Patent Description:
During testing of composite components for gas turbine engines the components are tested to determine how long the component continues to burn, e.g. produce a flame, after an external heat source has been removed from the component. The time taken for the flame to extinguish is referred to as the "burn-on time". It is desirable for the burn-on time of composite components to be reduced.

Japanese patent application <CIT> discloses a fastening structure for a duct in a gas turbine. The fastening structure comprises a plurality of layers that can be of different lengths and form a J shaped structure. The structure can be connected to the casing through the use of bolts and mechanical fasteners. United States patent application <CIT> discloses an upstream flange of a fibre reinforcement for a fan casing comprising a J shaped portion of overlapping fibres. European patent application <CIT> discloses an airfoil component in a gas turbine engine. The airfoil component is made up of a plurality of layers. United Kingdom patent application <CIT> discloses a fibre-reinforced composite casing for a gas turbine engine, which has a circumferentially extending flange or rib that defines a projection. A metal cap is attached around the flange or rib to allow ancillary components to be attached. United States patent <CIT> discloses a thermal tapered composite substrate for mounting electronic circuit boards. United States patent <CIT> discloses a composite panel having an edge shock protected edge. In order to protect the edge a border comprising a metallic cloth is connected to the structure of the panel by an adhesive agent, which is diffused through the cloth.

The present invention provides a method of reducing a burn-on time of a composite component for a gas turbine engine after the composite component has been heated, a composite component for a gas turbine engine, and an assembly for an electrical system for a gas turbine engine, as set out in the appended claims.

According to a first aspect, there is provided a method of reducing a burn-on time of a composite component for a gas turbine engine after the composite component has been heated as set out in claim <NUM>.

The edge portion comprises the edge face. The edge portion may further comprise a portion of the body adjacent, e.g. immediately adjacent, to the edge face that is shaped or otherwise treated in order to reduce burn-on time. Burn-on time is the time a flame continues to burn after an external source of heat has been removed.

The edge face or portion thereof may be formed at an angle of between <NUM> degrees and <NUM> degrees relative to the front face.

The edge portion may be shaped by layering the fibres of the body such that the edges of two or more of the layers of fibres are staggered in a direction parallel to the front face of the body prior to curing the composite component. Additionally or alternatively, the edge portion may be shaped by machining the edge face after the body of the composite component has been formed, e.g. cured.

Shaping the edge portion of the body may comprise bending or folding the body of the composite component out of a plane of the front face to form a bend portion extending along a length of the edge portion, e.g. along a length of the edge face. The bend portion may create a rim or overlap at the edge.

The bend portion may extend along the body parallel to the edge face. The body may be bent through an angle of approximately <NUM> degree at the bend portion. Alternatively, the body may be bent through an angle of less than <NUM> degrees at the bend portion. Alternatively again, the body may be bent through an angle of greater than <NUM> degree at the bend portion. For example, the body may be bent through an angle of approximately <NUM> degrees at the bend portion, e.g. such that the body is bent or folded back on itself at the bend portion.

An edge cap may extend at least partially over the front face and the rear face. The edge cap may be resilient and may resist separation of the layers of fibres forming the body by virtue of its resilience. The edge cap may be adhered to the body, e.g. at the front face, rear face and/or the edge face. The edge cap may comprise a layer of fibres and associated matrix material positioned over the layers of fibres forming the body. The edge cap may comprise an intumescent coating or paint, a ceramic material, such as fire cement or a ceramic adhesive, or a metallic material, such as titanium.

The method may further comprise installing one or more fasteners, such as rivets, nuts and bolts or any other fasteners, passing through the layered fibres of the body, such that the fasteners act to resist separation of the layers of fibres at the edge face. The fastener may comprise first and second opposing shoulders abutting the front face and/or the rear face of the body. The fasteners may be installed adjacent to the edge, e.g. in the edge portion. The fasteners may be spaced apart along a length of the edge face. The fasteners may be arranged in a line parallel to the edge face. The fasteners may pass through the edge cap, e.g. the portion of the edge cap extending over the front face and/or the rear face of the body. The method may further comprise providing a strip of resilient material, e.g. a metallic material, such as titanium, over the front and/or rear faces of the body. The strip of resilient material may extend at least partially along the length of the edge face. The fasteners may pass though the strip of resilient material. The strip of resilient material may be provided over the end cap or between the end cap and the body.

The fasteners may pass though the body at the bend portion, on one side of the bend portion, e.g. between the bend portion and the edge face or on an opposite side of the bend portion to the edge, or at both sides of the bend portion, e.g. if the body is bent or folded back on itself at the bend portion.

The profile of the edge face undulates in a direction parallel to a plane of the front face, e.g. in a direction perpendicular to the direction in which the edge face extends.

The composite component may be a support structure for an electrical component of the gas turbine engine, such as a printed circuit board and/or an electronic controller of the gas turbine engine.

According to a second aspect, there is provided a composite component for a gas turbine engine as set out in claim <NUM>.

According to a third aspect, there is provided an assembly for an electrical system for a gas turbine engine that comprises the composite component of the second aspect and an electrical component coupled to, e.g. mounted on, the body of the composite component. The electrical component may comprise a printed circuit board. Additionally or alternatively, the electrical component may comprise an electronic controller of the gas turbine engine. The electronic controller may be mounted on the printed circuit board.

As noted elsewhere herein, the present invention relates to a gas turbine engine.

Arrangements of the present invention may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox.

The gas turbine engine as described herein may have any suitable general architecture.

In any gas turbine engine as described herein, a combustor may be provided axially downstream of the fan and compressor(s).

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 <NUM>% span position, to a tip at a <NUM>% 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: <NUM>, <NUM>, <NUM><NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>,or <NUM>. 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: <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches) cm or <NUM> (around <NUM> 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 <NUM> rpm, for example less than <NUM> 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 <NUM> to <NUM> (for example <NUM> to <NUM>) may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> 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 <NUM> to <NUM> may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> 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 <NUM> on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip<NUM>, where dH is the enthalpy rise (for example the <NUM>-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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> (all units in this paragraph being Jkg-<NUM>K-<NUM>/(ms-<NUM>)<NUM>). 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 invention 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. 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 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 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>. 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: <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s, <NUM> Nkg-<NUM>s or <NUM> Nkg-<NUM>s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described 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 <NUM> (ambient pressure <NUM>. 3kPa, temperature <NUM>), 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. 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: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. 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 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 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 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 invention may apply to engines with or without a VAN.

The fan of a gas turbine as described herein may have any desired number of fan blades, for example <NUM>, <NUM>, <NUM>, or <NUM> fan blades.

As used herein, cruise conditions may mean cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of decent. Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example <NUM> to <NUM>, for example on the order of Mach <NUM>, on the order of Mach <NUM> or in the range of from <NUM> to <NUM>. 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 <NUM> or above Mach <NUM>.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions at an altitude that is in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM> (around <NUM> ft), for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> (around <NUM> ft) to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example on the order of <NUM>. 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 <NUM>; a pressure of <NUM> Pa; and a temperature of -<NUM>.

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 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 <NUM> or <NUM>) gas turbine engine may be mounted in order to provide propulsive thrust.

The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines <NUM>, <NUM> before being exhausted through the core exhaust nozzle <NUM> to provide some propulsive thrust.

It will be appreciated that the arrangement shown in <FIG> is by way of example only, and various alternatives are within the scope of the present invention.

Accordingly, the present invention 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.

Other gas turbine engines to which the present invention 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 <FIG> has a split flow nozzle <NUM>, <NUM> meaning that the flow through the bypass duct <NUM> has its own nozzle that is separate to and radially outside the core exhaust nozzle <NUM>. However, this is not limiting, and any aspect of the present invention may also apply to engines in which the flow through the bypass duct <NUM> and the flow through the core <NUM> 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 invention 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 <NUM> may not comprise a gearbox <NUM>.

One or more components of the gas turbine engine <NUM> may be manufactured from a composite material, such as a carbon fibre reinforced polymer material. With reference to <FIG>, a composite component <NUM> for a gas turbine engine <NUM> comprises a body <NUM> comprising a plurality of fibres <NUM>, e.g. carbon fibres, layered between a front face 110a of the body and a rear face 110b of the body.

In the arrangement shown in <FIG>, the fibres <NUM> are layered within the composite component to form three layers <NUM>, e.g. a first layer 114a, a second layer 114b and a third layer 114c. However, in other arrangements, the body <NUM> may comprise any other number of layers. The fibres <NUM> are embedded within a matrix material <NUM>, such a polymer resin, e.g. a thermoset or thermoplastic polymer resin.

The body <NUM> comprises an edge face 110c formed between the front and rear faces 110a, 110b. As depicted in <FIG>, the edge face 110c is at least partially formed by edges <NUM> of the layers of fibres (which may be covered by the matrix material <NUM>).

It is desirable for the composite components and structures of the gas turbine engine <NUM> to be fire resistant or fire proof. Burn-on time is the time taken from a component to stop burning, e.g. stop producing a flame, after an external heat source, which was sufficient to cause combustion of the component material, has been removed from the component. Burn-on time is one parameter that can be used to quantify how fire resistant a component is. It is desirable for the burn-on time of composite components of the gas turbine engine to be reduced or minimised. For example, it may be desirable for burn-on time of the components to be less than approximately <NUM> seconds or less than approximately <NUM> second.

When composite components, such as the composite component <NUM>, are exposed to an external heat source, the matrix material <NUM> can evaporate, sublime and/or thermally decompose to produce a vapour. The vapour may be contained between the layers <NUM> of the fibres <NUM> and may pass between the layers <NUM> to reach an edge of the body, where the vapour may be released, e.g. from the edge face 110c.

The vapour produced by heating the matrix material <NUM> is flammable, and hence, whilst an external heat source is applying a sufficient amount of heat to the component, the vapours being released from the composite component <NUM> may ignite to produce a flame.

When the external heat source is removed, the burning vapour may continue to heat the body <NUM> of the composite component <NUM> causing more vapour to be produced. Because the vapour passes between the layers of fibres and is released at the edge face, a concentration of the vapour at the edge face, e.g. a local concentration at one or more positions along the edge face, can be sufficient to sustain a flame, increasing the burn-on time. In some arrangements, the concentration of vapours at the edge face may produce an approximately stoichiometric mixture of vapours and oxygen for a combustion reaction of the vapours at the edge.

With reference to <FIG>, the layers <NUM> of fibres in the body <NUM> may become partially separated from one another or "lofted" due to the heat of combustion at the edge. Lofting of the layers <NUM> of the composite body <NUM> may allow more air between the layers <NUM>, which may encourage combustion of the vapours.

With reference to <FIG>, the composite component <NUM> may be treated according to a method <NUM> in order to reduce the burn-on time of the composite component. The method <NUM> comprises a first step <NUM>, in which a fire retarding treatment is applied to an edge portion of the of the body in order to control, e.g. reduce, a local concentration of vapours from the matrix material at the edge face of the body, such that the concentration of vapours is insufficient for combustion to be sustained when the external heat source is removed.

For example, the edge portion of the body may be shaped, as described below, in order to control the local concentration of vapours from the matrix material at the edge face 110c of the body.

With reference to <FIG>, a composite component <NUM> treated according to the method <NUM> will now be described. The composite component is similar to the composite component <NUM> described above and comprises a body <NUM> comprising a plurality of layers <NUM> of fibres <NUM> that are layered between a front face 110a and a rear face 110b of the body <NUM> within a matrix material <NUM>. Edges <NUM> of the layers <NUM> of fibres at least partially form an edge face 110c of the body <NUM>.

According to the invention, the composite component differs in that an edge portion <NUM> of the body <NUM> is shaped such that positions of the edges <NUM> of two or more of the layers <NUM> of fibres forming the body <NUM> are staggered in a direction D parallel to the front face 110a of the body. By staggering the layers <NUM> of the fibres <NUM> at the edge portion <NUM> in this way, at least a portion of the edge face 110c is formed at a non-perpendicular angle relative to the front face 110a.

As depicted, the edge portion <NUM> comprises the edge face 110c. Additionally, the edge portion <NUM> may comprise a portion of the body <NUM> adjacent, e.g. immediately adjacent, to the edge face 110c that is treated, e.g. shaped, in order to control the concentration of vapours. As depicted in <FIG>, an angle A of the edge face 110c relative to the front face 110a may be approximately <NUM> degrees. In other arrangements, the angle A may be between <NUM> degrees and <NUM> degrees.

The edge portion <NUM> may be shaped by machining the edge portion after the body <NUM> of the composite component <NUM> has been formed, e.g. after the body has been cured. In other words, the edge face 110c may be cut using a cutting tool, e.g. a machining tool, such as a milling tool, after the body <NUM> has been cured.

Alternatively, the edge portion <NUM> may be shaped by layering the fibres <NUM> of the body <NUM>, e.g. prior to curing of the body <NUM>, to form the edge face 110c into the desired shape. For example, the fibres <NUM> of the body <NUM> may be layered such that the edges <NUM> of two or more of the layers <NUM> of fibres are staggered in the direction D parallel to the front face 110a of the body at the edge face 110c.

Forming the edge face 110c at a non-perpendicular angle relative to the front face 110a increases the area over which the vapours from the matrix material <NUM> are released from the body <NUM>. As a result, local concentrations of the vapours at positions along the edge face 110c are reduced. In particular, the local concentrations of the vapours may be reduced to an extent that the concentration of vapour is insufficient to sustain a flame at the edge.

With reference to <FIG>, shaping the edge portion of the body <NUM> may comprise bending or folding the body <NUM> of the composite component out of a plane of the front face 110a to form a bend portion <NUM>. The bend portion may extend at least partially along the length of the edge face 110c.

As shown in <FIG>, the body <NUM> may be bent through an angle of approximately <NUM> degrees at the bend portion <NUM>. In other words, the body <NUM> may be turned down, e.g. to create a rim along the edge. Alternatively, as depicted in <FIG>, the body <NUM> may be bent through an angle of greater than <NUM> degrees, such as approximately <NUM> degrees at the bend portion <NUM>, e.g. such that the body is bent or folded back on itself at the bend portion <NUM>.

The presence of the bend portion <NUM> may act to restrict the passage of vapours between the layers <NUM> to the edge face 110c, e.g. by creating a tortuous path for the vapour passing towards the edge. The bend portion <NUM> may therefore reduce the amount of vapour passing between the layers to reach the edge face 110c and may thereby reduce the concentration of vapours at the edge face 110c.

In the arrangement shown in <FIG>, the edge face 110c of the body <NUM> is formed substantially perpendicularly to the front face 110a at the edge. However, in other arrangements, the edge face 110c of the body <NUM> may be formed at a non-perpendicular angle relative to the front face 110a, e.g. as described above with reference to <FIG>, in addition to the bend portion <NUM> being formed on the body <NUM>.

Returning to <FIG>, the method <NUM> may comprise a second step <NUM> in which an edge cap <NUM> (depicted in <FIG>) is provided at the edge portion <NUM>, around the edge face 110c of the body, e.g. covering at least a portion of the edge face.

As shown in <FIG>, the edge cap <NUM> may extend at least partially over the front and rear faces 110a, 110b of the body <NUM>, and between the front and rear faces around the edge face 110c.

The edge cap <NUM> may be made from a resilient material. For example, the edge cap may comprise a metallic material, such as titanium. Alternatively, the edge cap <NUM> may comprise one or more layers of fibres (and associated matrix material) positioned over the layers of fibres forming the body <NUM>. The edge cap <NUM> may resist lofting, e.g. separation, of layers <NUM> of the body <NUM> at the edge face 110c when the composite component is heated, e.g. by virtue of its resilience. As depicted in <FIG>, the edge cap <NUM> may be adhered to the body, e.g. at the front face, rear face and/or the edge face.

Additionally or alternatively to resisting lofting of the layers <NUM>, the edge cap <NUM> may be configured to restrict vapours that have passed between the layers <NUM> from being released from the body <NUM> at the edge face 110c. For example, the edge cap <NUM> may be configured to create a seal for vapours at the edge face 110c and/or create a tortuous path for vapours being released from the edge face 110c of the body <NUM>. In some arrangements, the edge cap <NUM> may comprise fire resistant material, e.g. an intumescent coating or paint, a ceramic material, such as fire cement or a ceramic adhesive.

In the arrangement shown in <FIG>, the edge face 110c of the body <NUM> is formed substantially perpendicularly to the front face 110a, e.g. where the front face meets the edge face. However, in other arrangements, the edge face 110c of the body <NUM> is formed at a non-perpendicular angle relative to the front face 110a, e.g. as described above with reference to <FIG>, in addition to the edge cap being provided. In such cases, the shape of the edge cap <NUM> may be configured to match the shape of the edge portion. For example, a part of the edge cap <NUM> extending across the edge face 110c may extend in a direction parallel with the edge face 110c. Alternatively, the edge cap <NUM> may extend across the edge face 110c in a direction that is not parallel with the edge face 110c. For example, the edge cap <NUM> may extend across the edge face 110c in a direction perpendicular to the front face 110a of the body <NUM>.

Returning to <FIG>, the method <NUM> may comprise a third step <NUM> in which one or more fasteners <NUM> (depicted in <FIG>) are installed in the edge portion of the body <NUM>.

As depicted in <FIG>, the fasteners <NUM> are arranged to pass through the layered fibres <NUM> at the edge portion <NUM> of the body <NUM>, such that the fasteners <NUM> act to resist separation of the layers <NUM> of fibres at the edge face 110c. The fasteners <NUM> may comprise any type of fastener that can be configured to clamp against the front and rear faces 110a, 110b of the body in order to resist separation of the layers <NUM>. For example, the fasteners may comprise rivets or nuts and bolts.

As depicted in <FIG>, the fasteners <NUM> may each comprise a first shoulder <NUM>, arranged to apply a clamping force against the front face 110a of the body <NUM>, and a second shoulder <NUM> arranged to apply a clamping force against the rear face 110b of the body. When the fasteners <NUM> comprise a nut and bolt, one of the first and second shoulders <NUM>, <NUM> may be formed by the nut and the other of the first and second shoulders <NUM>, <NUM> may be formed by the bolt.

The fasteners <NUM> may be spaced apart along at least a portion of the length of the edge face 110c, such that the fasteners <NUM> act to resist separation of the layers along the length or portion of the edge face.

As depicted in <FIG>, a first strip of resilient material <NUM> may be provided over the front face 110a of the body <NUM> and a second strip of resilient material <NUM> may be provided over the rear face 110b of the body <NUM>. The strips of resilient material may extend at least partially along the length of the edge face 110c. As depicted, the fasteners <NUM> may pass through the first and second strips <NUM>, <NUM> of resilient material. The strips of resilient material may spread the clamping force applied by the fasteners <NUM> along the length of the edge face 110c in order to resist lofting of the layers of fibres along the edge between the fastener locations. In other arrangements, the first and/or second strips of resilient material <NUM>, <NUM> may be omitted.

In the arrangement shown in <FIG>, the edge face 110c of the body <NUM> is formed substantially perpendicularly to the front face 110a at the edge. However, in other arrangements, the edge face 110c of the body <NUM> may be formed at a non-perpendicular angle relative to the front face, e.g. as described above with reference to <FIG>.

Furthermore, in the arrangement shown in <FIG>, the body <NUM> does not comprise a bend portion <NUM>, e.g. as depicted in <FIG>. However, in other arrangements, the body <NUM> may comprise a bend portion <NUM> and the fasteners <NUM> may be arranged to pass through the layers <NUM> of fibres at the bend portion <NUM>, between the bend portion and the edge face 110c or on an opposite side of the bend portion <NUM> to the edge face 110c. As depicted in <FIG>, when the body <NUM> portion bends through an angle of approximately <NUM> degrees at the bend portion <NUM>, the fasteners <NUM> may be configured to pass through the layers <NUM> or both sided of the bend portion <NUM>.

It will be appreciated that when the bend portion <NUM> bends through an angle of approximately <NUM> degrees, such that the body <NUM> is bent or folded back on itself, the first and second shoulders <NUM>, <NUM> of the fasteners may both act against the front face 110a or rear face 110b of the body <NUM>. Similarly, in arrangements in which the strips of resilient material <NUM>, <NUM> are provided, both of the strips may be provided over the front face 110a of the body or over the rear face 110b of the body, e.g. between the front face 110a and the shoulders <NUM>, <NUM> of the fasteners <NUM>.

With reference to <FIG>, in the present invention further requires the edge portion <NUM> to be shaped such that the edge face 110c has an undulating profile along the length of the edge face. Shaping the edge face 110c to have an undulating profile increases the length of the edge face 110c, and hence, the local concentration of vapours at a position along the edge face may be reduced.

Shaping the edge portion <NUM> such that the edge face 110c has an undulating profile and forming the edge face at a non-perpendicular angle relative to the front face 110a is performed in addition to any of the other treatments for controlling local concentration of vapours at the edge face that are described above. For example, the edge portion <NUM> is shaped such that the edge face 110c has an undulating profile and such that the edge face is formed at a non-perpendicular angle relative to the front face in addition to providing an edge cap across the edge face 110c, providing one or more fasteners through the layers <NUM> and/or providing one or more strips of resilient material over the front and/or rear faces 110a, 110b of the component body <NUM>.

With reference to <FIG>, the composite component depicted in <FIG> and described above may be manufactured using a method <NUM>. The method <NUM> comprises a first step <NUM>, in which a plurality of fibres are layered to form a body of the composite component. The fibres are layered between a front face and a rear face of the body and an edge face of the body is formed between the front face and the rear face. The edge face is at least partially formed by edges of the layers of fibres. The method <NUM> further comprises a second step <NUM>, in which the composite component is cured.

When performing the method <NUM>, the composite component may be manufactured such that the edge portion of the body is treated in any of the ways claimed in the claims.

The third step <NUM> may be at least partially performed prior to the second step <NUM>. For example, during manufacturing of the composite component, e.g. during the first step <NUM>, the plurality of fibres may be layered such that the positions of the edges of two or more of the layers of fibres are staggered in a direction parallel to the front face of the body.

Additionally or alternatively, the third step <NUM> may be performed at least partially after the second step <NUM>. For example, the composite component may be machined after the second step <NUM>, such that the two or more of the layers of fibres forming the edge face are staggered in the direction parallel to the front face of the body. In this way, the edge face, or a portion of the edge face, may be formed at a non-perpendicular angle relative to the front face. In a similar way, the plurality of fibres may be arranged during the first step <NUM>, and/or the composite component may be machined such that the edge face follows an undulating profile along its length.

In the third step <NUM>, the edge portion of the composite component may be treated in any of the ways described above with reference to <FIG> in order to reduce a burn-on time of the composite component. In some arrangements, the third step <NUM> may comprise performing the method <NUM> described above.

With reference to <FIG>, the composite component <NUM> may form part of an assembly <NUM> for an ancillary system of the gas turbine engine <NUM>, such as an electrical system. As depicted in <FIG>, the assembly <NUM> may comprise the composite component <NUM> and an ancillary component <NUM> of the gas turbine engine <NUM>. The assembly <NUM> may be arranged about the core <NUM> of the gas turbine engine <NUM>.

In the arrangement shown in <FIG>, the assembly <NUM> is part of an electrical system of the gas turbine engine, and the ancillary component is an electrical component. However in other arrangements, the assembly may be part of any other ancillary system of the gas turbine engine and the ancillary component may be any other ancillary component.

As depicted in <FIG>, the electrical component comprises a printed circuit board <NUM> mounted on the front or rear face of the composite component body <NUM>. A controller <NUM>, e.g. of the gas turbine engine <NUM>, is mounted on the printed circuit board <NUM>.

In the arrangement depicted in <FIG>, the edge portion of the composite component body <NUM> has been shaped such that the edge face 110c is at a non-perpendicular angle relative to the front face 110a of the body, the composite component body <NUM> being further shaped by forming an undulating profile along the length of the edge face 110c, e.g. as depicted in <FIG>.

Additionally, the edge portion of the composite component provided within the assembly <NUM> may be treated by providing an edge cap across the edge, providing one or more fasteners through the layers <NUM> and/or providing one or more strips of resilient material over the front and/or rear faces 110a, 110b of the component body <NUM>, in order to reduce the local concentration of vapours at the edge face 110c.

Claim 1:
A method of reducing a burn-on time of a composite component (<NUM>) for a gas turbine engine after the composite component has been heated, wherein the composite component comprises a body (<NUM>) comprising a plurality of fibres (<NUM>) layered between a front face (110a) of the body and a rear face (110b) of the body within a matrix material (<NUM>) to form layers of fibres, wherein the body comprises an edge face (110c) between the front face and the rear face at least partially formed by edges (<NUM>) of the layers of fibres, the method comprising:
shaping an edge portion (<NUM>) of the body in order to control a local concentration of flammable vapours from the matrix material at the edge face of the body, the flammable vapours having been produced during heating of the composite component and having passed between the layers (<NUM>) of fibre of the body to the edge face;
wherein the edge portion (<NUM>) is shaped such that positions of the edges (<NUM>) of two or more of the layers (<NUM>) of fibres are staggered in a direction (D) parallel to the front face (110a) of the body, such that at least a portion of the edge face (110c) is formed at an non-perpendicular angle relative to the front face and the edge portion (<NUM>) is also shaped such that the edge face (110c) has an undulating profile along the length of the edge face.