GB2606342A

GB2606342A - Fibre reinforced gas turbine engine component

Info

Publication number: GB2606342A
Application number: GB2105996.9A
Authority: GB
Inventors: Hillier Steven; Edmonds Ian; Razzell Anthony
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2021-04-27
Filing date: 2021-04-27
Publication date: 2022-11-09
Also published as: GB202105996D0

Abstract

Method of manufacturing a ceramic matrix composite component, the component having an aperture 90a extending through a surface. To produce the component, an aperture mould is inserted into a fibre preform, locally deflecting the fibres around the mould away from their previous alignment. After depositing a matrix material in the fibre preform, the aperture mould is removed from the fibre preform to form a corresponding aperture 90a in the matrix material. The aperture mould may comprise a tapered insertion portion. The fibres and matrix material may each be formed of silicon carbide. The preform may be held inside an external tool and/or or an internal tool where the aperture mould extends at least partially into the external and/or internal tool. The component may be an aerofoil component for a gas turbine engine, where the holes may provide passages for transport of coolant.

Description

FIBRE REINFORCED GAS TURBINE ENGINE COMPONENT

The present invention relates to a component of a gas turbine engine, and particularly to a component which contains one or more holes extending through a surface of the component. Holes may be formed in such components, for example, to form passages for the transport of coolant therethrough.

In order to illustrate the related art, it is suitable to consider gas turbine engine components such as an aerofoil component. The performance of the simple gas turbine engine cycle, whether measured in terms of efficiency or specific output is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbine at the highest possible temperature. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature will always produce more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of cooling mechanisms.

In modern engines, the high pressure (HP) turbine gas temperatures are now much hotter than the melting point of typical blade materials. Cooling of the blades is therefore required.

In some engine designs the intermediate pressure (IP) and low pressure (LP) turbines are also cooled. During its passage through the turbine the mean temperature of the gas stream decreases as power is extracted. Therefore the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the HP stage(s) through the IP and LP stages towards the exit nozzle.

Internal impingement, internal convection, and external films are the prime methods of cooling the gas-path aerofoils, for example aerofoils, platforms, shrouds, shroud segments and turbine nozzle guide vanes (NGVs). Air is conventionally used as a coolant and is flowed in and around the gas-path aerofoils.

Ceramic Matrix Composites (CMSs) are materials that comprise ceramic fibres embedded in a ceramic matrix. CMGs typically exhibit desirable mechanical, chemical and physical properties at high temperatures. For example, CMCs are typically more resistant to oxidation at high temperatures than are metals. CMGs are generally tougher than monolithic ceramics and exhibit damage tolerance. SiC/SiC CMCs are one example of a composite material that exhibits excellent high temperature mechanical, physical and chemical properties. Such materials are suitable for a number of high temperature applications, such as for use in producing hot sector components of gas turbine engines. SiC/SiC CMC engine components allow gas turbine engines to operate at much higher temperatures than engines having superalloy metal components.

CMCs may be produced by a variety of processes. One process for producing CMCs uses chemical vapour infiltration (CVO to deposit the matrix material on a network of fibres.

In consideration of the suitability of CMCs for forming high temperature components for gas turbine engines, the present invention aims to address drawbacks in the formation of holes in the components, in particular for the passage of coolant. For metallic components, suitable holes can be formed by machining without deleteriously affecting the mechanical properties of the component. However, for CMC components, typical machining processes would sever some of the reinforcing fibres in the CMC material, locally reducing mechanical properties. Furthermore, typical machining processes may induce cracks within the matrix, increasing the rate of environmental deterioration of the component.

According to a first aspect there is provided a method of manufacturing a ceramic matrix composite component having a first surface and an aperture extending through the first surface, the method comprising: providing a fibre preform comprising an array of ceramic fibres, at least a first group of ceramic fibres in the array of ceramic fibres extending aligned in a first direction; inserting an aperture mould into the fibre preform, the insertion of the aperture mould deflecting one or more fibres of the first group of ceramic fibres so that the fibres extend around the aperture mould, deviated locally from their previous alignment in the first direction; depositing a matrix material in the fibre preform; and removing the aperture mould by extracting the aperture mould from the fibre preform, thereby forming a corresponding aperture in the matrix material, the aperture extending through the first surface of the ceramic matrix composite component.

According to a second aspect there is provided a ceramic matrix composite component comprising an array of ceramic fibres embedded in a ceramic matrix, the ceramic matrix composite component having a first surface and an aperture formed to extend though the first surface, there being defined a proximal region of the component proximal to the aperture and a distal region of the component distal to the aperture, wherein there is provided at least a first group of ceramic fibres in the array of ceramic fibres, the first group of ceramic fibres extending aligned in a first direction in the distal region of the component, and wherein the first group of fibres extend continuously from the distal region through the proximal region of the component, and wherein in the proximal region of the component, one or more fibres of the first group of ceramic fibres are deviated from alignment with the first direction to extend around the aperture, thereby providing fibre continuity around the aperture.

According to a third aspect there is provided a gas turbine engine for an aircraft incorporating a component according to the second aspect.

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

The step to remove the aperture mould may occur after a step of at least partially degrading the aperture mould within the fibre preform. Thus, removal may occur after partial degradation of the aperture mould within the fibre preform.

Insertion of the aperture mould may deflect at least two fibres of the first group of ceramic fibres so that the two fibres are deflected in opposing directions to be further spaced apart and thereby extend around the aperture mould. In this way, deflection of the fibres to create or extend the aperture by inserting an aperture mould into the fibre preform before depositing a matrix material may improve durability and long term mechanical performance.

There may be an accommodation of the deflected fibre(s) by a corresponding thickness variation in the component. This increase in thickness may be presented at an interior surface of the component, or an exterior surface of the component, or a combination of these such that the increase in thickness is shared between the interior surface and exterior surface of the component.

One advantage provided by some embodiments is that the aperture may be formed at any desired angle with respect to an exterior or interior surface of the component. This is in contrast to many machining techniques.

Note that it is permitted to fill or block the aperture with a temporary fill in order to allow downstream processing (for example CV1/31/M1). This is intended to reduce or avoid unintended blockage of the aperture due to such downstream processing.

The aperture may lie in the path of the alignment of the first group of ceramic fibres in the distal region. Therefore, in the proximal region, at least one of the fibres of this first group is displaced from its notional position, this notional position being a position based on a notional continuation of the alignment from the distal region to the proximal region with the aperture notionally being absent.

Typically, the ceramic fibres in the preform are laid in plies of fibres. Accordingly, the first group of ceramic fibres may be formed in a ply of fibres, the deflection of the ceramic fibres being within the ply. Thus, considering the direction of insertion of the aperture mould, the deflection of the ceramic fibres may be in a direction normal to the direction of insertion of the aperture mould.

The fibre preform may include fibres aligned in different directions. Accordingly, the array of ceramic fibres may include a second group of ceramic fibres extending aligned in a second direction, wherein the second direction is not aligned with the first direction. The second direction may be perpendicular to the first direction. The second direction may be oblique to the first direction. Alternatively, the fibres may be unidirectional (UD), the fibre layup therefore being a UD fibre layup.

As for the first group of fibres, the insertion of the aperture mould may cause deflection of one or more fibres of the second group of ceramic fibres so that the fibres extend around the aperture mould, deviated locally from their previous alignment in the second direction.

In the context of the ceramic matrix composite component, the second group of fibres may extend continuously from a distal region through the proximal region of the component. In view of the different alignment of the second group of fibres, this may be a different distal region than considered with respect to the first group of fibres. In the proximal region of the component, one or more fibres of the second group of ceramic fibres may be deviated from alignment with the second direction to extend around the aperture, thereby providing fibre continuity around the aperture.

The ceramic matrix composite component may include at least a first interior space. The aperture may extend from the first surface to the first interior space. There may be provided a first array of apertures, formed using corresponding aperture moulds. The first array of apertures may extend from the first surface to the first interior space. In this way, the apertures may provide channels for the passage of a cooling fluid, in use. The cooling fluid may be conducted from the first interior space.

The ceramic matrix composite component may include at least a second interior space. There may be provided a second array of apertures, formed using corresponding aperture moulds. The second array of apertures may extend from the first surface to the second interior space.

The aperture mould may comprise an outer surface formed for contact with the fibres and/or the matrix material. In some embodiments, the outer surface of the aperture mould may comprise a material facilitating extraction of the aperture mould the matrix material. Thus, in some examples, the outer surface of the aperture mould may comprise a material facilitating physical extraction, to remove the aperture mould from the matrix material. The outer surface of the aperture mould may be relatively more susceptible to, for example, one or more of oxidation, sublimation, evaporation, combustion, or vaporising, melting, dissolving, chemical leaching, cooling, or shrinking, than the fibres and/or the matrix material, or the main body of the aperture mould. In some examples, at least a portion of the outer surface of the aperture mould may be degraded using a chemical degradation agent that selectively degrades the material of the outer surface of the aperture mould in preference to the composite material.

In some examples, the aperture mould may be a preform aperture mould. Additionally or alternatively, in some examples, the aperture mould may be a deposition aperture mould. In some embodiments, the preform aperture mould may be used to deflect the fibres in the fibre preform. Such a preform aperture mould may for example be one or more of a metallic, organic, ceramic, polymeric, or a wax-based material. For the process of depositing the matrix material, the preform aperture mould may be replaced with the deposition aperture mould. The deposition aperture mould may be formed, at least at its outer surface, of a material that facilitates removal from the matrix material and/or from the fibres. Suitable materials may be inert during the deposition process. One suitable material, for example, may be graphite. Thus, in some examples, where one or more pins are used as preform aperture moulds, fibres may be temporarily held in place with a binder or an adhesive, before the or each pin is removed and replaced with a deposition aperture mould. In some examples, the preform aperture mould may comprise a pointed or tapered end to facilitate insertion. In further examples, the deposition aperture mould may be relatively less pointed or tapered than the preform aperture mould.

In some examples, the aperture (and thus the preform aperture mould) may extend through a ply (or plys) of the fibre preform. However, in some embodiments, the aperture (and thus the preform aperture mould) may extend between adjacent plys of the fibre preform. For example, this latter configuration may apply to apertures formed at or close to a trailing edge of the component. However, such an arrangement may still result in deflection of fibres.

The outer surface of the aperture mould may be provided by a coating. This coating may be formed on a core aperture mould.

In some embodiments, the aperture mould may be physically and entirely removed from the fibre preform. In some embodiments, at least a portion of the aperture mould may be degraded by one or more of oxidation, burning, melting, vaporising, dissolving, chemical leaching, cooling, or shrinking before physical removal or machining of the aperture mould from the fibre preform. In such examples, the aperture mould may be relatively more susceptible to, for example, one or more of oxidation, sublimation, evaporation, combustion, or vaporising, melting, dissolving, chemical leaching, cooling, or shrinking, than the fibres and/or the matrix material, or the main body of the aperture mould.

Considering the structure of the aperture mould in more detail, the aperture mould may comprise an insertion end, an intermediate portion and a removal end. The intermediate portion may correspond to the aperture to be formed in the component. The insertion end may be tapered relative to the intermediate portion. This may facilitate deflection of the fibres during insertion, and reduce the chance of fibre breakage. The insertion end may terminate in a sharp edge or point. The insertion end may comprise a lubricious or anti-wear coating.

The removal end may include one or more features to facilitate removal of the aperture mould from the ceramic matrix composite component. Such features may provide an indication of the depth to which the aperture mould has been inserted into the fibre preform.

The fibre preform may be held inside an external tool for deposition and/or consolidation of the matrix material. The aperture mould may extend at least partially into the external tool in order to retain the aperture mould in position.

The fibre preform may be held on an internal tool for deposition and/or consolidation of the matrix material. The aperture mould may extend at least partially into the internal tool in order to retain the aperture mould in position.

The composite may be a continuous fibre reinforced ceramic matrix composite. The fibres may for example be formed of an oxide ceramic or a carbide ceramic. A suitable carbide ceramic may be SiC. The matrix may for example be formed of an oxide ceramic or a carbide ceramic. A suitable carbide ceramic may be SiC. Therefore in one embodiment the continuous fibre reinforced ceramic matrix composite may be formed of SiC fibres embedded in an SiC matrix. The reinforcing fibres may be contained in layered plies, the orientation of the fibres in the plies optionally alternating from ply to ply.

The matrix material may for example be deposited by chemical vapour infiltration (CV!). CVI may be carried out a temperature in the range of about 300 °C to about 2000°C. CV! may be carried out a pressure in the range of about 0.1 torr to about 10 atm.

In the fibre preform, the fibre volume may be in the range 15-50 vol %. This is intended to express the actual volume occupied by the fibre compared with the apparent total volume of the fibre preform. The fibre volume may be in the range 30-40 vol %.

The fibres may have one or more surface coatings. These may be applied before or after assembly of the fibres into the fibre preform. Suitable surface coatings may include coatings to affect the high temperature oxidation resistance of the fibres.

One suitable SiC fibre may be HI-NICALON TM TYPES manufactured by Nippon Carbon Co., Ltd. of Japan. This is a multi-filament SiC fibre. Typical filament diameter is in the range 10-pm. There may be about 500 filaments per tow.

The aperture may have a rounded cross sectional shape. This assists to avoid stress concentration. The aperture may for example have a circular cross sectional shape. In some examples, the diameter of the aperture may be between about 0.2 mm and about 10 mm. In further examples, the diameter of the aperture may be between about 0.3 mm and about 5 mm. In further examples, the diameter of the aperture may be between about 0.4 mm and about 3 mm. In further examples, the diameter of the aperture may be at least 0.4 mm. In yet further examples, the diameter of the aperture may be at most 3.0 mm. Where the aperture has a cross sectional shape other than circular, the maximum and/or minimum width of the aperture may correspond to the diameter ranges specified above. For example, where the aperture has a cross sectional shape of a slot, either of the diameter or width of the aperture may be at most 8.0 mm.

The component may for example be a turbine guide vane. The component may for example be a turbine rotor blade.

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).

It is to be understood here that the gearbox used to drive the fan is different from the accessory gearbox. The expression "gearbme' used in this specification is to be understood as the gearbox used to drive the fan (unless the context demands otherwise) and the expression "accessory gearbox" used in this specification is to be understood as that used to drive the accessories.

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-limitafive 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 Ufip. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/U1ip2, where dH is the enthalpy rise (for example the 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: 0.3, 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-if(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 TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). 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 On terms of time and/or distance) between top of climb and start of descent.

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.

Embodiments will now be described by way of example only, with reference to the accompanying figures, in which: Figure I 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 a schematic perspective partial cut-away view of an aerofoil component according to an embodiment; Figure 5 is a schematic cross sectional view of a method for manufacturing a component according to an embodiment; Figure 6 is a schematic plan view of the relationship between a machined hole and the reinforcing fibres in a CMC component, for reference; Figure 7 is a schematic plan view of the relationship between a hole and the reinforcing fibres in a CMC component according to an embodiment; Figure 8 is a schematic side view of an aperture mould for use in an embodiment; Figure 9 shows a schematic flow diagram of the process for providing the fibre preform; and, Figures 10, 11, 12 and 13 show schematically suitable processes for consolidation of the CMC.

Figure 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 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 invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the Figure 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of Figure 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in Figure 2.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

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

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in Figure 1 has a split flow nozzle 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.

In one embodiment, this disclosure provides an aerofoil component for the high pressure turbine 17 of the gas turbine engine described above. For example, the component may be a turbine guide vane or rotor blade.

Figure 4 shows a schematic perspective partial cut-away view of the aerofoil component 50 comprising a ceramic matrix composite material 52. Component 50 includes first interior space 54. A first array of apertures 56 extend from first surface 58 to the first interior space 54. Similarly, there is a second interior space 60 and a second array of apertures 62 extend from first surface 58 to the second interior space 60. Furthermore, there is a third interior space 64 and a third array of apertures 66 extend from first surface 58 to the third interior space 64. In some examples, the aerofoil is a nozzle guide vane (NGV). In some examples, the aerofoil is a turbine blade.

The apertures 56, 62, 66 are formed using corresponding arrays of aperture moulds during the manufacture of the component, as will now be described. The apertures provide channels for the passage of a cooling fluid, in use. The cooling fluid is conducted from the interior spaces 54, 60, 64 to form a cooling film over the surface of the component in use.

Figure 5 is a schematic cross sectional view of a method for manufacturing a component according to an embodiment. A fibre preform 70 (shown in blank outline in Figure 5) is held inside an external tool 72 formed of graphite for subsequent deposition and/or consolidation of the matrix material. The fibre preform 70 is held on an internal tool 74 formed of graphite also during this process. As shown in the drawing, the external tool has upper and lower halves to allow splitting of the tool in order to extract the component after this process.

As shown in Figure 5, aperture moulds 76, 78, 80 extend through the external tool 72 and through the fibre preform 70 and extend at least partially into the internal tool 74 in order to retain the aperture moulds in position with respect to the fibre preform 70 and the external tool 72 and internal tool 74.

In the example shown, at least a portion of the aperture moulds 76, 78, 80 are comprised of a material which is at least partially resistant to for example, one or more of oxidation, sublimation, evaporation, combustion, or vaporising, melting, dissolving, chemical leaching, cooling, or shrinking. In further examples, at least a portion of the aperture moulds 76, 78, 80 is relatively more resistant to for example, one or more of oxidation, sublimation, evaporation, combustion, or vaporising, melting, dissolving, chemical leaching, cooling, or shrinking than a further portion of the fibre preform 70 and/or the matrix material. Thus, the aperture moulds 76, 78, 80 retain structural integrity during and after each of the manufacturing processes carried out on the fibre preform 70. Thus, the aperture moulds 76, 78, 80 may be extracted from the aperture mould 76, 78, 80, 110 at the relevant stage in the manufacturing process.

In further examples, an outer coating of the aperture moulds 76, 78, 80 may be relatively more susceptible to, for example, one or more of oxidation, sublimation, evaporation, combustion, or vaporising, melting, dissolving, chemical leaching, cooling, or shrinking, than a further portion of the aperture moulds 76, 78, 80. Thus, by exposing the aperture moulds 76, 78, 80 to a further material which degrades the aperture moulds 76, 78, 80, at least a portion of the aperture moulds 76, 78, 80 may be preferentially degraded, relative to the fibre preform 70.

The composite to be produced by the process may be a continuous fibre reinforced ceramic matrix composite. The fibres may for example be formed of an oxide ceramic or a carbide ceramic. A suitable carbide ceramic is SiC. The matrix may for example be formed of an oxide ceramic or a carbide ceramic. A suitable carbide ceramic is SiC. Therefore in one embodiment the continuous fibre reinforced ceramic matrix composite is formed of SiC fibres embedded in an SiC matrix. The reinforcing fibres may be contained in layered plies, the orientation of the fibres in the plies optionally alternating from ply to ply.

In forming the fibre preform 70, a fibre network is produced. For example, the fibre network can be a near net shape preform of the component. This is referred to here as a fibre preform. The fibre volume in the fibre preform may range between about 15% and about 50%. More specifically, the fibre volume in the fibre preform may typically range between about 30% and about 40%.

The fibre may be continuous. In principle, it is possible for a continuously wound single fibre to be formed into the required fibre preform 70. In this case, the advantageous effects are seen as if the fibre preform was formed of an array of long individual fibres. Thus, references herein to "fibres" are intended also to refer to portions of a continuously wound single fibre, as appropriate.

The fibre preform 70 may be coated with one or more interface coatings. The interface coatings can be selected to perform a number of functions, such as resisting crack propagation, increasing toughness of the matrix, improving bonding between the matrix and the fibres or producing other desirable results. The fibres may be coated by CVI or other methods.

The main matrix material, or a precursor thereof, is then deposited on the fibres. The materials may be deposited by a CV! process in which gases are introduced into a furnace and are deposited on the fibre preform through chemical reaction and infiltration. The gases may be activated by thermal energy, ultraviolet light, microwaves or other means. Deposition typically occurs at a temperature between about 300 °C to about 2000 °C, depending on the process and materials selected. Furnace pressure typically ranges from about 0.1 torr to about 10 atm. The thickness of the matrix material can be adjusted to vary the initial load sharing conditions between the fibres and the matrix.

The CMC may undergo further processing after deposition of the matrix material. For example, the CMC may be further processed by polymer infiltration and pyrolysis (PIP), slurry infiltration, melt infiltration, further CVI, heat treating to obtain a desired material microstructure or combinations of the foregoing. In some cases it may be desirable to precondition a component by stressing it at elevated temperature to increase the matrix cracking stress. This may be particularly desirable when the matrix exhibits greater creep than the fibre.

One suitable SiC fibre is HlNlCALONTM TYPE S manufactured by Nippon Carbon Co., Ltd. of Japan. This is a multi-filament SIC fibre. Typical filament diameter is in the range 10-20 pm. There may be about 500 filaments per tow.

Additional explanations of different processes for forming CMC composites are now set out. CMC manufacture typically consists of four main steps. The first step is fibre forming. Individual fibres are gathered into groups called tows. A 'size' is applied to protect the fibre from damage, the size later being washed or burnt off. Tows are woven into cloth (such as a 5-harness satin) or 3D form. Cloth is laid into a tool and formed to the correct shape. 3D forms fit directly into a tool. A binder may be used to fix the relative position of the fibres. In the context of the present embodiments, a binder may or may not be used, but if a binder is used, it is still possible locally to deflect the fibres to accommodate the aperture tool.

The second step is consolidation. A number of techniques can be used in isolation or combination to form the matrix material around the fibres. The most commonly used are CVI (Chemical Vapour Infiltration), SI (Slurry Infiltration), MI (Melt Infiltration) and SMI (Slurry Melt Infiltration). Some process are slow or may need repeating multiple times to achieve the required thickness of deposition or component density.

The third step is machining. The final component is shaped usually with minimal machining to prevent cutting or exposing fibres to the environment. An additional processing step may be after machining to 'seal' the cut end of the fibres.

The fourth step is coating. TBC (Thermal Barrier Coating) or EBC (Environmental Barrier Coatings) may be applied to some or all of the component to prevent degradation in use.

Figure 9 shows a schematic flow diagram of the process for providing the fibre preform.

Figures 10, 11, 12 and 13 show schematically suitable processes for consolidation of the CMC. These processes may be combined as appropriate. Following the consolidation of the CMC, the aperture mould is removed.

Following the consolidation, the component may be finish machined in order to produce the finished component.

Considering the interaction between the fibres and the aperture mould in more detail, it is useful to review first Figure 6 which shows a schematic plan view of the relationship between a machined hole and the reinforcing fibres in a CMC component, for reference.

Machined hole may be formed for example by drilling, spark cutting or the like. However, because the hole is formed only after deposition of the matrix material 92, the effect of forming the hole 90 is to cut fibres 94, 96, 98, 100 (in this schematic example).

Figure 7 is a schematic plan view of the relationship between a hole and the reinforcing fibres in a CMC component according to an embodiment. Before deposition of the matrix material, the fibre preform 70 is treated as follows. The preform 70 comprises an array of ceramic fibres. In this embodiment there is a first group of ceramic fibres 94a, 96a in the array of ceramic fibres extending aligned in a first direction. There is also a second group of ceramic fibres 98a, 100a in the array of ceramic fibres extending aligned in a second direction. An aperture mould 76, 78, 80 is inserted into the fibre preform 70, the insertion of the aperture mould 76, 78, 80 deflecting ceramic fibres 94a, 96a and ceramic fibres 98a, 100a so that the fibres extend around the aperture mould 76, 78, 80, deviated locally from their previous alignment. Matrix material 92a is then deposited into the fibre preform 70 and the aperture mould 76, 78, 80 is extracted from the fibre preform 70 and the matrix material 92a. Thereby, a corresponding aperture 90a is formed in the matrix material, and therefore in ceramic matrix composite component.

Considering Figure 7 in more detail, it can be seen that the schematic drawing shows the fibre arrangement close to the aperture 90a. Therefore the region shown containing the fibres in Figure 7 can be designated the proximal region. Regions 102a, 104a can be designated distal regions. In the distal regions, ceramic fibres 94a, 96a and ceramic fibres 98a, 100a retain their respective alignment. However, in the proximal region, the fibres are deviated from the alignment they have in the distal regions to extend around the aperture, thereby providing fibre continuity around the aperture 90a.

As shown in Figure 7, insertion of the aperture mould deflects at least two fibres of the first and second groups ceramic fibres so that the two fibres are deflected in opposing directions to be further spaced apart and thereby extend around the aperture mould.

Figure 8 is a schematic side view of an aperture mould 110 for use in an embodiment, or any such previously described embodiment.

The aperture mould 110 comprises an outer surface 112 formed for contact with the fibres and/or the matrix material. In some embodiments, the outer surface of the aperture mould presents a material facilitating removal of the aperture mould from the matrix material.

In some embodiments, a preform aperture mould may be used to deflect the fibres in the fibre preform. Such a preform aperture mould may for example be metallic. For the process of depositing the matrix material, the preform aperture mould may be replaced with a deposition aperture mould. At least a portion of the deposition aperture mould may be formed, at least at its outer surface, of a material that facilitates removal from within the matrix material and/or from the fibres. Suitable materials are inert during the deposition process. One suitable material for example is graphite.

The outer surface of the aperture mould may be provided by a coating (not shown). This coating may be formed on a core aperture mould (not shown).

In the embodiments discussed above, the aperture mould is typically removed from the component by physical extraction of the intact aperture mould. In other embodiments, at least a portion of the aperture mould may be removed by degradation. For example, at least a portion of the aperture mould may be removed by melting, sublimation, evaporation, combustion, etc. Alternatively, the aperture mould may be removed using a chemical degradation agent that selectively degrades the material of the aperture mould in preference to the composite material. Furthermore, following at least partial degradation of at least a portion of the aperture mould, the portion of the aperture mould may be removed from the component by physical extraction.

Considering the structure of the aperture mould in more detail, the aperture mould has an insertion end 114, an intermediate portion 116 and a removal end 118. The intermediate portion 116 corresponds to the aperture 90a to be formed in the component, and has a corresponding diameter D, after accounting as necessary for shrinkage due to subsequent processing of the component after removal of the aperture mould. The insertion end 114 is tapered relative to the intermediate portion 116. This facilitates deflection of the fibres during insertion, and reduces the chance of fibre breakage. The insertion end terminates in a sharp edge or point 120. The insertion end transitions to the intermediate portion 116 via a transitional region 122. The removal end may include one or more features (not shown) to facilitate removal of the aperture mould 110 from the ceramic matrix composite component. Such features may provide an indication of the depth to which the aperture mould has been inserted into the fibre preform.

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

CLAIMS1. A method of manufacturing a ceramic matrix composite component (50) having a first surface (58) and an aperture (56, 62, 66) extending through the first surface, the method 5 comprising: providing a fibre preform (70) comprising an array of ceramic fibres, at least a first group of ceramic fibres in the array of ceramic fibres extending aligned in a first direction; inserting an aperture mould (76, 78, 80, 110) into the fibre preform, the insertion of the aperture mould deflecting one or more fibres of the first group of ceramic fibres so that the fibres extend around the aperture mould, deviated locally from their previous alignment in the first direction; depositing a matrix material in the fibre preform; and removing the aperture mould (76, 78, 80, 110) by extracting the aperture mould (76, 78, 80, 110) from the fibre preform (70), thereby forming a corresponding aperture (90a) in the matrix material, the aperture extending through the first surface of the ceramic matrix composite component.
2. A method according to claim 1, wherein insertion of the aperture mould (76, 78, 80, 110) deflects at least two fibres (94a, 96a) of the first group of ceramic fibres so that the two fibres are deflected in opposing directions to be further spaced apart and thereby extend around the aperture mould.
3. A method according to claim 1 or claim 2 wherein the ceramic fibres in the preform are laid in plies of fibres, and the deflection of the ceramic fibres is within the ply. 25
4. A method according to any one of claims 1 to 3 wherein the array of ceramic fibres includes a second group of ceramic fibres extending aligned in a second direction, wherein the second direction is not aligned with the first direction, wherein the insertion of the aperture mould (76, 78, 80, 110) causes deflection of one or more fibres of the second group of ceramic fibres so that the fibres extend around the aperture mould, deviated locally from their previous alignment in the second direction.
5. A method according to any one of claims 1 to 4 wherein the aperture mould (76, 78, 80, 110) has an outer surface formed for contact with the fibres and/or the matrix material, the outer surface of the aperture mould presenting a material facilitating extraction of the aperture mould from the matrix material.
6. A method according to any one of claims 1 to 5 wherein the aperture mould (76, 78, 80, 110) comprises an insertion end (114), an intermediate portion (116) and a removal end (118), the intermediate portion (116) corresponding to the aperture (90a) to be formed in the component, the insertion end (114) being tapered relative to the intermediate portion (116).
7. A method according to any one of claims 1 to 6 wherein the fibre preform (70) is held inside an external tool (72) for deposition and/or consolidation of the matrix material, the aperture mould (76, 78, 80, 110) extending at least partially into the external tool (72) in order to retain the aperture mould in position.
8. A method according to any one of claims 1 to 7 wherein the fibre preform (70) is held on an internal tool (74) for deposition and/or consolidation of the matrix material, the aperture mould (76, 78, 80, 110) extending at least partially into the internal tool (74) in order to retain the aperture mould in position.
9. A method according to any one of claims 1 to 8 wherein the aperture mould is physically and entirely removed from the fibre preform.
10. A ceramic matrix composite component (50) comprising an array of ceramic fibres embedded in a ceramic matrix, the ceramic matrix composite component having a first surface (58) and an aperture (90a) formed to extend though the first surface, there being defined a proximal region of the component proximal to the aperture and a distal region (102a) of the component distal to the aperture, wherein there is provided at least a first group of ceramic fibres (94a, 96a) in the array of ceramic fibres, the first group of ceramic fibres extending aligned in a first direction in the distal region (102a) of the component, and wherein the first group of fibres extend continuously from the distal region (102a) through the proximal region of the component, and wherein in the proximal region of the component, one or more fibres of the first group of ceramic fibres (94a, 96a) are deviated from alignment with the first direction to extend around the aperture (90a), thereby providing fibre continuity around the aperture.
11. A ceramic matrix composite component according to claim 10 wherein the aperture (90a) lies in the path of the alignment of the first group of ceramic fibres in the distal region (102a) so that, in the proximal region, at least one of the fibres of the first group is displaced from its notional position, the notional position being a position based on a notional continuation of the alignment from the distal region to the proximal region with the aperture notionally being absent.
12. A ceramic matrix composite component according to claim 10 or claim 11 wherein a second group of fibres (98a, 100a) extend continuously from a distal region (104a) through the proximal region of the component so that, in the proximal region of the component, at least one of the fibres of the second group of ceramic fibres is deviated from alignment with the second direction to extend around the aperture (90a), thereby providing fibre continuity around the aperture.
13. A ceramic matrix composite component according to any one of claims 10 to 12 wherein the ceramic matrix composite component includes at least a first interior space (54), the aperture extending from the first surface to the first interior space.
14. A ceramic matrix composite component according to claim 13 wherein there is provided a first array of apertures (56), the first array of apertures extending from the first surface (58) to the first interior space (54).
15. A ceramic matrix composite component according to any one of claims 10 to 14 wherein the composite is a continuous fibre reinforced ceramic matrix composite.
16. A ceramic matrix composite component according to any one of claims 10 to 15 wherein the fibres are formed of SiC and the matrix is formed of SiC.
17. A ceramic matrix composite component according to any one of claims 10 to 16 wherein component is an aerofoil component for a gas turbine engine.
18. A gas turbine engine (10) for an aircraft incorporating a ceramic matrix composite component according to any one of claims 10 to 17.
19. A gas turbine engine (10) according to claim 18 further comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades and 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.
20. A gas turbine engine according to claim 19, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the core shaft is a first core shaft (26); the engine core further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.