Patent Publication Number: US-2017356296-A1

Title: Gas turbine engine component

Description:
The present disclosure is a continuation of application Ser. No. 14/044,460 filed Oct. 2, 2013, which claims benefit to British Patent Application No. 1217650.9, filed Oct. 3, 2012, and British Patent Application No. 1217652.5, filed Oct. 3, 2012, which relates to a gas turbine engine component having a cavity defining shell which receives an insert therein, the contents of which are hereby incorporated by reference. The disclosure finds particular use in ceramic matrix composite shells but also in more traditional metal shells. 
    
    
     BACKGROUND 
     Field of the Invention 
     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 always produces more specific thrust (e.g. engine thrust per unit of air mass flow). However, as turbine entry temperatures increase, the life of an uncooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling. 
     In modern engines, the high pressure (HP) turbine gas temperatures are now much hotter than the melting point of the blade materials used, and 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 convection and external films are the main methods of cooling the aerofoils. HP turbine nozzle guide vanes (NGVs) consume the greatest amount of cooling air on high temperature engines. HP blades typically use about half of the NGV cooling air flow. The IP and LP stages downstream of the HP turbine use progressively less cooling air. 
       FIG. 1  shows an isometric view of a conventional HP stage cooled turbine. Block arrows indicate cooling air flows. The stage has NGVs  100  with inner  102  and outer  104  platforms and HP rotor blades  106  downstream of the NGVs, blade platform  112  and shroud  114 . 
     Cooling air can enter NGVs as a single end feed (i.e. in one direction) or a dual end feed (i.e. an inboard and an outboard feed). An aim of the dual feed is to ensure that adequate backflow margin exists at all flight conditions. 
     The NGVs and HP blades are cooled using high pressure (HP) air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000K. Mainstream gas temperatures can be in excess of 2100K. 
     The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Extracting coolant flow therefore has an adverse effect on the engine operating efficiency. Thus, it is important that the cooling air is used as effectively as possible. 
     Improvements in Ceramic Matrix Composite (CMC) technology have resulted in its use in HP turbine components becoming more common. CMC can be used to replace metal static components such as high temperature seal segments, and also, more recently NGVs and other aerofoil components. 
     CMC materials have a high temperature capability and low thermal conductivity. Environmental barrier coatings (EBC) are typically applied to the CMC material. It can be shown that using coated CMC materials such as SiC—SiC, where long multi-strand fibres of silicon carbide are integrated into a silicon carbide matrix, cooling mass flows can be reduced by approximately 40% relative to similar NGV designs made from single crystal nickel alloys. 
     The introduction of CMCs does not eliminate the need for cooling, although the quantity of coolant required to ensure adequate durability reduces considerably. CMCs may be formed by a laser sintering manufacturing process. However, this process can only be used to produce relatively simple non-detailed structures such as a hollow aerofoil shape with a centrally located divider wall. A composite produced by laser sintering will generally be porous, but the addition of a protective coating can help to protect against environmental attack. 
     It is known that additional cooling of a hollow turbine engine component can be achieved by providing sheet metal inserts such as tubes or plates which provide impingement cooling by directing cooling air onto the inside walls of the hollow component. The sheet metal inserts may be adapted to provide location supports in the form of pressed dimples. 
     Wth engine cycle gas temperatures rising and combustion temperature profiles becoming flatter, as a consequence of the drive to reduce NOx and CO 2  emissions, there is an increasing need to make better use of the cooling air in addition to utilising the advantages provided by the CMC material. 
     Although the use of CMC material shells with the inserts of the invention is particularly advantageous, the inserts can be used with non-CMC materials, such as traditional metal shells which may be cast as is known in the art. EP0392664 describes a blade for a combined cycle turbine in which inserts are used to define conduits for the transportation and recovery of steam for cooling purposes. However, how the blades are constructed with the inserts is not described. 
     The present invention seeks to provide inserts which may be placed within shells having irregular cavities which may not ordinarily be able to receive an insert. 
     SUMMARY OF THE INVENTION 
     The invention provides a gas turbine according to the appended claims. In particular there is provided a gas turbine engine component ( 100 ), comprising: a shell having an internal cavity for receiving a multi-part insert; a multi-part insert located within the cavity, wherein the multi-part insert comprises separate insert parts assembled in an abutting relation with one another within the cavity to provide the multi-part insert; an insertion aperture within a wall of the shell which is sized to receive each of the insert parts individually and wherein the multi-part insert cannot be withdrawn from the cavity through the insertion aperture when assembled. 
     The component may include an aerofoil for a gas turbine engine. The component may be a blade or a vane. The blade or vane may be for use in the turbine of the gas turbine engine. 
     The cavity may include an insertion aperture or portion into which the insert parts are inserted, and a receiving portion in which at least one of the insert parts is located when the insert is assembled. The receiving portion may be at least partially obscured by a wall or an internal protuberant feature of the shell when viewed from the insertion aperture. The insertion aperture may be defined by wall of a cavity, or may be defined as part of a larger opening. The insertion aperture may be defined by a portion of a larger opening through which an insertion part can be inserted. The receiving portion may be different for each insert part. The insertion aperture may be different for each insertion part. An shell may have an insertion aperture in each end thereof, each for a different insert part. 
     The obstruction of the receiving portion may be caused by a twist along the length of the cavity. The obstructing wall may be the leading or trailing edge of the aerofoil or the pressure or suction surface wall. Alternatively or additionally, the obstructing wall may be a dividing wall. The obstruction may be due to a distortion in the shape of the cavity. The cavity may be irregularly shaped along the length thereof. The cavity may be twisted or bent along the length thereof. The twist may be chordal. That is, the twist may be provided by an angular offset between a first end and a second end of the aerofoil relative to the longitudinal axis of the aerofoil. The cavity may include one or more features around which the insert must be placed. The one or more features may include cooling holes or projections. 
     The twisting may be due to the aerodynamic profiling of the outer surface of the component. The cavity of the shell may be provided by a wall of the shell. The internal surface of the shell may be smooth. That is, the internal surface of the shell may be devoid of surface features. Such features may include but are not restricted to cooling and turbulating features such as pedestals and trip strips. 
     The cavity may widen along the length thereof and the receiving portion may be located within the widened portion of the cavity. The cavity may include a recess. The recess may provide a receiving portion for part of an insert. The recess may be towards the trailing edge of the blade. The recess may be provided by another part of the multi-part insert. 
     The maximum width of the assembled multi-part insert may be greater than that of the maximum width of the insertion aperture. 
     The assembled insert may include at least one retention part, wherein the retention part acts to engage with a portion of the cavity and the at least one other insert part so as to retain the assembled insert within the cavity. 
     The retention part may provide an interference fit with other insert parts and or a wall of the shell so as to provide a chock. The retention part may provide a resilient bias. 
     The retention of the assembled insert with the retention part may be for assembly purposes only. As such, the insertion aperture may be partially or completely blocked after the insert is located within the cavity. For example, the insertion aperture may be covered with a cap or plate attached over the insertion aperture. 
     The retention part may provide a resilient bias which acts to urge the retention part and or another insert part against one or more walls of the shell. 
     The retention piece may include two members joined at a hinge portion. The hinge portion may be sprung loaded to provide the resilient bias. The hinge portion may be plastically deformed prior to assembly. The hinge portion may be connected to the members so as to provide an angle of separation between the two members. The angle of separation between the members may be greater prior to assembly such that the arms need to be forcibly moved together for insertion into the cavity. Forcing the arms of the retention part together can elastically deform the hinge part such that it is resiliently biased against a wall of a cavity or another one of the insert parts when the retention part is placed in situ. 
     The retention part may be oversized relative to the size required when in situ such that inserting the retention part into the cavity requires a deformation of the part and a resulting stressing to provide the bias. 
     At least one insert piece may be made by additive layer manufacturing. The insert may include formations which support the insert within the shell and guide the cooling air around the inner surface of the shell. The formations may include projections. The shell may be a ceramic matrix composite shell. 
     The projections may be fins. The fins may be pin-fins. The formations may form one or more chambers, between the insert and the inner surface of the shell, the or each chamber being configured so that, in use, the chamber receives cooling air from the one or more flow channels, the cooling air pressure being lower in the chamber than in the flow channels. The insert may form a plurality of flow channels in fluid communication with one another to define a multi-pass cooling arrangement. 
     At least one of the insert parts may predominantly include trip strip formations which lie along the inner surface of the shell when the insert is assembled. The separate insert parts may include one or more support structures for engagement with the insert and the trip strip formations. The strip trip formation and or support structures may be elongate members in the form of bars or rods. The strip trip insert part may have a ladder like construction. 
     The shell forms an aerofoil and includes a divider wall which divides the shell into a front cavity at a leading edge region of the component and a rear cavity at a trailing edge region of the component. The multi-part insert may be located in the front cavity or a rear insert located in the rear cavity. 
     The divider wall may include apertures which provide fluid communication between the front and rear cavity. At least one part of the multiple insert parts may include a sealing plate to restrict or prevent the flow of cooling air across the divider wall. The sealing plate may be incorporated on the retaining part. Alternatively or additionally, the sealing plate may be formed by one or more insert parts. 
     In another aspect, the invention provides a method of forming the gas turbine engine component according to any one of the previous claims, the method including the steps of: providing the shell; providing a plurality of insert parts which are configured to be assembled in an abutting relation with one another within the cavity to provide the multi-part insert; wherein the assembled insert includes at least one retention part which engages with a wall of the cavity and at least one other insert part so as to retain the assembled insert within the cavity. 
     Other preferred features include , a gas turbine engine component having a shell and an insert located inside the shell, the insert forming one or more flow channels which, in use, receive a flow of cooling air; wherein the insert is made by additive layer manufacturing; and wherein the insert includes formations which support the insert within the shell and guide the cooling air around the inner surface of the shell. 
     The formations formed as part of an insert made by additive layer manufacturing (ALM) may be intricate features which cannot be formed as part of the shell and which cannot be formed with a high level of dimensional accuracy on inserts produced from sheet metal. By providing ALM inserts with supporting formations, the cooling properties of the component can be greatly improved. 
     The use of ALM for the production of metal inserts can also be advantageous in that the walls of the insert including the impingement holes can be manufactured in one procedure without requiring a separate tooling step to manufacture the holes. Further, the inserts can be readily modified without a need for expensive re-tooling and the time taken to manufacture inserts can be reduced. Where the insert is metallic, the ALM process can be direct laser deposition (DLD) (also known as direct metal deposition (DMD)). 
     The shell may be a ceramic matrix composite shell. By providing ALM inserts with supporting formations, the cooling properties of a component having a CMC shell can be greatly improved. More generally, it is possible to add to the benefits provided by a CMC shell, such as its thermal properties, by providing detailed structures that cannot be manufactured as part of the CMC shell. 
     Alternatively, the shell may be a metal shell, such as single crystal nickel alloy shell. The formations may include fins. These fins may extend to the inner surface of the shell to support the insert within the shell. The fins may be pin-fins which advantageously enhance the heat transfer level by increasing the turbulence of the cooling air flow and providing mixing of the cooling air. The insert may also include impingement holes for jetting cooling air from one or more flow channels onto the inner surface of the shell. 
     The formations may form one or more chambers between the insert and the inner surface of the shell, the or each chamber being configured so that, in use, the chamber receives cooling air from the one or more flow channels, the cooling air pressure being lower in the chamber than in the flow channels. Each chamber can contain cooling air at a different pressure. The or each chamber can supply film cooling holes formed in the shell, the pressure of cooling air at the film cooling holes being matched to the local external pressure. 
     The insert may be tubular so that it forms a central flow channel and fits inside the shell in a nested arrangement with formations protruding outwardly from an outer wall of the insert towards the inner surface of the shell. In this way, the chambers can be located around the central flow channel. Another option is for the insert to be a plate which extends from one part of the inner surface of the shell to another part of the inner surface of the shell to form a flow channel on at least one side of the insert. 
     The insert may form a plurality of flow channels in fluid communication with one another to define a multi-pass cooling arrangement. In such a multi-pass cooling arrangement, the cooling air can flow in opposite directions through successive channels. Integral plates may be located at end walls of the component to create suitable bend geometries between channels. 
     The insert may include trip strip formations which lie along the inner surface of the shell. These formations can improve heat transfer to the cooling air. 
     The gas turbine engine component may be an aerofoil. More particularly, the gas turbine engine component may be a nozzle guide vane (NGV) or a rotor blade. However, it is also possible that the gas turbine engine component could be an NGV platform, a shroud segment or a shroud liner. 
     When the component is an aerofoil, the shell can include a divider wall which divides the shell into a front cavity at a leading edge region of the component and a rear cavity at a trailing edge region of the component. The divider wall can help to prevent the aerofoil structure from rupturing under pressure loads and also can help to prevent unwanted ballooning of the aerofoil shape. The insert may be a front insert located in the front cavity or a rear insert located in the rear cavity. Indeed, the aerofoil may include respective inserts in both the front cavity and the rear cavity. 
     When the insert is a rear insert, one or more chambers defined by the insert can supply cooling air to trailing edge discharge holes or slots, with the holes or slots receiving cooling air at a pressure matched to the local external pressure. 
     The insert may include a sealing plate to prevent the flow of cooling air across the divider wall. Such a sealing plate can allow the divider wall to be discontinuous. In preventing the flow of cooling fluid across the divider wall, the sealing plate can help to reduce thermal induced stresses associated with hot external walls and a cold divider. 
     The insert may be a unitary body, or may be formed from two or more separately insertable insert parts. Forming the insert from a plurality of insert parts can allow the insert to be fitted into a shell which has, for example, a re-entrant cavity or is otherwise configured in such a way as to prevent a unitary body from being inserted. 
     A gas turbine engine component may be provided having a shell and an insert located inside the shell, the insert may include: a first wall containing first impingement holes which, in use, jet cooling air onto a first region of the inner surface of the shell; a second wall containing second impingement holes which, in use, jet cooling air onto a second region of the inner surface of the shell; and a fluid pathway formed between the two walls, the pathway recycling the cooling air jetted onto the first region to the inlets of the second impingement holes for jetting onto the second region. 
     Advantageously, the insert allows jetted cooling air to be used twice. In this way, film cooling effectiveness and film coverage can be increased for a given quantity of cooling air mass flow. 
     The shell may be a ceramic matrix composite shell. Alternatively, the shell may be a metal shell, such as single crystal nickel alloy shell. 
     The insert may be made by additive layer manufacturing (ALM) or by casting. Where the insert is metallic, the ALM process can be direct laser deposition (DLD) (also known as direct metal deposition (DMD)). An insert made by ALM or casting can be produced with a high level of intricacy and with high speed and repeatability. For example, ALM facilitates the production of features such as thin walls and internal cooling holes, as well as internal heat transfer augmentation features like trip-strips, pedestals, pin-fins etc. 
     The insert may include heat transfer formations at the first and second regions which support the insert within the shell and which guide the cooling air around the inner surface of the shell. In this way, the cooling air can remove more heat from the walls of the shell. In addition, as the insert supports itself, there may be no need for extra support structures which can add to manufacturing time and cost. 
     The geometry of the heat transfer formations at the first region in particular may be chosen to restrict the flow rate of the cooling air and to increase the pressure drop through the pathway. The heat transfer formations may be pedestals or pin-fins, in which case the flow rate of the cooling air may be controlled by the number of pedestals/pin-fins, their density and their diameter. Additionally or alternatively, the number of the impingement holes and/or the diameter of the impingement holes can be used to control the flow rate of the cooling air. 
     The shell may include exterior film cooling holes fed by cooling air that has been jetted onto the second region of the inner surface. This further recycling of the cooling air helps to make even more effective use of the air. 
     The insert may include trip strip formations which lie along the inner surface of the shell. 
     The gas turbine engine component may be an aerofoil. More particularly, the gas turbine engine component may be a nozzle guide vane (NGV) or a rotor blade. However, it is also possible that the gas turbine engine component can be an NGV platform, a shroud segment or a shroud liner. 
     Where the component is an aerofoil, the first and second regions may be located at the suction side of the aerofoil. 
     The shell of the aerofoil may include a divider wall which divides the shell into a front cavity at a leading edge region of the aerofoil and a rear cavity at a trailing edge region of the aerofoil. The insert can then be a front insert located in the front cavity, or a rear insert located in the rear cavity. Indeed, the aerofoil may have respective inserts in both the front cavity and the rear cavity. The divider wall can help to prevent the aerofoil structure from rupturing under pressure loads and also helps to prevent unwanted ballooning of the aerofoil shape. The insert may include a sealing plate to prevent a flow of cooling air across the divider wall. In preventing such a flow, the sealing plate can reduce thermal induced stresses associated with hot external walls and a cold divider. 
     Where the insert of the aerofoil is a front insert, the pathway may guide the recycled cooling air in an upstream direction towards the leading edge. In this way, for the front cavity, the first region of the inner surface of the shell may be located further away from the leading edge of the aerofoil and the second region of the inner surface of the shell may be located closer to the leading edge. Any exterior film cooling holes fed by cooling air that has been jetted onto the second region may therefore lie at a position close to the leading edge, and can contribute to a cooling film on the suction side of the aerofoil. 
     Where the insert is a rear aerofoil insert, the pathway may guide the recycled cooling air in a downstream direction towards the trailing edge. In this way, for the rear cavity, the first region of the inner surface of the shell may be located further away from the trailing edge of the aerofoil and the second region of the inner surface of the shell may be located closer to the trailing edge. 
     The insert may also include a bank of further heat transfer formations, such as pedestals or pin fins, along the inner surface of the shell to guide the cooling air along the inner surface of the shell after it has been jetted onto the second region. In respect of a rear aerofoil insert, the bank of further heat transfer formations preferably guides the recycled cooling air in a downstream direction towards the trailing edge to feed exit holes or slots at the trailing edge. 
     The aerofoil insert may define one or more flow channels which, in use, collect cooling air from one or both ends of the aerofoil and distribute the cooling air through the shell, at least a portion of the cooling air being distributed to the inlets of the first impingement holes for jetting onto the first region. 
     The insert may be a unitary body, or may be formed from two or more separately insertable insert parts. Forming the insert from a plurality of insert parts can allow the insert to be fitted into a shell which has, for example, a re-entrant cavity or is otherwise configured in such a way as to prevent a unitary body from being inserted. 
     Further optional features of the invention are set out below. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which: 
         FIG. 1  shows an isometric view of a conventional HP stage cooled turbine; 
         FIG. 2  shows a longitudinal cross-section through a ducted fan gas turbine engine; 
         FIG. 3  shows cross sectional views of (a) a ceramic matrix composite shell of a nozzle guide vane found in the circled region labelled R in  FIG. 2  and (b) front and rear inserts to the shell; 
         FIG. 4  shows a cross-sectional view of the nozzle guide vane of  FIG. 3  with the inserts fitted inside the shell and cooling flows indicated by arrows; 
         FIG. 5  shows a cross-sectional view of a second nozzle guide vane; 
         FIG. 6  shows a cross-sectional view of the nozzle guide vane of  FIG. 5  with cooling flows indicated by arrows; 
         FIG. 7  shows a cross-sectional view of variant inserts for the nozzle guide vane of  FIGS. 5 and 6 ; and 
         FIG. 8  shows cross-sectional views of (a) a ceramic matrix composite shell of a nozzle guide vane found in the circled region labelled R in  FIG. 2 , and (b) front and rear inserts to the shell; 
         FIG. 9  shows a cross-sectional view of the aerofoil of  FIG. 8  with the inserts fitted inside the shell; and 
         FIGS. 10 a  to 12 b    show various aerofoil embodiments having multi-part inserts according to the invention. 
     
    
    
     DETAILED DESCRIPTION AND FURTHER OPTIONAL FEATURES OF THE INVENTION 
     Wth reference to  FIG. 2 , a ducted fan gas turbine engine incorporating the invention is generally indicated at  10  and has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake  11 , a propulsive fan  12 , an intermediate pressure compressor  13 , a high-pressure compressor  14 , combustion equipment  15 , a high-pressure turbine  16 , and intermediate pressure turbine  17 , a low-pressure turbine  18  and a core engine exhaust nozzle  19 . A nacelle  21  generally surrounds the engine  10  and defines the intake  11 , a bypass duct  22  and a bypass exhaust nozzle  23 . 
     During operation, air entering the intake  11  is accelerated by the fan  12  to produce two air flows: a first air flow A into the intermediate pressure compressor  13  and a second air flow B which passes through the bypass duct  22  to provide propulsive thrust. The intermediate pressure compressor  13  compresses the air flow A directed into it before delivering that air to the high pressure compressor  14  where further compression takes place. 
     The compressed air exhausted from the high-pressure compressor  14  is directed into the combustion equipment  15  where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines  16 ,  17 ,  18  before being exhausted through the nozzle  19  to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors  14 ,  13  and the fan  12  by suitable interconnecting shafts. 
     A first example of a component having an insert will be described with reference to  FIGS. 3 and 4 .  FIG. 3  shows cross sectional views of (a) a ceramic matrix composite (CMC) shell of a gas turbine engine component in the form of a nozzle guide vane (NGV) as found in the circled region labelled R in  FIG. 2 , and (b) front and rear inserts to the shell.  FIG. 4  shows a cross-sectional view of the aerofoil of  FIG. 3  with the inserts fitted inside the shell and arrows indicating cooling air flows. 
     The NGV shell includes a divider wall  203  which divides the shell into a front cavity  201  at a leading edge region of the aerofoil and a rear cavity  202  at a trailing edge region of the aerofoil. The dividing wall may have apertures  204 . A front insert  210  made by direct laser deposition (DLD) (a form of additive layer manufacturing) is located inside the front cavity  210  and a rear insert  220  also made by DLD is located inside the rear cavity  202 . 
     The CMC shell includes film cooling holes  206  located at a region of the suction side of the aerofoil closest to the leading edge. Film cooling holes  206  are also located along the pressure side of the aerofoil. A cooling flow outlet  207  is located at the trailing edge of the CMC shell, in fluid communication with the rear cavity, and may take the form of exit holes or slots. 
     Each of the DLD inserts  210 ,  220  of  FIGS. 3 and 4  has a tubular shape similar to the shape of the front and rear cavities so that the front insert  210  is located inside the front cavity  201  in a nested arrangement, and the rear insert  220  is located inside the rear cavity  202  in a nested arrangement. Each tubular insert defines a central flow channel  211 ,  212 , and cooling air is bled out from each central flow channel to the inner surface of the shell via impingement holes  216  formed in the walls of the insert. 
     Each DLD insert  210 ,  220  includes formations  218 ,  219  which extend outwards from an outer surface of the insert to an inner surface of the shell to support the insert within the shell and guide cooling air around the inner surface of the shell. The formations include pin-fin formations  218  and chamber-forming formations  219 . 
     The rear insert  220  includes a sealing plate  270  located along the divider wall  203  of the shell to help prevent the flow of cooling air across the divider wall  203 . 
     The flow of cooling air will now be described with reference to  FIG. 4 . Large shaded arrows depict the flow of cooling air into the aerofoil, inboard  311  and outboard  312  flows entering the front cavity  201 , and a single inboard flow of cooling air  313  entering the rear cavity  202 . Where the flow is a dual feed (an inboard and outboard flow), the insert  210  preferably includes a baffle plate (not shown). The baffle plate reduces differential pressures caused by the dual feed, thereby reducing unwanted ‘blow through’ effects. The baffle plate can be formed as an integral part of the insert  210 , which advantageously reduces the part count and cost, and improves reliability. 
     The chamber-forming formations  219  form a plurality of chambers  229  between each insert and the inner surface of the shell  200 . Each chamber  229  is configured to receive cooling air from a flow channel  211 ,  212  via impingement holes  216 , the pressure of the cooling air being lower in the chambers than in the flow channel. Cooling air from the chambers  229  is used to supply film cooling holes  206 . The formations  219  of the front insert of the aerofoil shown in  FIG. 4  form four chambers between the insert  210  and the inner surface of the shell  200 . A first chamber supplies cooling air to film cooling holes  206  on the suction side, a second chamber supplies cooling air to showerhead cooling holes  206  at the leading edge region of the pressure side, and third and fourth chambers supply cooling air to film cooling holes on the pressure side further away from the leading edge. 
     The number of impingement holes  216  supplying a given chamber and the number of film cooling holes  206  fed by that chamber are selected so that each chamber is maintained at a different pressure. Cooling air can therefore be supplied to the film cooling holes  206  and the film cooling outlet  207  at pressures which match the local external pressure. The front flow channel  211  has an internal pressure level which is controlled to ensure adequate blowing rates through these film cooling holes, while maintaining a safe backflow pressure margin to prevent hot gas ingestion throughout the flight cycle. 
     A second example of a component having an insert will be described with reference to  FIGS. 5, 6 and 7 .  FIG. 5  shows a nozzle guide vane  300  according to the second example,  FIG. 6  shows the cross-sectional view of  FIG. 5  with arrows indicating cooling air flows, and  FIG. 7  shows a cross-sectional view of variant inserts for the nozzle guide vane of  FIGS. 5 and 6 . The NGV has a CMC shell  400 , including a divider wall  403  which divides the shell into a front cavity at a leading edge region of the aerofoil and a rear cavity at a trailing edge region of the aerofoil. The dividing wall may have apertures  404 . A front insert  410  made by DLD is located inside the front cavity and a rear insert  420 , also made by DLD, is located inside the rear cavity. Each insert  410 ,  420  includes a sealing plate  470  to prevent the flow of cooling air across the divider wall  403 . 
     The front insert  410  of the aerofoil shown in  FIGS. 5 and 6  has formations, including a plate end  419  and pin-fins  418 , which extend from an insert plate  440  to the inner surface of the shell to support the insert within the front cavity of the shell, and thereby define a flow channel  411  at the pressure side of the front cavity between the front insert and the inner surface of the shell, and a chamber at the suction side of the front cavity between the front insert and the inner surface of the shell. The chamber on the suction side receives cooling air from the flow channel  411  via impingement holes  416 . 
     The rear insert  420  of the aerofoil shown in  FIGS. 5 and 6  has formations in the form of sealing walls  475  which extend outwardly from a central insert plate  430 , to the inner surface of the shell. There are four sealing walls  475  which, in addition to the sealing plate  470 , define a plurality of flow channels  441 ,  442 ,  443  in fluid communication with one another to form a multi-pass cooling arrangement. 
     In  FIG. 6 , large straight arrows  512 ,  513  and  514  depict flows of cooling air into the aerofoil  400 . The multi-pass cooling arrangement includes, in flow series, a pair of parallel first pass chambers  441  (one on the pressure side and one on the suction side) corresponding to a first pass flow channel, a pair of parallel second pass chambers  442  (one on the pressure side and one on the suction side) corresponding to a second pass flow channel and a common third pass chamber  443  corresponding to a third pass flow channel. The third chamber is located at a trailing edge region of the rear cavity and feeds trailing edge discharge holes or slots  407 . The first pass chamber  441  on the pressure side supplies film cooling holes  406  on the pressure side of the NGV. Similarly, the second pass chamber  442  on the pressure side supplies film cooling holes  406  on the pressure side of the NGV. 
     Integral plates at end walls (not shown) create suitable bend geometries to guide cooling air from the first pass chambers  441  to the second pass chambers  442  and from the second pass chambers to the third chamber  443  in order that the chambers operate as the rearward flowing, 3-pass cooling arrangement shown by the curved arrows. 
       FIG. 7  shows variant front and rear inserts similar to those of  FIGS. 5 and 6  but having additional trip strip formations  460  which lie along the inner surface of the shell. The trip strip formations are ladder-like in construction having a pair elongate parallel rails which provide support for a linear array of equally spaced trip strips or bars which run therebetween. The trip strips are set at a compound angle to the rails, Trip strips are known in the art and can locally enhance heat transfer to the cooling air. 
     Although not shown in the above Figures, formations defining a contra-flow cooling system can be incorporated into an insert, as an alternative or an addition to the cooling structures described above. 
     Any holes  216 ,  416  in the insert can be formed during the DLD process so there is no need for subsequent machining of the inserts. 
     In addition, the DLD process facilitates modification and development of the insert design during the manufacturing process as no tooling changes are required. For example, features such as formations  218 ,  219 ,  418 ,  419 ,  460 ,  475  may be altered slightly between the manufacture of different aerofoils  100 ,  400  of a single engine  10  depending on the position of the respective aerofoils within the engine to give a relative increase or decrease in the cooling mass flow of the aerofoil. 
       FIG. 8  shows cross-sectional views of (a) a ceramic matrix composite (CMC) shell  800  of a gas turbine engine component  100  in the form of a nozzle guide vane (NGV) found in the circled region labelled R in  FIG. 2 , and (b) front  210  and rear  820  inserts to the shell.  FIG. 9  shows a cross-sectional view of the aerofoil of  FIG. 8  with the inserts fitted inside the shell and arrows indicating cooling air flows. 
     The shell  800  includes a divider wall  803  which divides the shell into a front cavity  801  at a leading edge region of the aerofoil and a rear cavity  802  at a trailing edge region of the aerofoil. The front insert  810  is located inside the front cavity  801  and the rear insert  820  is located inside the rear cavity  802 . 
     The CMC shell  800  includes exterior film cooling holes  806  located at the region of the suction side of the aerofoil closest to the leading edge. More exterior film cooling holes  806  are located along the pressure side of the aerofoil. The CMC shell  800  also includes exit holes or slots  807  at its trailing edge. 
     Each of the front and rear inserts includes a first wall  811 ,  821  having first impingement holes  813 ,  823  formed therein and a second wall  812 ,  822  having second impingement holes  814 ,  824  formed therein. For each insert, a fluid pathway  815 ,  825  is formed between the first wall  811 ,  821  and the second wall  812 ,  822 . 
     The first impingement holes  813 ,  823  lie opposite a first region  833 ,  843  of the inner surface of the shell and the second impingement holes  814 ,  824  lie opposite a second region  834 ,  844  of the inner surface of the shell. The first and second regions of the aerofoil of  FIGS. 3 and 4  are both located at the suction side of the aerofoil. For each insert, the fluid pathway is formed between the first region  833 ,  843  and the inlets of the second impingement holes  814 ,  824  to recycle cooling air which has been jetted onto the first region for jetting onto the second region. 
     The fluid pathway  815  of the front insert guides recycled cooling air in an upstream direction towards the leading edge so that, for the front cavity, the first region  833  is located further away from the leading edge of the aerofoil and the second region  834  is located closer to the leading edge of the aerofoil. The fluid pathway  825  of the rear insert guides recycled cooling air in a downstream direction so that, for the rear cavity, the first region  843  is located furthest away from the trailing edge of the aerofoil and the second region  844  is located closest to the trailing edge  807  of the aerofoil. 
     Heat transfer formations  853  are located at the first region  833 ,  843  and the second region  834 ,  844 . The heat transfer formations shown in  FIGS. 3( b )  and  4  are pin-fins. 
     In addition to the first wall  821  and second wall  822 , the rear insert  820  shown in  FIGS. 3( b )  and  4  includes a bank of pin-fins  863  which extend along the inside surface of the shell from the second region to the trailing edge. The rear insert also defines a plurality of chambers  881 ,  882  at the pressure side of the rear cavity. The chambers are interconnected via internal passageways  829  so that they are in fluid communication with each other. Two chambers  881 ,  882  are shown in the rear insert of the aerofoil of  FIGS. 3( b )  and  4 . 
     Each insert  810 ,  820  includes a sealing plate  870  which lies along the divider wall  803  of the CMC shell  800  to prevent a flow of cold air across the divider wall. The rear insert  820  also includes trip strip formations  816  which lie along the inner surface of the shell at the pressure side of the cavity to improve heat transfer to the cooling air at this location. 
     The flow of cooling air will now be described with reference to  FIG. 9 . Large shaded arrows depict the flow of cooling air into the aerofoil: inboard  911  and outboard  912  flows entering the front cavity  801 , and a single inboard flow of cooling air  913  entering the rear cavity  802 . Where the flow is a dual feed (an inboard and an outboard flow), the insert preferably includes a baffle plate (not shown). The baffle plate reduces differential pressures caused by the dual feed, therefore reducing unwanted ‘blow through’ effects. The baffle plate can be formed as an integral part of the insert which advantageously reduces the part count and cost and improves reliability. 
     In the front cavity  801 , the first wall  811  defines a front flow channel  860  at the pressure side of the cavity. Cooling air is distributed from this front flow channel to the inlets of the first impingement holes  813  for jetting onto the first region  833 . The front flow channel also supplies cooling air at a high pressure to film cooling holes  806  on the pressure side in the form of a leading edge showerhead cooling head arrangement. The front flow channel has an internal pressure level which is controlled to ensure adequate blowing rates through these cooling holes, while maintaining a safe backflow pressure margin to prevent hot gas ingestion throughout the flight cycle. Cooling air which has been recycled and jetted onto the second region  834  will have a reduced pressure compared to the cooling air supplied directly by the front flow channel and can therefore be used to feed exterior film cooling holes  806  on the suction side. 
     In the rear cavity  802 , the plurality of chambers  881 ,  882  on the pressure side form a plurality of rear flow channels. Cooling air enters the first chamber  881  and is distributed therefrom to the inlets of the first impingement holes  823  for jetting onto the first region  843 . This first chamber also supplies cooling at a high pressure to exterior film cooling holes  806  on the pressure side of the aerofoil, as well as supplying cooling air to the second chamber  882  via internal passageways  829 . The second chamber supplies cooling air to the bed of pin-fins  863  as well as to further exterior film cooling holes  806  on the pressure side. Both chambers have internal pressure levels which are controlled to ensure adequate blowing rates through their cooling holes, while maintaining a safe backflow pressure margin to prevent hot gas ingestion throughout the flight cycle. 
     The CMC shell may be SiC—SiC and a protective coating may be applied to the outside and/or inside surfaces of the shell  800  to prevent environmental attack. The inserts  810 ,  820  may be cast (e.g. using the lost wax process) and then machined (e.g. for hole drilling), or may be made using additive layer manufacturing such as direct laser deposition (also known as direct metal deposition). Additive layer manufacturing, and particularly direct laser deposition, enables all of the detailed features of the inserts to be manufactured in one procedure, including the impingement holes  813 ,  814 ,  823 ,  824 . Further, it allows cooling schemes to be easily changed, without the need for re-tooling. 
     The gas turbine component of the present invention can be an NGV aerofoil, as described in detail in above, but can be any other gas turbine aerofoil, including a rotor blade. The gas turbine component may alternatively be an NGV platform, a shroud segment, or a shroud liner. 
     The inserts described above can be used instead of, or in combination with, sheet metal inserts. 
     Instead of forming each insert as a unitary body, as shown in  FIGS. 3 to 9 , another option is to form the inserts from two or more insert parts. This allows the inserts to be fitted into cavities where a receiving portion in which part of the insert would ideally be located is obstructed in some way such that a complete insert cannot be directly inserted. The obstruction in question may be provided by a wall of the cavity or by a protuberant feature which extends from one or between two walls of the cavity. An obstructed portion may be as viewed from outside the shell through an insertion aperture, or by a part of the insert having to enter the cavity along a first trajectory before being located in a receiving portion along a second trajectory which is different to the first trajectory. For example, an elongate insert part having a longitudinal axis may be inserted into the cavity with an axially extending trajectory, before being pushed laterally into a recess or an otherwise obscured portion of the cavity. 
       FIGS. 10 a , and 10 b    show a perspective view of an aerofoil having a front insert  1010  which is a variant of the front insert of  FIGS. 5 and 6 , and a rear insert  1020  which is a variant of the rear insert of  FIG. 7 , the CMC shell  1000  being drawn as a transparent body. 
     Thus, in  FIGS. 10 a  and 10 b    there is shown an aerofoil in the form of a vane similar to the NGV shown in  FIG. 1 . The aerofoil includes an elongate shell  1000  having internal front  1001  and rear  1003  cavities. The outer surface of the shell has a predetermined aerodynamic shape suitable for use as an NGV. As such, the aerofoil is distorted from a straight radially extending form and includes a chordal twist along its length. This distortion can be best seen in  FIG. 10 b    where the first end  1000   a  and second end  1000   b  of the aerofoil are angularly offset from each other when viewed approximately along the longitudinal axis of the aerofoil  1000 . This means that the front  1001  and rear cavities which extends along the radial axis of the interior of the aerofoil  1000  have an irregular shape with obstructed portions when viewed from the first end along the longitudinal axis of the shell  1000 . 
     It will be appreciated that the distortion of the cavities is also affected by the internal profile of the shell walls which may be varied but will typically be determined by the weight and mechanical and thermal requirements of the aerofoil rather than the fit of an insert. In the described example, the walls of the shell have substantially uniform thickness. 
     The front cavity  1001  has a multi-part insert  1010  located therein, which, in the described example, is made up from two separate insert parts  1010   a,b  assembled in an abutting relation to one another so as to provide the multi-part insert  1010 . The rear cavity  1003  also includes a multi-part insert  1020  having multiple separate insert parts  1020   a - f . The rear cavity insert  1020  is made up from two main body parts  1020   a,b  and several trip-strip insert parts  1020   c - f  which abut and engage the main body portions  1020   b  of the rear insert  1020 , and also the wall of the shell  1000 . Thus, the front insert  1010  is a multi-part insert formed from two insert parts  1010   a,b  and the rear insert  1020  is formed from six insert parts  1020   a - f . In each cavity, the last insert part to be installed locks the completed insert in place and ensures a tight fit between the insert and the shell  1000  while accommodating manufacturing tolerances. 
     To construct the vane with the assembled inserts  1010 ,  1020 , the insert parts  1010   a,b ,  1020   a - f , are placed within the respective cavities via an insertion aperture  1050 . The insertion aperture  1050  may be any suitable entrance to the cavity and may be covered and optionally sealed after the inserts  1010 ,  1020  have been correctly located within the shell  1000 . In the described example, the insertion aperture  1050  is provided by the open end of the aerofoil and is as large as can be accommodated by the walls of the shell  1000 . It will be appreciated that some constructions of the component, particularly one which is cast for example, may only include a partial opening in the end of the aerofoil. Further, the insertion aperture may be defined by the walls of the shell, or a particular portion or zone of a larger opening. 
     Although the insertion aperture  1050  of the rear cavity  1020  is as large as can be accommodated, the irregular shape of the rear cavity  1020  means that the insertion of the assembled or unitary insert  1010 ,  1020  into the cavity  1020  would not be possible. This is because an insert which is shaped to match and abut the internal walls of the cavity may be too large in parts to fit through the insertion aperture  1050 . Alternatively, the curvature or twist of the insert may prevent it from being inserted along the length of the cavity. Further, there may be features or recesses within the cavity which the insert must either go around or be placed within when being inserted. Thus, although the use of prior art inserts has provided some benefits, applications have been limited due to the restrictions placed on the inserts. 
     Providing a multi-part insert allows a first insert part to be loaded into the cavity via an insertion aperture and subsequently located into a receiving portion of the cavity. Thereafter, the second insert part, or retaining part, is passed into the cavity and engaged with the first insert part in an abutting manner. The retaining part may provide a biasing force which acts to urge the first insert part against a wall of the cavity so as to retain it there, or may be manufactured to have an interference fit with the first insert part so as to provide chock. Thus, there is provided an assembled insert within the cavity which cannot be withdrawn from the insertion aperture (or inserted if assembled outside of the shell), but which can be located against the wall of the shell. 
     In some embodiments, the resilient part may be the first or an intermediate part loaded into the cavity. In this instance, the loading of the resilient part will occur upon insertion of the last part which will act to put the resilient part in a stressed condition. 
     In the described example of  FIGS. 10 a  and 10 b   , a receiving portion  1060  can be taken to the rearmost portion of the rear cavity  1003  in which the first insert part  1010   a  is located. The insertion aperture  1050  can be taken to be at the first end  1000   b  of the aerofoil toward the divider wall  1004 . Thus, the first insert part  1020   a  is inserted into the rear cavity  1003  through the insertion aperture  1050  which is located at the wider end of the open ended aerofoil towards the divider wall  1004  and with a trajectory which is coincidental with the plane of the divider wall  1004 . Once in place, the first insert part  1020   a  can be moved toward the rear of the cavity until the distal ends of partitioning walls  1021  abut the walls of the cavity. It will be appreciated that the trip strip formations  1020   c  and  1020   d  can be mated to the first insert part before or after the insertion depending on the particular design, but it is envisaged that they are mated to the first main body insert part  1020   a  prior to being loaded into the rear cavity  1003 . Next, the second main body insert part  1020   b  and third trip strip  1020   e  can be placed within the rear cavity  1003  via the insertion aperture  1050  and pushed home to provide a chock for retaining the first insert part  1020   a  in place. The final insert part is the fourth trip strip  1020   f  formation which is slid between a free end of a web of the first main insert part  1020   a,  and a shoulder  1005  which protrudes into the rear cavity along the length of the divider wall  1004  where the divider wall meets the shell wall. 
     It will be noted from  FIG. 10 b   , that the shape of the rear cavity  1003  would prevent the insertion of the assembled insert  1020  into the cavity from the open end of the vane due to the variance in amount of the chordal twist required between the front and rear parts of the assembled insert. 
     The two insert parts  1010   a,b  of the front cavity  1001  include a curved member  1010   a  which sealably contacts the interior of the leading edge of the aerofoil and extends around the suction side toward the divider wall  1004 . The second insert part  1010   b  is in the form of a sealing plate  1014  which sealably abuts the divider wall  1004 . The sealing plate  1014  includes a short wall along its length which includes a rebate for receiving the corresponding free end of the first insert part  1010   a.    
     The first insert part  1010   a  is made to be slightly flatter than required when in situ such that the free end is closer to the divider wall  1004  and inserting the second insert part  1010   b  urges the first part  1010   a  towards the leading edge so as to provide the biasing force for retaining the assembled insert  1010  in place. 
     In order to provide a correct fit, the insert parts  1010   a,b  are arranged to be held in an abutting relation with a resilient bias provided by one of the insert parts. The resilient bias in the case of the front cavity is provided by the fore insert part  1010   a  which is inserted after the sealing plate which is described above. The fore insert part may be oversized slightly with respect to the space in which it is designed to accommodate such that it must elastically deform during insertion. 
     The elastic deformation is such that the part is sufficiently stressed so as to provide the resilient bias between a wall of the cavity and sealing plate. Alternatively, the insert part may be made so as to be partly collapsible or compressible so that the shape of the part is altered to allow it to be inserted. In order to provide the collapsibility and compressibility, the insert part may be made to size for the cavity before being plastically deformed prior to insertion of the part. 
     The insert parts can incorporate rebates or other features to allow them to be secured in an abutting relation and to provide opposing surfaces for the retention of the parts via the resilient bias. Hence, as seen in  FIG. 10 b   , the sealing plate insert part  1010   b  in the front cavity  1001  and the free ends of the first and second main body parts in the rear cavity  1003  include rebates for receiving corresponding parts of abutting insert parts. Further, the rails of trip-strip insert parts  1020   c - f  include protuberant lips which engage with corresponding rebates in the main body portions. 
     It will be noted that the shell is constructed from a CMC material and as such has smooth outer and inner walls, principally due to the difficulties of forming discrete features in a CMC material. However, this may not always be the case, and the inserts are applicable to other non-CMC constructed shells. 
       FIGS. 11 a  and 11 b    provide another example in which the rear insert  1120  comprises three insert parts  1120   a - c . The first insert part  1120   b  is a V-shaped part having two plate-like members  1121   a  and  1121   b  which are joined at a hinged portion  1121   c.  The free ends of the members  1121   a,b  (or arms) are tapered from the first end to the second end so as to provide a smaller sectional area at the first end so that it can be manoeuvred more readily into the insertion aperture  1150 , and to provide a generally wedge shaped insert part. The second  1120   b  and third  1120   c  insert parts join along a mid-line of the sealing plate and form a wedge shaped part in unison which provides a chock for the first part  1120   a  when the insert parts  1120  are assembled into a complete insert. It will be appreciated that the second and third parts are inserted from the opposite end of the cavity through a second insertion aperture. 
     The V-shaped first insert part  1120   a  is fabricated such that the angle between the arms is greater than angle between the corresponding portions of the rear cavity. Thus, to insert the part, the arms are forceably moved together so as to elastically stress the hinge portion as it is passed through the insertion aperture. Once inside the cavity, the insert part can be pushed into the receiving portion  1160  with the resilient bias of the arms retaining the part in place. 
     The front cavity multi-part insert  1110  includes three parts  1110   a - c . Here, the first insert part  1110   a  extends from the divider wall toward the leading edge against the pressure surface of the front cavity  1101 . The second part  1110   b  abuts the free end of the first insert part  1110   a  which is local to the leading edge and extends around the suction surface toward the suction surface. The third insert  1110   c  is generally L shaped with rebates provided on the free ends of long and short members. The rebates provide a flange which resides on the inside of the free ends of the corresponding ends of the first and second insert parts. The arms are joined at a hinge portion. 
     The first  1110   a  and second  1110   b  insert parts are made to fit in a neutral or stress-free state within the front cavity  1101  whilst abutting the walls of the shell  1100 . The third L-shaped insert part is fabricated to have a larger angle than required such that the hinge portion elastically deformed upon insertion so as to provide a restoring force to bias against the free ends of the first and second insert parts against the wall of the cavity via the rebated portions. 
     A further example is shown in  FIGS. 12 a  and 12 b    which corresponds to the component described in  FIGS. 8 and 9  above, but with multiple insert parts in the front  1201  and rear cavities  1203 . Hence, the front  1210  and rear  1220  inserts each include two insert parts  1210   a,b ,  1220   a,b , having similar features to those described above in relation to  FIGS. 10 a  to 11 b   . In this instance, the front cavity  1201  has a first insert part  1210   a  which is inserted first and provides the resilient bias once the sealing part is inserted. The rear cavity  1203  has a first insert part  1220   a  which is inserted into the rear cavity via the insertion aperture  1250  along a first trajectory before being pushed rearward into the trailing edge which it is located in its corresponding receiving portion  1260 . The second insert  1220   b  provides the sealing plate and a portion of wall which defines a cooling chamber with the cavity wall. The wall is connected to the sealing plate via a hinge portion which provides the resilient bias for retaining the first insert part in place. 
     In addition to the above, it is possible in some embodiments that multiple insert parts can be fitted inside one another so that a single shell cavity includes an insert formed from two or more nested insert parts. Each insert shown in  FIGS. 3 and 4  seals its cavity, as well as providing formations to support the insert and guide cooling air around the inner surface of the shell. If two nested insert parts are used in a cavity, the outer of the two insert parts can provide the formations, and the inner of the two insert parts can be configured to balloon under the pressure of the inboard and/or outboard flows of cooling air to provide a sealing load. 
     While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. For example, the shell may be a metal shell rather than a CMC shell. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.