Arrangement of components

Abstract

An arrangement of components the arrangement comprising an outer component 21 and an inner component 30 arranged coaxially about a central axis wherein, the arrangement comprises a plurality of supports 42, each support being elongate and having first and second ends attached to a surface of the outer component and having a location 40 between the first and second ends where the support is attached to a surface of the inner component. The inner and outer component may be cylindrical. Preferably, each end 40 of the first or second ends of the support is attached to the surface of the outer component by multiple bolts and the support is attached to the inner component by a single bolt. There may be 3, 4, 5 or 6 supports arranged symmetrically about the central axis. A gas turbine engine, where the outer component is an outer casing of the gas turbine engine, and the inner component is an engine core casing is also claimed. The gas turbine may have a bypass ratio in the range 0.1 to 2.5 and the gas turbine engine may comprise a fan having a plurality of fan blades each with a fan tip, wherein the radial distance between the central axis and the fan tip is in the range 90mm to 225mm.

Description

SECONDARY TITLE Supporting apparatus for concentric casings within a gas turbine. CROSS REFERENCE TO RELATED APPLICATIONS This represents the first application directed towards the subject-matter. FIELD The present disclosure relates to a gas turbine engine for air vehicles, and more specifically to a gas turbine engine with specified relative component dimensions. BACKGROUND Gas turbine engines have been developed to provide power for many different air applications including hobby, defence and civil. Each industry has their own requirements and although the present disclosure has been written with a primary focus on civil it will be appreciated that an engine with the specified relative component dimensions may also be used for appropriate products in the defence industry. Whilst some individuals may naively suggest that conventional gas turbines could be scaled to a small size the skilled person would appreciate that simply scaling components of a known engine type may not provide a corresponding scaling of power, thrust and / or efficiency. Particularly for smaller engines the size required for installation of a scaled engine may provide insufficient thrust, power or efficiency and concomitantly a scaled engine providing sufficient thrust, power or efficiency may be inappropriately large for installation on an airframe. Reconsideration of engine parameters may therefore be appropriate. Small hobby aircraft sometimes use gas turbines engines as their power source. The scale of these may be too small to be beneficial for commercial use which may require transport of high value cargo items, such as people and / or cargo many times more valuable than the cost of the engine. Commercial use engines may face a more stringent regulatory approval than model aircraft. Accordingly model aircraft may be simplified below a level that would be approved in the commercial market and the skilled person would appreciate that an engine produced for entertainment purposes may not provide sufficient reliability, etc. to meet the requirements set by the FAA, the CAA, or EASA. Additionally, commercial aircraft may have to operate for longer periods than hobby aircraft without refuelling. Thrust levels must be maintained whilst carrying the requisite combined weight of fuel and load. Hobby engines in their simplicity may lack the fuel efficiency and / or thrust levels required to deliver cargo over the required range for an amount of carried fuel. The skilled person would appreciate that, if the overall size of a gas turbine engine is decreased, one problem that may need to be addressed is howto ensure a sufficient thrust level is maintained. This is of interest for air vehicles where a finite amount of fuel is available and particular thrust and fuel efficiency is required to maintain the vehicle in flight for a given operational period. It would be appreciated that this may not be a factor for ground-based gas turbine engines such as those providing electrical generative power, or those where fuel is easily resupplied, or the fuel-tank size is not a significant limiting factor. The skilled person would also appreciate that greater thrust levels may be achieved in a smaller size engine that is arranged as a turbojet rather than a turbofan. However, the skilled person would also understand that turbojets are noisier than turbofans and have a significantly worse Specific Fuel Consumption (SFC) and may require significantly higher volumes of fuel. The noise is likely to be unacceptable in civil operations and may require extra fuel to be carried at the expense of passengers and cargo. Manufacturing tolerances of components and systems at smaller scales become increasingly important. Axial compressors become highly sensitive at very small scale with, for example, surface finish, tip clearance, stage spacing, etc. having to be so precise that quality control in manufacturing would reject an increasing number of parts resulting in uneconomic manufacture. Similarly, in flight damage, wear, dirt, or other minor issues would take the component or system away from acceptable tolerance leading to early replacement and additional cost that is likely to be unacceptable to economic operation. Reconsideration of engine parameters or component design may therefore be appropriate The skilled person would appreciate that simply scaling components of a known engine type may not provide a corresponding scaling of the operational window of the engine. For example, insufficient air and fuel mixing volumes and / or combustion volumes may induce problems with flame stability, relight and stable thrust. Whilst an engine core may be cantilevered from an upstream component the core may require downstream support to limit deflections which may arise from manoeuvre, vibration or thermal loads and which may affect rotor-dynamics or the clearance of rotor blade tips within the core. In an example a set of struts, or supports, may be used to span across a bypass duct, which is a duct between provided between the engine core and an outer casing. In larger engines these struts have ball-joints where they meet the casings to allow the struts to articulate without bending. This is enabled through the bypass duct having a significant radial annulus that permits the ball joints to be relatively small and compact compared to the strut lengths and thus they can be fairly easily accommodated in recesses in the engine cases. Such a construction is also considered appropriate to resist fatigue and stress damage within the struts that may be caused by repeated cycle of engine use and limit the necessity of a frequent replacement of components and / or catastrophic failure of the engine. At the scale of a very small turbofan the height of the bypass annulus makes a similar design impractical as the ball-joints, fasteners, etc. do not scale down in proportion and may add a significant bypass flow blockage. The relative size of the struts to the joints is also affected as the struts must be shortened to accommodate the larger joints and this makes them less able to cope with the movement between the core and the casing. As used herein, a range "from value X to value Y" or "between value X and value Y", or the likes, denotes an inclusive range; including the bounding values of X and Y. As used herein, the term "axial plane" denotes a plane extending along the length of an engine, parallel to and containing an axial centreline of the engine, and the term "radial plane" denotes a plane extending perpendicular to the axial centreline of the engine, so including all radial lines at the axial position of the radial plane. Axial planes may also be referred to as longitudinal planes, as they extend along the length of the engine. A radial distance or an axial distance is therefore a distance in a radial or axial plane, respectively. As used herein, the terms upstream and downstream denote relative locations along the axial centreline of the engine in the direction of the main intended use airflow through the engine. Axially forward and axially rearward denote relative locations along the axial centreline of the engine in the direction of the main intended use airflowthrough the engine. The skilled person will appreciate that equivalent units should be selected for the fan face area, turbine diameter and core length-e.g. if the area is given in m2 , the lengths should both be provided in metres. If the volume is given in m3, the lengths should be provided in metres. SUMMARY According to an aspect there is provided a gas turbine engine comprising an outer component and an inner component arranged coaxially about a central axis wherein, the arrangement comprises a plurality of supports, each support being elongate and having first and second ends attached to a surface of the outer component and a point between the first and second ends where the support is attached to a surface of the inner component; and wherein the outer component is an outer casing of the gas turbine engine and the inner component is an engine core casing. The gas turbine may have a bypass ratio in the range 0.1 to 2.5. The gas turbine may have a bypass ratio in the range 2.0 to 3.5. The gas turbine engine may comprise a fan having a plurality of fan blades each with a fan tip, wherein the radial distance between the central axis and the fan tip is in the range 90mm to 225mm. According to an aspect there is provided an arrangement of components the arrangement comprising an outer component and an inner component arranged coaxially about a central axis wherein, the arrangement comprises a plurality of supports, each support being elongate and having first and second ends attached to a surface of the outer component and a point between the first and second ends where the support is attached to a surface of the inner component. The supports may be arranged symmetrically about the central axis. The outer component may be cylindrical. The inner component may be cylindrical. Each of the first or second ends of the support may attached to the surface of the outer component by multiple bolts or rivets or by a weld, adhesive or braze joint. The support may be attached to the inner component by a single bolt. There may be 3, 4 or 6 supports. In the gas turbine the fan tip radius may be measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge. This may equivalently be described as the fan tip radius being defined as the radial distance between a centreline of the engine and an outermost tip of each fan blade at its leading edge. The fan face area may be equal to it multiplied by the square of the fan tip radius. The fan tip radius, measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge, may be in the range from 90mm to 225mm, for example in the range from 110mm to 225mm, or alternatively in the range from 150mm to 200mm. Within these ranges the fan tip radius may be greater than any of: 90mm, 95mm, 100mm, 105mm, 110mm, 115mm, 120mm, 125mm, 130mm, 135mm, 140mm, 145mm, 150mm, 155mm, 160mm, 165mm, 170mm, 175mm, or 180mm, or 185mm, or 190mm, or 195mm, or 200mm, or 205mm, or 210mm, or 215mm, or 220mm. The fan tip radius may be around any of: 90mm, 95mm, 100mm, 105mm, 110mm, 115mm, 120mm, 125mm, 130mm, 135mm, 140mm, 145mm, 150mm, 155mm, 160mm, 165mm, 170mm, 175mm, or 180mm, or 185mm, or 190mm, or 195mm, or 200mm, or 205mm, or 210mm, or 215mm, 220mm, or 225mm. Aircraft gas turbine engines provide both propulsive and electric power for aircraft. Consequently, such engines may comprise electric machines such as generators for providing electric power. Further electric machines in the form of motors configured to drive fluid pumps such as oil and fuel pumps may also be provided. The function of the generators and motors can be combined, such that, for instance, a generator may operate as a starter motor during starting. As electrical power demands increase, and electrical motors are increasingly used to replace hydraulic and pneumatic motors in aircraft, the size of electric machines on aircraft is steadily increasing relative to the size of the gas turbine engines. Consequently, packaging of such machines within the limited space available is becoming increasingly difficult. The person of skill in the art would appreciate that by providing a gas turbine with the above dimensions additional space may be created within the fan hub that may be used to house an electric generator or other components. The fan may be driven by a turbine that may have one or more stages. A shaft may connect the turbine to the fan. The length of the engine, defined as the distance between the leading edge of the first fan blade and the final turbine stage trailing edge, may be in the range: 420mm to 1020mm, optionally in the range 420 to 780mm, and further optionally in the range 740 to 1020mm. Within these ranges the engine length may be above 420mm, or 440mm, or 460mm or 480mm, or 500mm or 520mm or 540mm or 560m or 580mm, or 600mm, or 620mm or 640mm or 660mm, or 680mm, or 700mm, or 720mm, or 740mm, or 760mm, or 780mm, or 800mm, or 820mm, or 840mm or 860mm, or 880mm, or 900mm, or 920mm or 940mm or 960mm, or 980mm, or 1000mm. For a gas turbine with a fan tip ratio of 110mm to 225mm the length of the engine may be between 420mm and 1020mm, and optionally between 510 and 1020mm. Optionally, for example for an engine with a fan tip radius in the range from 150mm to 180mm the engine length may be in the range from 730mm to 790mm. The ratio of fan radius to engine length may be in the range 0.18 to 0.22, and further optionally in the range 0.19 to 0.21. The gas turbine may have an axial compressor supplied with air from the upstream fan. The compressor may comprise one or more compressor stages. Each stage of the compressor 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. There may be 3 or 4 or 5 or more compressor stages. The number of compressor stages divided by the fan tip radius (in mm) is in the range 0.013 to 0.055 In use air may be presented to the first (i.e. upstream) compressor stage in an axial direction and air may leave the first stage in an axial direction. The air may be presented to the first (i.e. upstream) compressor stage in an axial direction and air may leave the first stage in a radial direction. The one or more compressor stages may comprise a lowest pressure stage and may further comprise one or more compressor stages of increasing pressure to a highest pressure compressor stage. The lowest pressure compressor stage may be located furthest upstream along the gas flow path. The further higher pressure stages may be spaced axially along the gas flow path through the compressor in a downstream (rearward) direction. Each row of rotor blades provided in the compressor may form an annular array of rotor blades carried by a respective rotor hub (or rotor disc or drum). Each of the rotor blades may be coupled to the hub via a root received in a corresponding slot in a peripheral edge of the hub. Each of the rotor blades may be integrally coupled to the hub or integrally formed with the hub. Where the rotor blade is integrally coupled to the hub this may be achieved through friction welding of the blade to the hub and an optional step of machining to bring the blade and hub to a finished form. Where each rotor blade is integrally formed with the hub the blade and hub may be formed as a blank that is subsequently machined to provide the final form. Multiple stages may be welded together to form an integral drum, forward and rearward shafts may be welded to the integral drum to form a compressor assembly module. Each rotor blade may be defined as having a radial span extending from the root (or hub) at a radially inner gas-washed location, or 0% span position, to an outer most radial tip at a 100% span position. 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 rotor blade. The radial span of each rotor blade refers to the gas-washed portion of the rotor blade, i.e. the portion radially outside any platform at which it is coupled to the hub. The compressor tip radius may be measured between a centreline of the engine and an outermost tip of a compressor blade at its leading edge. The compressor tip radius may be the same for each compressor stage. The compressor tip radius may be different for one or more compressor stage. The compressor tip radius, measured between a centreline of the engine and an outermost tip of a compressor blade at its leading edge, may be in the range from 45mm to 202mm, for example in the range from 55mm to 202mm, or alternatively in the range from 55mm to 115mm, or alternatively in the range from 110mm to 202mm. Within these ranges a compressor tip radius may be greater than any of: 45mm, 55mm, 65mm, 75mm, 85mm, 95mm, 105mm, 115mm, 125mm, 135mm, 145mm, 155mm, 165mm, 175mm, 185mm, 195mm, or 200mm. The compressor tip radius may be around any of: 45mm, 55mm, 65mm, 75mm, 85mm, 95mm, 105mm, 115mm, 125mm, 135mm, 145mm, 155mm, 165mm, 175mm, 185mm, 195mm, or 200mm. The compressor may be connected to a turbine via an interconnecting shaft. The core length may be defined as the axial distance between a forward region of the compressor and a rearward region of the turbine that drives the turbine via the interconnecting shaft. The core length may be measured along a centreline of the engine from a mean radius point of the first stage of the compressor blade leading edge to a mean radius point of the lowest pressure turbine rotor stage blade trailing edge of the turbine. The core length may be in the range from 220mm to 560mm, and optionally 220mm to 400mm. Optionally, for example for an engine with a fan tip radius in the range from 110mm to 225mm the core length may be in the range from 270mm to 560mm. Optionally, for example for an engine with a fan tip radius in the range from 150mm to 180mm the core length may be in the range from 380mm to 440mm. Where the compressor is a multistage compressor the number of stages divided by the fan tip radius in mm may be in the range from 0.013 to 0.055, and optionally in the range 0.013 to 0.028, and optionally in the range 0.018 to 0.044, and optionally in the range 0.017 to 0.026, and optionally in the range 0.022 to 0.055, and optionally in the range 0.022 to 0.045, and optionally in the range 0.022 to 0.031. The number of stages may be 3 or 4 or 5. The compressor may have compressor blades integrally mounted to a compressor hub driven by a turbine. The engine may have just one compressor positioned downstream of the fan. The skilled person will appreciate that by providing an axial compressor in a gas turbine with the above dimensions it is possible to provide an engine within a small diameter nacelle. By providing air to the compressor in an axial direction, rather than a radial one, the air flow channels may be simplified along with providing a smaller diameter engine. The axial compressor, for flight applications, offers improved fuel efficiency over radial compressors for the same pressure ratios. This may permit a smaller and lighter fuel supply further increasing the efficiency of the engine. In a smaller engine the comparative weight saving is proportionally greater than in a larger engine further enhancing the fuel efficiency. Where the compressor is an integrally bladed rotor there may a weight saving over a bolted compressor. In a smaller engine the comparative weight saving is proportionally greater than in a larger engine further enhancing fuel efficiency and / or the operational efficiency of the engine. Where the compressor is an integrally bladed rotor there may be a length saving over a bolted compressor. The skilled person will appreciate that by providing an axial compressor in a gas turbine with the above dimensions it is possible to address the challenge provided by scaling a conventional engine that results in components and engine systems that are extremely small where In such engines there is a likelihood that tolerances on manufacture or during operation are so extreme that manufacturing rejections and engine life results in uneconomic or unacceptable manufacture or operation. The skilled person would appreciate that shortening an engine of a known engine type may create additional difficulties of creating sufficient compression within the engine. Particularly for low diameter engines known short compressors such as radial compressors may have too large a diameter. Physical limitations of air flow through the engine may provide a minimal axial spacing between components such that scaling a known engine configuration may not be appropriate resulting in a structure where certain aerodynamic flow parameters, for example turn vs axial velocity or volume vs polytrophic efficiency are outside acceptable values. In these arrangements any valid operating window may be unacceptably small. The gas turbine may have an annular combustion chamber downstream of the compressor and upstream of the turbine. The combustor may have one or more fuel injectors at an upstream end. The combustor may have one or more guide vanes at a downstream end. The combustor chamber may be defined between a radially inner wall and a radially outer wall. The combustor radial distance (CRD ) is defined as the distance between the inner wall and the outer wall and may be in the range: 50mm to 110mm, optionally in the range 50mm to 85mm, and further optionally in the range 80mm to 110mm. Within these ranges the radial distance may be above 50mm, or 55mm, or 60mm or 65mm, or 70mm or 75mm or 80mm or 85mm or 90mm, or 95mm, or 100mm or 105mm. The ratio of combustor radial distance to fan radius may be of the order 0.35 to 0.60, and optionally in the range 0.40 to 0.55, and further optionally in the range 0.45 to 0.55, and further optionally in the range 0.40 to 0.50. The combustor volume may be in the range: 1.4L to 20L, optionally in the range 2.3L to 13.3L, and further optionally in the range 2.3L to 7L, and further optionally in the range 5L to 13.3L. Within these ranges the combustor volume may be above 1.5L, or 2.5L or 3.5L, or 4.5L, or 5.5L, or 6.5L, or 7.5L, or 8.5L or 9.5L, or 10.5L, or 11.5L, or 12.5L, or 13.5L, or 14.5L, or 15.5L, or 16.5L, or 17.5L, or 18.5L or 19.5L. Defining the combustor dimensions may allow more stable operation at various altitudes and speeds. A requirement of aero gas turbines that is not shared with land or marine based gas turbines or with hobby aircraft that typically fly below 400ft is that altitude and therefore ambient pressures and temperatures will vary significantly during the operation of the engine. The gas turbine should be able to operate continuously throughout its flight or have a robust restart mechanism. The combustor dimensions defined herein are configured to assist relight. The ratio of combustor dimensions to the fan tip radius may be greater than in known gas turbine engines. It will be appreciated that scaling the combustor from known engine dimensions may provide a combustor that does not enable sufficient energy to be extracted from the fuel and power the engine over the required operational parameters, fail to ignite at altitude, or exhibit a greater than acceptable risk of flame-out during operation. The engine may comprise a nacelle which surrounds a fan and the engine core comprising a turbine, a compressor and a core shaft connecting the turbine to the compressor. The engine core may comprise a casing (also referred to as an inner fixed structure). 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. A 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. The bypass ratio may be in the range: 0.1 to 2.5, optionally the bypass ratio may be in the range 0.2 to 1.5, optionally the bypass ratio may be within the range 0.1 to 0.7, optionally the bypass ratio may be in the range 0.7 to 1.5, optionally the bypass ratio may be up to 3.5. Within these ranges the bypass ratio may be greater than (or on the order of) any of the following: 0.2, 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1.0, 1.1, 1.2, 1.3, 1.4, 1.5, 1.6, 1.7, 1.8, 1.9, 2.0, 2.1, 2.2, 2.3, 2.4. 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). Optionally, for example for an engine with a fan tip radius in the range from 110mm to 225mm, the bypass ratio may be in the range from 0.2 to 1.5. Optionally, for example for an engine with a fan tip radius in the range from 150mm to 180mm the bypass to core ratio may be in the range from 0.5 to 0.9. The use of a bypass duct and a bypass ratio in an engine with a fan tip radius in the range from 110mm to 225mm may improve the Specific Fuel Consumption (SFC) of the engine. In a small engine with a fan tip radius in the range from 110mm to 225mm an improved SFC would increase the range of the engine for a given amount of fuel. Where cargo is being carried by the aircraft the lighter load of fuel may enable a higher cargo load to be carried. In a small diameter engine the use of a bypass flow may be counter-intuitive as it sits outside the engine core and may increases the overall diameter of the engine and / or place additional constraints on the dimensions of the engine core components such as the compressor, combustor, turbine, etc. 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. The skilled person will appreciate that a feature described above in relation to any one of the aspects may be applied, mutatis mutandis, to any other aspect of the invention. For example, in various embodiments any two or more of the conditions for ratios as defined above, and optionally all specified ratio ranges, may apply to any given aspect or embodiment. All aspects may apply to an engine of some embodiments. Furthermore, any feature described below may apply to any aspect and / or may apply in combination with any one of the claims. The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the drive connecting the compressor and the turbine may be a shaft extending between the turbine and the compressor. The drive may be provided by a drive arm extending rearwardly from the compressor that interfaces directly with the turbine such that torque is transmitted. The drive may be provided by a drive arm extending forwardly from the turbine that interfaces directly with the compressor such that torque is transmitted. The drive may be provided by a drive arm extending rearwardly from the compressor that interfaces directly with a drive arm extending forwardly from the turbine such that torque is transmitted between the turbine and the compressor. In any gas turbine engine as described and / or claimed herein, a combustor may be provided axially downstream of the fan and compressor. For example, the combustor may be directly downstream of (for example at the exit of) the compressor. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the turbine. The combustor may be provided upstream of the turbine. The compressor may comprise any number of 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 turbine 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. 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. Each stage of the fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades. BRIEF DESCRIPTION OF THE DRAWINGS Embodiments will now be described by way of example only with reference to the accompanying drawings, which are purely schematic and not to scale, and in which: FIG. 1 shows a radial cross-section through a gas turbine; FIG. 2 shows a simplified axial cross-section through a gas turbine; FIG. 3 shows a strut of FIG. 2 in more detail; FIG. 4 shows multiple simplified axial cross-sections though a gas turbine each with a different arrangement of struts; FIG. 5 shows a simplified axial cross-section through a gas turbine; FIG. 6 depicts an aircraft with a gas turbine. DETAILED DESCRIPTION FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine comprises a nose cone 25, 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 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a compressor 14, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A bypass casing 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. In use, the core airflow A is accelerated and compressed by the compressor 14. The compressed air exhausted from the compressor 14 is directed into the combustion equipment 16 where it is mixed with fuel delivered by an injector 27 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 compressor 14 by a suitable interconnecting drive. The drive may be a shaft or one or more drive arms extending from the compressor and / or turbine. The fan 23 may provide the majority of the propulsive thrust. Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of further example, the gas turbine engine 10 shown in FIG. 1 has a mixed flow nozzle meaning that flow through the bypass duct and the flow through the core are mixed or combined, before (or upstream of) a single nozzle. However, this is not limiting, and various aspects of the present disclosure may also apply to engines in which the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core engine nozzle 20 this is known as a split flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. 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 FIG. 1 ), and a circumferential direction. The axial and radial directions are mutually perpendicular. The compressor 14 comprises multiple compressor stages. In the arrangement shown in FIG. 1 , the compressor 14 comprises four compressor stages. Each stage of the compressor 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 one or more compressor stages may comprise a lowest pressure stage and may further comprise one or more compressor stages of increasing pressure to a highest pressure compressor stage. The lowest pressure compressor stage may be located furthest upstream along the gas flow path within the lowest pressure compressor 14. The further higher pressure stages may be spaced axially along the gas flow path through the compressor in a downstream (rearward) direction. The highest pressure turbine 17 similarly comprises one or more turbine stages. In the arrangement shown in FIG. 1 , the highest pressure turbine 17 comprises one stage. Each turbine stage may comprise a row of rotor blades 17a and a row of stator vanes 17c, 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 lowest pressure turbine 19 similarly comprises one or more turbine stages. In the arrangement shown in FIG. 1 , the lowest pressure turbine 19 comprises one stage. Each turbine stage may comprise a row of rotor blades 19a and a row of stator vanes 19c, 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 one or more turbine stages forming each turbine 17, 19 may comprise a highest pressure stage, and may further comprise one or more turbine stages of decreasing pressure to a lowest turbine pressure stage. The lowest pressure turbine stage may be located furthest downstream within each turbine 17, 19. The further pressure stages are spaced axially in an upstream (forward) direction along the gas flow path through the turbine. In embodiments with only one stage, the single stage is the lowest pressure stage. Each row of rotor blades provided in the compressor 14 may form an annular array of rotor blades carried by a respective rotor hub (or rotor disc). Each of the rotor blades may be integrally formed with the hub or joined to the hub to form an integral assembly. Where the blades are formed with the hub the entire assembly may be formed from a unitary piece of material that is subsequently machined to form both the shaft and hub configuration and the blades. Alternatively, each row of blades and a corresponding portion of the hub may be formed separately through machining of a blank and subsequently joined together to form the integral assembly. Any appropriate machining technique may be used both chemical and mechanical. For example, the assembly may be etched, milled, ground, etc. Any appropriate joining technique may be used e.g. welding, brazing, etc. amongst others. Alternatively, each blade may be coupled to the hub via a root received in a corresponding slot in a peripheral edge of the hub. An assembly may be formed by bolting together, or by using another appropriate mechanical technique, more than one rotor stage. Each rotor blade may be defined as having a radial span extending from the root (or hub) at a radially inner gas-washed location, or 0% span position, to an outer most radial tip at a 100% span position. 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 rotor blade. The radial span each rotor blade refers to the gas-washed portion of the rotor blade, i.e. the portion radially outside any platform at which it is coupled to the hub. Each of the rotor blades forming the compressor stages may have a leading edge mean radius point (or mid blade span) and a trailing edge mean radius point. The mean radius point is defined as the midpoint between the 0% span position and the 100% span position. It may be measured at the rotor blade leading edge (axially forward-most edge) or trailing edge (axially rearward-most edge) to give the leading edge mean radius point and the trailing edge mean radius point respectively. Each row of rotor blades provided in the turbine(s) 17, 19 may form an annular array of rotor blades 17a, 19a carried by a respective rotor hub 17b, 19b (or rotor disc), as shown byway of example in FIGS. 3 . Each of the rotor blades 17a, 19a may be integrally formed with the hub 17b, 19b or joined to the hub to form an integral assembly. Where the blades are formed with the hub the entire assembly may be formed from a unitary piece of material that may be subsequently machined to form both the shaft and hub configuration and the blades. For some uses the near net-shape forming may be sufficiently precise that machining of the aerofoils and the hub may not be required. Alternatively, each row of blades and a corresponding portion of the hub may be formed separately through machining of a blank and subsequently joined together to form the integral assembly. Any appropriate machining technique may be used both chemical and mechanical. For example, the assembly may be etched, milled, ground, etc. Any appropriate joining technique may be used e.g. welding, brazing, etc. amongst others. Alternatively, each blade 17a, 19a may be coupled to the hub 17b, 19b via a root received in a corresponding slot in a peripheral edge of the hub. The gas turbine engine may comprise an electric machine arrangement, which is configured to produce electrical power, and / or provide motive power to start the engine. The electric machine arrangement comprises a starter generator 130, which comprises an electric machine stator 132 and an electric machine rotor 134 separated by a radial air gap 152. The electric machine rotor 134 is coupled to the compressor forward drive arm 123, while the stator 132 is coupled to static structure. The electric machine 130 is provided forward of the compressor 14, and rearward of the fan 12, in an annular space provided radially inward of the gas flow path. The engine core 11 has a casing 30 that is radially outside the compressor, a casing 32 that radially outside the combustor, and a casing 34 that is radially outside the turbine. These casings 30, 32, 34 may be a unitary single piece component or they may be separate components that are bolted together at their axial ends. Bracing is provided between the bypass casing and the engine core to provide stability and prevent excessive movement in addition to absorbing the relative axial and radial thermal expansions of the core relative to the bypass. An exemplary configuration is depicted in Figure 2 which excludes the use of ball joints, pinned hinges or other complex mechanical devices. An inner casing 30 is supported within an outer casing by four struts 42 which span chordally from points on the inner surface of the outer casing to further points on the inner surface of the outer casing. At each end of a strut fasteners 40 are used to secure the strut to the inner wall of the outer casing 21. These may be removable fasteners such as bolts, pins, clips, etc. or it may be permanent fastening such as welding or brazing. At the end of each strut there is a pad 50 that has apertures through which a nut and bolt fastener (if used) may extend to attach the strut to the outer casing. The pad may be angled in the direction of the circumference of the outer casing to provide an increased surface area and better connection between the strut and the casing. At the mid-point of a strut 42 a fastener is provided to secure the strut to the outer wall of the inner casing 30. These may be removable fasteners such as bolts, pins, clips, etc. or it may be permanent fastening such as welding or brazing. A boss 44, or other projection, may be provided which extends radially outwardly from the inner casing around which the strut may engage via an aperture therein. The strut may have a central pad 46 that provides the aperture 48. The boss may be threaded to allow a nut to secure the strut to the inner casing 30. The struts 42 are unitary and tangential to the outer surface of the inner casing to provide a pair of strut elements. Each strut is formed from a high-strength metal sheet that has a thickness T that is small compared to the length L so that they can bend out of plane to absorb the relative radial expansion of the engine core casings. This is shown by the dotted lines in Figure 2. The struts may also be manufactured from a suitable composite laminate such as carbon fibre. A curved change of direction may be provided to the form of the strut. An angular change of direction may be provided to the form of the strut. The change of direction may be in the radial direction. The change of direction may be in the axial direction. The change of direction may be in a strut element in the portion of the strut between the mid-point and the first and / or second end. The change of direction may be at the mid-point. The change of direction may alter the stiffness provided by the struts. The width of the strut element is selected to allow relative axial movement between the outer casing 21 and the inner casing 30 that may be caused by vibration, thrust or thermal inertia. The deflection of the strut element in the axial direction is depicted by the dotted lines in Figure 3. The variables of number of struts, strut length, strut thickness and strut width is described by reference to Figure 4. It is possible to vary the number of struts to provide appropriate support and, for example, three, four, five or six struts may be used to accommodate different spacings between the inner casing and the outer casing. Each strut and strut element length L is limited by the space available in the bypass annulus between the engine core casing 30 and the bypass casing 21. A smaller annulus necessitates shorter chordal struts but also permits a greater number of these shorter struts. To absorb a given relative axial movement between the core and the outer casing with a given level of stress and strain a strut with a shorter length L needs a smaller width W so that it remains flexible. To maintain sufficient structural strength and stiffness the strut must then have a larger thickness T. It has been found that through the selection of appropriate strut element width, length, thickness, material and number the lifetime of the support may be maintained even where the bypass ratio is up to 3.5. Where the mid-point of the strut is attached to core casing 30 it has been found that as the engine core is moved axially relative to the outer casing 21 a couple is generated that may try to rotate the outer pads 40 around the axis of the fastener. This may be reacted by using two or more fasteners in each end pad. At the centre pad the moment applied to each strut is cancelled out so that single boss or fastener is required. In an alternative arrangement as shown in Figure 5 it is possible to form the struts along with an outer ring such that no bolts or fasteners are required to connect the struts to the outer casing. The article 60 may be formed from a machined single piece, or from multiple machined pieces stacked together. The machining may be physical or chemical. In an alternative method of manufacture the article 60 may be manufactured in an additive process. Where multiple sheets are stacked together it is possible to have the grain direction of the sheet metal of each ring angled in a different direction to avoid orthotropy in the mechanical properties of the assembled system. The outer ring comprises a plurality of holes that align with holes in adjacent casing elements 30, 32, 34 such that as the respective casings are bolted together the outer ring of the strut component is secured between them. In some arrangements this allows the joint of the struts to be in the lee of the flange, where the air is already disturbed, which may 5 minimise the pressure loss in a small bypass duct that would otherwise be caused by the strut arrangement. Figure 6 provides a schematic illustration of engines 10a located under the wings 52 of an aircraft 70 and also shows an engine 10b mounted to the top of the fuselage. Either of the engines 10a, or 10b may be removed of altered depending on the required use. It will be 10 appreciated that other configurations may be appropriate such as embedding the engine within the wing, within the fuselage, or elsewhere, and a duct is provided to supply air to the engine. It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the 15 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

1. An arrangement of components the arrangement comprising an outer component and an inner component arranged coaxially about a central axis wherein,the arrangement comprises a plurality of supports, each support being elongate and having first and second ends attached to a surface of the outer component and having a location between the first and second ends where the support is attached to a surface of the inner component.

2. An arrangement according to claim 1, wherein the outer component is cylindrical.

3. An arrangement according to any preceding claim, wherein the inner component iscylindrical.

4. An arrangement according to any preceding claim wherein each end of the first or second ends of the support is attached to the surface of the outer component by multiple bolts.

5. An arrangement according to any preceding claim, wherein the support is attached to the inner component by a single bolt.

6. An arrangement according to any preceding claim, wherein there are 3, 4, 5 or 6 supports.

7. An arrangement according to claim 6, wherein the supports are arranged symmetrically about the central axis.

8. A gas turbine engine comprising an arrangement according to any preceding claim, wherein the outer component is an outer casing of the gas turbine engine and the inner component is an engine core casing.

9. A gas turbine engine according to claim 8, wherein the gas turbine has a bypass ratio in the range 0.1 to 2.5.

10. A gas turbine engine according to claim 8 or claim 9, wherein the gas turbine engine comprises a fan having a plurality of fan blades each with a fan tip, wherein the radial distance between the central axis and the fan tip is in the range 90mm to 225mm.19

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