GAS TURBINE ENGINE COMPONENT

Abstract

Described is 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 multiple separate parts 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, the retention part engaging with a wall of the cavity and at least one other insert part so as to retain the assembled insert within the cavity.

Claims

1. A gas turbine engine component, comprising: a shell having a divider wall which divides an internal cavity of 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; an insert located within the front cavity or rear cavity, wherein the divider wall includes at least one aperture which provides fluid communication between the front and rear cavity, and the insert includes a sealing plate to prevent the flow of cooling air across the divider wall through the at least one aperture.

2. A gas turbine engine component as claimed in claim 1, further comprising a front cavity insert and a rear cavity insert, each of the front cavity insert and rear cavity insert including a sealing plate.

3. A gas turbine engine component as claimed in claim 1, wherein the shell comprises a ceramic matrix composite material.

4. A gas turbine engine component as claimed in claim 1, wherein the shell is an aerofoil.

5. A gas turbine engine component as claimed in claim 1, wherein the insert is a unitary body.

6. A gas turbine engine component as claimed claim 1, wherein the insert is resiliently biased between a wall of the shell and the divider wall so as to urge the sealing plate against the divider wall to aid sealing the apertures.

7. A gas turbine engine component as claimed in claim 1, wherein the insert includes two members joined at a hinge portion.

8. A gas turbine engine component as claimed in claim 7, wherein the hinge portion is sprung loaded to provide a resilient bias, wherein the resilient bias acts to urge the an insert part against one or more walls of the shell.

9. A gas turbine engine component as claimed in claim 1, wherein the insert includes: 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.

10. A gas turbine engine component as claimed in claim 1, wherein the insert includes heat transfer formations which support the insert within the shell and which guide the cooling air around the inner surface of the shell.

11. A gas turbine engine component as claimed in claim 1, wherein the shell includes exterior film cooling holes.

12. A gas turbine engine component as claimed in claim 1, wherein the insert includes trip strip formations which lie along the inner surface of the shell.

13. A gas turbine engine component as claimed in claim 4 wherein the insert defines 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.

14. A gas turbine engine component as claimed in claim 4, wherein the or each insert is a multi-part insert located within the respective cavity, wherein the multi-part insert comprises multiple separate parts assembled in an abutting relation with one another within the respective cavity to provide the multi-part insert, the multiple separate parts being separately insertable into the respective cavity; wherein the assembled multi-part insert includes at least one retention part, the retention part engaging with a wall of the respective cavity and at least one other insert part so as to retain the assembled multi-part insert within the respective cavity.

15. A gas turbine component as claimed in claim 14, wherein the retention part is configured to provide an interference fit with one or more other of the insert parts and or a wall of the respective cavity.

16. A gas turbine component as claimed in claim 15, wherein the retention part provides a resilient bias which acts to urge the retention part or another insert part against one or more walls of the shell.

17. A gas turbine component as claimed in claim 16, wherein the retention part includes at least two members connected by a hinge portion, wherein the hinge portion provides the resilient bias.

18. A gas turbine component as claimed in claim 17, wherein the respective cavity includes an insertion 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.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0066] Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which:

[0067] FIG. 1 shows an isometric view of a conventional HP stage cooled turbine;

[0068] FIG. 2 shows a longitudinal cross-section through a ducted fan gas turbine engine;

[0069] 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;

[0070] 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;

[0071] FIG. 5 shows a cross-sectional view of a second nozzle guide vane;

[0072] FIG. 6 shows a cross-sectional view of the nozzle guide vane of FIG. 5 with cooling flows indicated by arrows;

[0073] FIG. 7 shows a cross-sectional view of variant inserts for the nozzle guide vane of FIGS. 5 and 6; and

[0074] 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;

[0075] FIG. 9 shows a cross-sectional view of the aerofoil of FIG. 8 with the inserts fitted inside the shell; and

[0076] FIGS. 10a to 12b show various aerofoil embodiments having multi-part inserts according to the invention.

DETAILED DESCRIPTION AND FURTHER OPTIONAL FEATURES OF THE INVENTION

[0077] 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.

[0078] 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.

[0079] 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.

[0080] 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.

[0081] 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.

[0082] 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.

[0083] 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.

[0084] 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.

[0085] 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.

[0086] 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.

[0087] 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.

[0088] 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.

[0089] 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.

[0090] 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.

[0091] 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.

[0092] 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.

[0093] 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.

[0094] 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.

[0095] 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.

[0096] 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.

[0097] 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.

[0098] 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.

[0099] 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.

[0100] 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.

[0101] 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.

[0102] 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.

[0103] 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.

[0104] 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.

[0105] 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.

[0106] 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.

[0107] 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.

[0108] 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.

[0109] 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.

[0110] 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.

[0111] 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.

[0112] The inserts described above can be used instead of, or in combination with, sheet metal inserts.

[0113] 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.

[0114] FIGS. 10a, and 10b 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.

[0115] Thus, in FIGS. 10a and 10b 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. 10b where the first end 1000a and second end 1000b 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.

[0116] 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.

[0117] 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 1010a,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 1020a-f. The rear cavity insert 1020 is made up from two main body parts 1020a,b and several trip-strip insert parts 1020c-f which abut and engage the main body portions 1020b 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 1010a,b and the rear insert 1020 is formed from six insert parts 1020a-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.

[0118] To construct the vane with the assembled inserts 1010, 1020, the insert parts 1010a,b, 1020a-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.

[0119] 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.

[0120] 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.

[0121] 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.

[0122] In the described example of FIGS. 10a and 10b, a receiving portion 1060 can be taken to the rearmost portion of the rear cavity 1003 in which the first insert part 1010a is located. The insertion aperture 1050 can be taken to be at the first end 1000b of the aerofoil toward the divider wall 1004. Thus, the first insert part 1020a 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 1020a 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 1020c and 1020d 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 1020a prior to being loaded into the rear cavity 1003. Next, the second main body insert part 1020b and third trip strip 1020e 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 1020a in place. The final insert part is the fourth trip strip 1020f formation which is slid between a free end of a web of the first main insert part 1020a, 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.

[0123] It will be noted from FIG. 10b, 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.

[0124] The two insert parts 1010a,b of the front cavity 1001 include a curved member 1010a 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 1010b 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 1010a.

[0125] The first insert part 1010a 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 1010b urges the first part 1010a towards the leading edge so as to provide the biasing force for retaining the assembled insert 1010 in place.

[0126] In order to provide a correct fit, the insert parts 1010a,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 1010a 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.

[0127] 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.

[0128] 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. 10b, the sealing plate insert part 1010b 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 1020c-f include protuberant lips which engage with corresponding rebates in the main body portions.

[0129] 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.

[0130] FIGS. 11a and 11b provide another example in which the rear insert 1120 comprises three insert parts 1120 a-c. The first insert part 1120b is a V-shaped part having two plate-like members 1121a and 1121b which are joined at a hinged portion 1121c. The free ends of the members 1121a,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 1120b and third 1120c 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 1120a 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.

[0131] The V-shaped first insert part 1120a 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.

[0132] The front cavity multi-part insert 1110 includes three parts 1110a-c. Here, the first insert part 1110a extends from the divider wall toward the leading edge against the pressure surface of the front cavity 1101. The second part 1110b abuts the free end of the first insert part 1110a which is local to the leading edge and extends around the suction surface toward the suction surface. The third insert 1110c 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.

[0133] The first 1110a and second 1110b 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.

[0134] A further example is shown in FIGS. 12a and 12b 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 1210a,b, 1220a,b, having similar features to those described above in relation to FIGS. 10a to 11b. In this instance, the front cavity 1201 has a first insert part 1210a which is inserted first and provides the resilient bias once the sealing part is inserted. The rear cavity 1203 has a first insert part 1220a 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 1220b 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.

[0135] 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.

[0136] 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.