imitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Ut1p2, 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 5 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 Nkg1s, 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 10 may form upper or lower bounds), for example in the range of from 80 Nkg-ls to 100 Nkg-ls, or 85 Nkg-ls to 95 Nkg-ls. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

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

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TEl 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, m id-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint -in terms of time and/or distance -between top of climb and start of descent. Cruise conditions thus define an operating point of, the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

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

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

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

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

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

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

According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise 20 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 25 operation according to this aspect may include (or may be) operation at the m id-cruise of the aircraft, as defined elsewhere herein.

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

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; Figure 4 is an oblique view of a component of the present disclosure; Figure 5 is a side view of the component of the present disclosure; Figure 6 is a cut-away view of the component of the present disclosure; Figure 7 is an isometric view of a mounting arrangement of the present disclosure; Figure 8 is a cut-away view of the mounting arrangement of the present 15 disclosure; and Figure 9 is a cut-away view of a component of another embodiment of the present disclosure.

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 10 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

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

Note that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor". Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 5 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 10 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 20 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.

However, it will be appreciated that the present disclosure will find wider 5 application than merely as a component for mounting arrangements in a gas turbine engine. In some embodiments, the present disclosure may find application in nuclear, industrial power systems, automotive industry, construction and heavy industries, where vibration isolators and dampers are required to be used.

Figures 4 and 5 show an example component 100 for a mounting arrangement of the gas turbine engine 10 in accordance with an embodiment of the present disclosure. The component 100 may be part of an anti-vibration mounting arrangement. The component 100 defines a longitudinal axis "LA" along its length. The component 100 includes a first abutment member 110 and a second abutment member 120 spaced apart from the first abutment member 110. Specifically, the first abutment member 110 and the second abutment member 120 are spaced apart from each other relative to the longitudinal axis "LA". The first abutment member 110 includes a first surface 112 facing the second abutment member 120 and a second surface 114 opposite to the first surface 112. The second abutment member 120 includes a first surface 122 facing the first abutment member 110 and a second surface 124 opposite to the first surface 122. Regardless of the specific shapes of the first and second abutment members 110, 120, the second surface 114 of the first abutment member 110 and the second surface 124 of the second abutment member 120 are substantially planar. In some embodiments, the second surfaces 114, 124 of the respective first and second abutment members 110, 120 are substantially parallel to each other. Further, the first surfaces 112, 122 of the respective first and second abutment members 110, 120 may also be planar and substantially parallel to each other.

The component 100 further includes a mounting member 130 disposed between the first abutment member 110 and the second abutment member 120. Specifically, the mounting member 130 is disposed between the first abutment member 110 and the second abutment member 120 relative to the longitudinal axis "LA". In some embodiments, the mounting member 130 includes a plurality of mounting holes 132 extending therethrough. In the illustrated embodiment, the component 100 includes four mounting holes 132 that are arranged 5 uniformly on the mounting member 130. However, the mounting member 130 may include any number of mounting holes 132 that can be arranged uniformly or non-uniformly as per application requirements. The first surfaces 112, 122 of the respective first and second abutment members 110, 120 face towards the mounting member 130. Further, the second surfaces 114, 124 of the respective 10 first and second abutment members 110, 120 face away from the mounting member 130.

The mounting member 130 further includes a first surface and a second surface opposite to the first surface. The first surface faces the first abutment member 110. The second surface faces the second abutment member 120. In some embodiments, each of the first and second surfaces of the mounting member 130 is substantially planar.

In some embodiments, the first abutment member 110, the second abutment member 120 and/or the mounting member 130 are substantially disc-shaped. In the illustrated embodiments of Figures 4 and 5, each of the first abutment member 110, the second abutment member 120 and the mounting member 130 is disc-shaped. However, each of the first abutment member 110, the second abutment member 120 and the mounting member 130 may have any suitable shape, such as elliptical, oval, rectangular, polygonal, and so forth.

The component 100 further includes a first spring member 140 and a second spring member 150. The first spring member 140 is disposed between the first abutment member 110 and the mounting member 130. Specifically, the first spring member 140 is disposed between the first abutment member 110 and the mounting member 130 relative to the longitudinal axis "LA". The second spring member 150 is disposed between the mounting member 130 and the second abutment member 120. Specifically, the second spring member 150 is disposed between the second abutment member 120 and the mounting member 130 relative to the longitudinal axis "LA". In some embodiments, the first spring member 140 connects the first abutment member 110 to the mounting member 130. In some other embodiments, the first spring member 140 may be connected to intermediate members (not shown) disposed between the first abutment member 110 and the mounting member 130. In some embodiments, the second spring member 150 connects the mounting member 130 to the second abutment member 120. In some other embodiments, the second spring member 150 may be connected to intermediate members (not shown) disposed between the mounting member 130 and the second abutment member 120. The first and second spring members 140, 150 are not limited to any particular type of spring. Each of the first and second spring members 140, 150 may be any resilient or biasing member. However, as shown in Figures 4 and 5, each of the first and second spring members 140, 150 may have a curved shape. In some embodiments, a shape, a length, an orientation and a thickness of each of the first and second spring members 140, 150 may be adjusted to change the stiffness of the first and second spring members 140, 150 and therefore a natural frequency of the mounting arrangement including the component 100. This may allow specific frequencies for a particular part to be isolated. In some embodiments, a larger movement of each of the first and second spring members 140, 150 may enhance a structural damping of the mounting arrangement. In some cases, the first and second spring members 140, 150 may damp and/or isolate vibrations to reduce overall response of the mounting arrangement at higher order frequencies.

The component 100 further includes a conduit 160. The conduit 160 includes a through-hole 170 extending at least through the first abutment member 110, the mounting member 130, the first spring member 140 and the second spring member 150. The conduit 160 is connected to the second abutment member 120. Each of the first abutment member 110, the second abutment member 120 and the mounting member 130 may have a central through-aperture to at least partially receive the conduit 160 therein. In some embodiments, the conduit 160 may be centrally located in the component 100. The conduit 160 extends along the longitudinal axis "LA" of the component 100. In some embodiments, the conduit 160 may have substantially cylindrical symmetry about a central axis through the through-hole 170.

In some embodiments, the first and second abutment members 110, 120 may substantially surround the conduit 160. In some embodiments, the mounting member 130 may surround the conduit 160. In some embodiments, each of the first spring member 140 and the second spring member 150 at least partially surrounds the conduit 160. In the illustrated embodiment, the conduit 160 is substantially cylindrical in shape.

In some embodiments, the first abutment member 110, the second abutment member 120, the mounting member 130, the first spring member 140 and the second spring member 150 are integrally formed. In other words, the component 100 is formed as a single element by an additive manufacturing process, such as additive layer manufacturing (ALM). The overall size and shape of the component 100 may be varied as per application requirements. For example, the size and shape of the mounting member 130 may depend on a configuration of a part mounted on the mounting member 130. This may allow the component 100 to be used for vibrationally isolating parts of any size. The materials from which the component 100 is formed may be any material having properties suitable for load conditions, weight constraints and environment (e.g. temperature) in which the component 100 is to be used. For example, the material may include one or more of: steel and its alloys, aluminium and its alloys, titanium and its alloys, nickel and its alloys, copper and its alloys, polymers, metal coated polymers, and composite materials. In some embodiments, the component 100 can be made using different metal, polymer, elastomers, a combination of metals and polymers, or a combination of different metals. Therefore, the component 100 is suitable for isolating parts from vibrations and withstand a high temperature environment where traditional rubber components cannot be used, for example, a core casing of the gas turbine engine 10. In some other embodiments, the component 100 may also be used in relatively cold environments, such as a fan case. Moreover, the component 100 may be built from a model file. In other words, the component 100 can be adapted to suit applications of varying size, shape, weight and stiffness requirements. In some embodiments, the component 100 may be manufactured using ALM (for example, selective laser melting, electron beam melting, blown powder deposition or equivalent alternative ALM processes) to suit individual requirements based on derived stiffness characteristics from computer simulations. In some other embodiments, a wax model of the component 100 may be manufactured using ALM. Investment casting may be used to manufacture the component 100 from the wax model.

Figure 6 shows a cut-away view of the component 100 shown in Figures 4 and 5. In the illustrated embodiment of Figure 6, both the first abutment member 110 and the second abutment member 120 are fixedly connected to the conduit 160. Further, the conduit 160 extends from the first abutment member 110. In some other embodiments, the conduit 160 may extend from the second abutment member 120. In some other embodiments, the conduit 160 may extend beyond the first abutment member 110 relative to the longitudinal axis "LA". In some embodiments, the conduit 160 may be integrally formed with the first abutment member 110 and the second abutment member 120. In the illustrated embodiment, the first spring member 140, the second spring member 150, and the mounting member 130 are configured to move independently of the first abutment member 110, the second abutment member 120, and the conduit 160. Accordingly, the first spring member 140, the second spring member 150, and the mounting member 130 are not fixedly connected to the conduit 160. In the illustrated embodiment, the through-hole 170 extends through the first abutment member 110, the first spring member 140, the mounting member 130, the second spring member 150, the second abutment member 120. Further, the conduit 160 is connected to both the first abutment member 110 and the second abutment member 120.

When the component 100 is in the mounting arrangement, the component 100 is configured such that a shaft of a fastener passes through the through-hole 170, a head of the fastener abuts the first abutment member 110, a first part (not shown in Figure 6) to be mounted by the mounting arrangement is coupled to the mounting member 130, and a second part (not shown in Figure 6) to be coupled to the fastener abuts the second abutment member 120.

Figures 7 and 8 show a mounting arrangement 300 according to the present disclosure. The mounting arrangement 300 is configured to mount a first part 310 to a second part 320 of the gas turbine engine 10 (shown in Figure 1). The mounting arrangement 300 is configured to vibrationally isolate the first part 310 from the second pad 320 of the gas turbine engine 10.

In some embodiments, the first part 310 may be a mounting bracket. In some embodiments, the second part 320 may be a casing section of the core casing of the gas turbine engine 10. In some embodiments, the mounting arrangement 300 may further include a unit 340 mounted on the first part 310. In some embodiments, the unit 340 may be an Electronic Engine Controller (EEC). In some other embodiments, the unit 340 may be any module mounted on the core casing of the gas turbine engine 10. In some other embodiments, the mounting member 130 of the component 100 may be directly be coupled to the unit 340. In some embodiments, the first part 310 may be built using ALM. In some embodiments, the component 100 may be formed integrally with the first part 310. In some other embodiments, the mounting member 130 may be integral to the first part 310. Integrating the first pad 310 to the mounting member 130 may remove the need for mounting fasteners (e.g., bolts, rivets etc.) to attach the component 100 to the first part 310.

The mounting arrangement 300 includes the component 100. In the illustrated embodiment, the mounting arrangement 300 includes four components 100. In some embodiments, the mounting arrangement 300 may include more than four components 100. In some other embodiments, the mounting arrangement 300 may include less than four components 100. The first part 310 is coupled to the mounting member 130 and the second part 320 abuts the second abutment member 120. The mounting member 130 includes the plurality of mounting holes 132 for coupling the mounting member 130 to the first part 310. The mounting holes 132 may receive respective mounting fasteners (not shown) for coupling the mounting member 130 to the first part 310. The mounting fasteners may include, for example, nut and bolt assemblies, screws, rivets, and so forth. In some embodiments, one of the plurality of mounting holes 132 on the mounting member 130 may be offset (e.g., radially and/or circumferentially) with respect to the other mounting holes 132. An offset mounting hole (not shown) may ensure that the component 100 is fixed to the first part 310 in a correct orientation. The correct orientation may further ensure that the components 100 are facing a same direction. In some embodiments, orientations of the first and second spring members 140, 150 may affect stiffness of the first and second spring members 140, 150. The correct orientation may further ensure that the first and second spring members 140, 150 are located in a particular direction or orientation.

The mounting arrangement 300 further includes a fastener 330. The fastener 330 may be a bolt. The fastener 330 includes a shaft 332 extending through the through-hole 170 of the conduit 160 and received in the second part 320. The fastener 330 further includes a head 334 abutting the first abutment member 110. In some other embodiments, the fastener 330 may be, for example, a rivet, a nut and bolt assembly, a screw and so forth.

When the component 100 is in the mounting arrangement 300, the component 100 is configured such that the shaft 332 of the fastener 330 passes through the through-hole 170. The head 334 of the fastener 330 abuts the first abutment member 110. The first part 310 to be mounted by the mounting arrangement 300 is coupled to the mounting member 130 and the second part 320 to be coupled to the fastener 330 abuts the second abutment member 120. Contact between the second abutment member 120 and the second part 320 may enable transmission of a load from the fastener 330 to the second part 320 while leaving the mounting member 130 free to move along the central axis of the component 100. Specifically, the mounting member 130 may be free to move along the longitudinal axis "LA" shown in Figure 6.

The component 100 may provide damping of vibrations through the first and second parts 310, 320. The second abutment member 120 transmits the load of the fastener 330, received through the conduit 160, to the second part 320.

In the example shown in Figures 7 and 8, when the component 100 is in position in the mounting arrangement 300, vibrations imparted from the gas turbine engine 10 may cause the unit 340 mounted on the first part 310 to move along the central axis. The vibrations imparted from the gas turbine engine 10 may load the first and second spring members 140, 150 in opposite directions. Specifically, the vibrations imparted from the gas turbine engine 10 may alternately load the first spring member 140 in tension and the second spring member 150 in compression, and the first spring member 140 in compression and the second spring member 150 in tension. The loading of the first and second spring members 140, 150 in opposite directions may cause the unit 340 mounted on the first part 310 to remain substantially at its datum position. The movement of the first and second spring members 140, 150 in opposite directions may help in damping the vibrations imparted from the gas turbine engine 10. In other words, the compression and tension in the first and second spring members 140, 150 may lower the vibration output as compared to the vibration input imparted from the gas turbine engine 10 and thereby help to isolate the unit 340 mounted on the first part 310 from the vibrations imparted from the gas turbine engine 10. The component 100 may therefore prevent the vibrations imparted from the gas turbine engine 10 from entering into the unit 340 mounted on the first part 310. The component 100 may also substantially prevent the vibrations from being transmitted back to the first part 310. Thus, the component 100 may protect both the unit 340 and the first part 310 from fatigue effects of the vibrations. The component 100 may lower a transmissibility ratio defined by the ratio of the amplitude of vibration response to the amplitude of vibration input imparted from the gas turbine engine 10. The component 100 may provide vibration isolation in high temperature environments without requiring the unit 340 and/or the second part 320 to have larger size and weight. This may decrease an overall mass of the gas turbine engine 10.

Figure 9 shows a cut-away view of another example of a component 200 in accordance with another embodiment of the present disclosure. The component 200 is substantially similar to the component 100 of Figures 4 and 5 with each of features 210 to 270 corresponding to equivalent features 110 to 170 of the component 100. The component 200 defines a longitudinal axis "LA2" along its length. However, in this embodiment, the conduit 260 does not extend from the first abutment member 210. In other words, the conduit 260 is detached from the first abutment member 210. Instead, the conduit 260 extends only from the second abutment member 220. In other words, the conduit 260 is connected to the second abutment member 220.

In some embodiments, the conduit 260 may be integrally formed with the second abutment member 220. In the illustrated embodiment, the first abutment member 210, the first spring member 240, the second spring member 250, and the mounting member 230 are configured to move independently of the second abutment member 220 and the conduit 260. Accordingly, the first abutment member 210, the first spring member 240, the second spring member 250, and the mounting member 230 are not fixedly connected to the conduit 260. As shown in Figure 9, the conduit 260 further includes an end 262 that is not connected to the first abutment member 210. Specifically, the end 262 of the conduit 260 is a free end distal to the second abutment member 220. The end 262 is radially spaced from the second surface 214 of the first abutment member 210. In some embodiments, the end 262 of the conduit 260 may be axially spaced from the second surface 214 of the first abutment member 210.

Specifically, the end 262 may be spaced apart from the second surface 214 of the first abutment member 210 relative to the longitudinal axis "LA2". In some embodiments, the through-hole 270 may extend partially through the first abutment member 210 such that the end 262 of the conduit 260 is surrounded by the first abutment member 210. However, in some other embodiments, the through-hole 270 may only extend partially through the first spring member 240 such that the end 262 of the conduit 260 is surrounded by the first spring member 240. In some embodiments, a gap (not shown) may be present between the end 262 of the conduit 260 and the first abutment member 210. The gap may allow the first abutment member 210 to axially extend beyond the end 262 of the conduit 260. The gap may enable compression of the first and second spring members 240, 250 when a fastener (not shown) passing through the through-hole 270 is tightened. Specifically, the fastener may axially move the first abutment member 210 towards the second abutment member 220 resulting in the compression of the first and second spring members 240, 250.

Therefore, a length of the conduit 260 and hence the gap may be varied to control pre-compression in the first and second spring members 240, 250.

In this embodiment, when the component 200 is in the mounting arrangement, 5 the component 200 is configured such that a shaft of the fastener passes through the through-hole 270, a head of the fastener abuts the first abutment member 210, a first part (not shown in Figure 5) to be mounted by the mounting arrangement is coupled to the mounting member 230, and a second pad (not shown in Figure 5) to be coupled to the fastener abuts the second abutment 10 member 220.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

CLAIMS: 1. A component (100) for a mounting arrangement (300) of a gas turbine engine (10), the component (100) comprising: a first abutment member (110); a second abutment member (120) spaced apart from the first abutment member (110), the second abutment member (120) comprising a first surface (122) facing the first abutment member (110) and a second surface (124) opposite to the first surface (122); a mounting member (130) disposed between the first abutment member (110) and the second abutment member (120); a first spring member (140) disposed between the first abutment member (110) and the mounting member (130); a second spring member (150) disposed between the mounting member 15 (130) and the second abutment member (120); and a conduit (160) comprising a through-hole (170) extending at least through the first abutment member (110), the mounting member (130), the first spring member (140), and the second spring member (150), wherein the conduit (160) is connected to the second abutment member (120); and wherein, when the component (100) is in the mounting arrangement (300), the component (100) is configured such that a shaft (332) of a fastener (330) passes through the through-hole (170), a head (334) of the fastener (330) abuts the first abutment member (110), a first part (310) to be mounted by the mounting arrangement (300) is coupled to the mounting member (130) and a second part (320) to be coupled to the fastener (330) abuts the second abutment member (120).
2. The component (100) of any preceding claim, wherein the mounting member (130) comprises a plurality of mounting holes (132) extending 30 thereth rough.
3. The component (100) of any preceding claim, wherein the conduit (160) is substantially cylindrical in shape.
4. The component (100) of any preceding claim, wherein the first abutment member (110), the second abutment member (120) and/or the mounting member (130) are substantially disc-shaped.
5. The component (100) of any preceding claim, wherein the conduit (160) extends from the first abutment member (110).
6. The component (100) of any preceding claim, wherein the conduit (160) is integrally formed with the first abutment member (110) and the second abutment 10 member (120).
7. The component (100) of any preceding claim, wherein the first spring member (140) connects the first abutment member (110) to the mounting member (130).
8. The component (100) of any preceding claim, wherein the second spring member (150) connects the mounting member (130) to the second abutment member (120).
9. The component (100) of any preceding claim, wherein each of the first spring member (140) and the second spring member (150) at least partially surrounds the conduit (160).
10. The component (100) of any preceding claim, wherein the mounting 25 member (130) surrounds the conduit (160).
11. The component (100) of any preceding claim, wherein the first abutment member (110), the second abutment member (120), the mounting member (130), the first spring member (140) and the second spring member (150) are 30 integrally formed.
12. The component (100) of any preceding claim, wherein the component (100) is formed as one element by an additive manufacturing process.
13. A mounting arrangement (300) for vibrationally isolating a first part (310) from a second part (320) of the gas turbine engine (10), the mounting arrangement (300) comprising: the component (100) of any preceding claim, wherein the first part (310) is 5 coupled to the mounting member (130) and the second part (320) abuts the second abutment member (120); and a fastener (330) comprising: a shaft (332) extending through the through-hole (170) of the conduit (160) and received in the second part (320); and a head (334) abutting the first abutment member (110).
14. The mounting arrangement (300) of claim 13, further comprising a unit (340) mounted on the first part (310).
15. A gas turbine engine (10) for an aircraft, the gas turbine engine (10) comprising: the mounting arrangement (300) of claim 13 or 14.