BLADE AND A METHOD OF MANUFACTURING A BLADE

Abstract

Disclosed herein is a method of manufacturing a double walled section of an aerofoil (701) of a gas turbine engine (10), the method comprising: forming a plurality of columns (402) on an outer surface of a first structure; forming a plurality of columns (402) on a surface of each of a plurality of parts of a second structure; and forming a double walled section of an aerofoil (701) by attaching the ends of columns (402) on each part of the second structure to the ends of columns (402) on the first structure such that the first structure is an inner wall (401) of the section of the aerofoil (701) and the second structure is an outer wall (501) of the section of the aerofoil (701).

Claims

1. A method of manufacturing a double walled section of an aerofoil of a gas turbine engine, the method comprising: forming a plurality of columns on an outer surface of a first structure; forming a plurality of columns on a surface of each of a plurality of parts of a second structure; and forming a double walled section of an aerofoil by attaching the ends of columns on each part of the second structure to the ends of columns on the first structure such that the first structure is an inner wall of the section of the aerofoil and the second structure is an outer wall of the section of the aerofoil.

2. The method as claimed in claim 1, wherein: forming a plurality of columns on the outer surface of the first structure comprises imprinting the columns in its outer surface; and/or for each of the plurality of parts of the second structure, forming a plurality of columns on the surface comprises imprinting the columns in the surface.

3. The method as claimed in claim 1, further comprising performing a smoothing operation that flattens the ends of each of the columns after the columns have been formed.

4. The method as claimed in claim 1, wherein attaching the ends of columns on each part of the second structure to the ends of columns on the first structure comprises diffusion bonding or brazing the ends of the columns together.

5. The method as claimed in claim 1, wherein forming a plurality of columns on the outer surface of the first structure is performed by a plurality of manufacturing operations with each manufacturing operation forming a plurality of parallel columns in a single region of the first structure; and/or wherein forming a plurality of columns on the surface of each part of the second structure is performed by one or more manufacturing operations with each manufacturing operation forming a plurality of parallel columns in a single region of the second structure.

6. The method as claimed in claim 1, wherein the section of the aerofoil is comprised by a turbine blade or a compressor blade, wherein the turbine blade or compressor blade comprises a leading-edge region, a mid-chord region and a trailing edge region; and the section of the aerofoil is the leading edge region and mid-chord region of the blade.

7. The method as claimed in claim 1, further comprising attaching each part of the second structure to one or more other parts of the second structure.

8. The method as claimed in claim 1, wherein the first structure comprises a plurality of parts; and the method comprises attaching each of the parts of the first structure to one or more other parts of the first structure.

9. The method as claimed in claim 1, wherein the first structure and/or one or more parts of the second structure are formed by casting.

10. The method as claimed in claim 1, wherein the first structure and/or one or more parts of the second structure are single crystal super alloys.

11. The method as claimed in claim 1, wherein the first structure is a walled structure with the walls providing an enclosed inner part of the section of the aerofoil and the second structure is a walled structure with the walls enclosing the first structure, wherein the method further comprises: forming one or more holes through the walls of the first structure; and forming one or more holes through the walls of the second structure, such that air in the inner part of the aerofoil can flow out of the aerofoil via the holes in the first structure and second structures.

12. The method as claimed in claim 11, wherein the positioning of the holes in the walls of the first structure is staggered with respect to the positioning of the holes in the walls of the second structure.

13. An aerofoil that is manufactured according to the method as claimed in claim 1.

14. A double walled section of an aerofoil of a gas turbine engine, the double walled section comprising: an inner wall with a plurality of columns protruding from its outer surface; an outer wall with a plurality of columns protruding from its inner surface; joints between the end surfaces of the columns on the inner wall and the end surfaces of the columns on the outer wall.

15. The double walled section as claimed in claim 14, wherein the joints are diffusion bonds or braised connections.

16. The double walled section as claimed in claim 15, wherein the columns are grouped to provide one or more interface regions; and within each of the interface regions, the joints between the ends of the columns are all in the same plane.

17. The double walled section as claimed in claim 16, wherein the section of the aerofoil is comprised by a turbine blade or a compressor blade.

18. The double walled section as claimed in claim 17, wherein the turbine blade or compressor blade has a leading-edge region, a mid-chord region and a trailing edge region; and the section of the aerofoil is the leading edge region and mid-chord region.

19. A gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; wherein one or more of the turbine and the compressor comprise one or more double walled sections of an aerofoil as claimed in claim 14.

20. The gas turbine engine as claimed in claim 19, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft; wherein one or more of the second turbine and second compressor comprise one or more double walled sections of an aerofoil as claimed in claim 14.

Description

DESCRIPTION OF THE DRAWINGS

[0055] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0056] FIG. 1 is a sectional side view of a gas turbine engine;

[0057] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0058] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0059] FIG. 4 is a cross-section of an inner wall of a blade after pedestals have been formed on the inner wall according to an embodiment;

[0060] FIG. 5 is a cross-section of both an inner wall and a plurality of parts of an outer wall according to an embodiment;

[0061] FIGS. 6A and 6B are cross-sections of a blade according to an embodiment;

[0062] FIG. 7 shows a blade according to an embodiment with a portion of the blade cut away; and

[0063] FIG. 8 is flowchart of a method according to an embodiment.

DETAILED DESCRIPTION

[0064] FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30. In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place.

[0065] The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0066] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

[0067] Note that the terms low pressure turbine and low pressure compressor as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the low pressure turbine and low pressure compressor referred to herein may alternatively be known as the intermediate pressure turbine and intermediate pressure compressor. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

[0068] The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in FIG. 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

[0069] The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

[0070] It will be appreciated that the arrangement shown in FIGS. 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 2.

[0071] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

[0072] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0073] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

[0074] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

[0075] The present disclosure provides an improved method of manufacture of a blade. The blades that are manufactured according to the present disclosure can be used as the blades of any of the components of the above-described gas turbine engine 10. The blades that are manufactured according to the present disclosure may be used as either high or low pressure turbine blades. However, the blades according to the present disclosure may be used in any of the other components of the gas turbine engine, such the first or second compressor.

[0076] In order to clearly present the context of the present disclosure, details of background technology are provided below.

[0077] An important property of blades in a gas turbine engine is how efficiently the blades can be cooled. Known techniques for cooling blades include transpiration cooling, as well as similar techniques such a effusion cooling and porous multi-wall cooling. Laminated multi-wall cooling schemes, in which a film is formed over at least part of the surface of the blade, are known to provide a practical implementation of these cooling schemes. The cooling schemes provide good cooling uniformity due to a series of highly distributed micro cooling channels. The heat transfer is also highly efficient as the cooling air is brought close to the outer surface of a blade.

[0078] Known multi-wall cooling blade designs with film cooling are disclosed in U.S. Pat. Nos. 6,514,042B2 and 5,702,232A. Both of these blades comprise a double wall configuration in the mid-chord region with radial feed passages on each side of the blade between the inner and outer wall. A problem with these known designs of blades is that they are difficult to manufacture. A method of manufacture is described in U.S. Pat. No. 5,640,767A. A partially-hollow blade support wall, that is formed by casting, with a desired outer profile has longitudinally-extending recesses, that are filled with channel fillers. A blade skin of a second material is then deposited, by methods comprising physical, chemical vapour deposition, thermal spraying and plating, onto the blade support wall such that the skin conforms and is metallurgically bonded to the blade support wall. The recesses' channel filler is removed and the combination of outer blade skin and blade support wall forms a double wall blade structure with the voids in the recess creating an integral internal chamber within the double walled structure. Another known manufacturing method of double walled blades is the Shell-and-Spar approach, as disclosed in Novikov et al, Creation of High Efficiency Turbine Cooled Blades With Structural Electron Beam Coatings, in collection of papers Electron Beam and Gas-Thermal Coatings, pages 87 to 97. This is similar to the manufacturing method described in U.S. Pat. No. 5,640,767A.

[0079] The known manufacturing techniques of double walled blades therefore comprise additive manufacturing. A problem with this manufacturing technique is that the layering processes introduce imperfections and it is difficult to construct surfaces with intricate details and without defects.

[0080] The present disclosure provides a new method of manufacturing double walled blades that avoids the problems associated with additive manufacturing processes. The blades may be effusion cooled aerofoils. The blades may be fabricated with a single crystal structure using known materials for blades. The blades may be made with intricate features, high surface quality and high mechanical strength.

[0081] The blades according to the present disclosure may be aerofoils that comprise a leading edge region, a mid-chord region and a trailing edge region. The leading edge region and mid-chord region may be constructed with double walls with the trailing edge region only having a single wall. The double wall is provided by an outer wall and an inner wall, which may alternatively be referred to as an outer skin and inner skin. The inner and outer walls are connected to each other by pedestals. A pedestal is a column. The columns can have a variety of shapes and may be linear with a substantially circular cross-section and a constant diameter along their length. The columns may alternatively be linear with a substantially square cross-section that is constant along their length. The pedestals protrude from corresponding surfaces of the inner and outer walls. The ends of the pedestals are flat and can therefore be joined together by, for example, diffusion bonding. The pedestals are effectively pin fins. The inner wall has holes through it that are referred to as impingement holes. The outer wall also has holes through it that are referred to as effusion holes. The impingement and effusion holes are positioned in a staggered arrangement with respect to each other. Accordingly, air can pass from an inner cavity of the blade, that is enclosed by the inner walls, through the impingement holes past the pedestals and out through the effusion holes. The air forms a transpiration-like film on at least part of the outer surface of the blade and is very effective at cooling the blade.

[0082] The present disclosure is described in more detail below with reference to FIGS. 4, 5, 6A, 6B and 7.

[0083] The inner wall 401 is initially manufactured without pedestals 402 in its surface. The pedestals 402 are then formed in its surface. The initial structure of the inner wall 401 may be manufactured through casting. The cast material may be directionally solidified or single crystal super alloy.

[0084] The pedestals 402 may be formed in the surface of the initial structure of the inner wall 401 by an imprinting process that removes material from the surface. The imprinting process may be milling/spark electrode discharge machining, EDM, using a graphite/tungsten tool. The tool may be configured with an exact negative of the desired pedestal pattern and may effectively be plunged into the surface of the initial structure of the inner wall 401 in order to form an array of a plurality of pedestals 402. The pedestals 402 are therefore part of the initial structure of the inner wall 401 and protrude from the surface of the inner wall 401 due to the imprinting process removing material around the pedestals 402 from the surface.

[0085] A plurality of imprinting processes are performed over the surface of the initial structure of the inner wall 401. FIG. 4 shows the inner wall 401 after the imprinting processes have been performed. The pedestals 402 formed in each imprinting process are all aligned in the same direction. In order for pedestals 402 to be formed over the entire surface of the inner wall 401, a plurality of imprinting processes may be performed from different directions. The number of directions that the pedestals 402 are aligned in may therefore the same as the number of different directions of the imprinting processes.

[0086] After the pedestals 402 have been formed, a smoothing operation may be performed in order to flatten the tops of the pedestals 402. This may also ensure that all of the pedestals 402 have the desired height. The smoothing operation may be, for example, a grinding operation.

[0087] The distance between two adjacent pedestals 402 on the inner wall 401 may be in the range 0.5 mm to 1.5 mm. If a pedestal 402 on the inner wall 401 has a substantially square cross-section that is constant along the length of the pedestal 402, the length of a side of the square cross-section of the pedestal 402 may be in the range 0.5 mm to 1.5 mm. The height of a pedestal 402 on the inner wall 401 after the smoothing operation may be in the range 0.5 mm to 1.5 mm.

[0088] The outer wall 501 may be initially manufactured as a structure with a plurality of separate parts. Each of the parts of the initial structure of the outer wall 501 may be initially manufactured without pedestals 402 in its surface. The pedestals 402 may then formed in a surface of each part. Each part may be manufactured through casting. The cast material may be directionally solidified or single crystal super alloy.

[0089] The pedestals 402 may be formed in the surface of each part of the initial structure of the outer wall 501 by an imprinting process that removes material from the surface. The imprinting process may be milling/spark electrode discharge machining, EDM, using a graphite/tungsten tool as described above for the formation of pedestals 402 on the outer surface of the inner wall 401. The tool may be configured with an exact negative of the desired pedestal pattern and may effectively be plunged into the surface in order to form an array of a plurality of pedestals 402. The pedestals 402 are therefore comprised by the part of the initial structure of the outer wall 501 and protrude from the surface of the part due to the imprinting process removing material around the pedestals 402 from the surface.

[0090] Each part of the initial structure of the outer wall 501 may have one or more imprinting processes performed on it. FIG. 5 shows the plurality parts of the initial structure of the outer wall 501 with pedestals 402 formed in a surface of each part. As described for the formation of pedestals 402 in the outer surface of the inner wall 401, the pedestals 402 formed in each imprinting process are all aligned in the same direction. Imprinting processes may be performed from different directions on the same part.

[0091] After the pedestals 402 have been formed in a surface of each part of the initial structure of the outer wall 501, a smoothing operation is performed in order to flatten the tops of the pedestals 402. This may also ensure that all of the pedestals 402 have the desired height. The smoothing operation may be, for example, a grinding operation.

[0092] The distance between two adjacent pedestals 402 on the outer wall 501 may be in the range 0.5 mm to 1.5 mm. If a pedestal 402 on the outer wall 501 has a substantially square cross-section that is constant along the length of the pedestal 402, the length of a side of the square cross-section of the pedestal 402 may be in the range 0.5 mm to 1.5 mm. The height of a pedestal 402 on the outer wall 501 after the smoothing operation may be in the range 0.5 mm to 1.5 mm.

[0093] On the inner wall 401, each group of adjacent pedestals 402 with flattened ends that are all in the same plane provides an attachment interface to the inner wall 401. On each part of the outer wall 501, each group of adjacent pedestals 402 with flattened ends that are all in the same plane provides an attachment interface to the part of the outer wall 501. Each attachment interface to the inner wall 401 has a corresponding attachment interface on a part of the outer wall 501. Each pedestal on the inner wall 401 is therefore arranged to be attached to a corresponding pedestal on a part of the outer wall 501.

[0094] FIGS. 6A and 6B show cross-sections of a blade 701 according to the present disclosure. FIG. 7 shows a blade 701 according to the present disclosure with a portion of the outer wall 501 cut away so as to show the pedestals 402. The trailing edge region of the blade 701 may comprise TE slots.

[0095] A blade 701 according to the present disclosure is formed by attaching each of the parts of the outer wall 501 to the inner wall 401. Each part of the outer wall 501 may be attached to the inner wall 401 by joining the end surfaces of all of the pedestals 402 in the one or more interfaces of the part of the outer wall 501 and the corresponding one or more interfaces on the inner wall 401. The end surfaces of the pedestals 402 may be joined together by diffusion bonding or brazing. This technique is possible since the end surfaces of each of the pedestals 402 being joined are all flat and parallel to each other. The joints in each of the interfaces may therefore all be in the same plane.

[0096] Adjacent parts of the outer wall 501 may also be directly attached to each other, such as by diffusion bonding or brazing, so that the parts of the outer wall 501 are no longer separated from each other.

[0097] As shown in FIGS. 6B and 7, there are impingement holes 601 through the inner wall 401 and effusion holes 602 through the outer wall 501. The initial structures of both the inner wall 401 and each part of the outer wall 501 may be made with these holes. For example, the initial structures may be cast with these holes. Alternatively, the holes may be formed in the initial structures of both the inner wall 401 and each part of the outer wall 501 after the initial structures have been manufactured.

[0098] Possible advantages of the above manufacturing technique may include the pedestals 402 all being integral with either the inner or outer wall 501. The attachment of each pedestal to the wall may therefore be very strong. Another advantage may be that the outer surface of the blade 701 does not need to be altered after each part of the initial structure of the outer surface has been manufactured. The features of outer surface may therefore be made with high precision.

[0099] As explained above, a plurality of machining operations are performed in order to form pedestals 402 in the inner wall 401. Each machining operation cuts out some of the surface in order to form an interface comprising a plurality of parallel columns. Due to the non-planar shape of the inner wall 401, adjacent interfaces may not be co-linear with each other and there may be an oblique or reflex angle between adjacent interfaces. The portion of the outer surface of the inner wall 401 between any two adjacent interfaces may therefore not have any material cut away by a machining operation. There may also be a corresponding portion on one or more parts of the outer wall 501 without pedestals 402. Since there may be no pedestals 402 in these portions between interfaces, the portions may not be cooled as much as the parts of the walls with pedestals 402 and, in particular on the outer wall 501, may become hot spots. If such hot spots are too large then they overheat and degrade operation. The size of the hot spots can be reduced by increasing the number of interfaces that are formed in the inner wall 401 so that the plurality of linear interfaces more closely approximate the shape of the inner wall 401. However, increasing the number of machining operations increases the cost of manufacture. There is therefore a trade off between the number of interfaces and the cost of manufacture.

[0100] The number, shape, size and position of all of the pedestals 402, effusion holes 602 and impingement holes 601 can be varied in order to achieve minimal aerodynamic losses and to maximise the cooling of the blade 701. The position of the effusion holes 602 and impingement holes 601 may be staggered with respect to each other.

[0101] Advantageously, the flow of air from inside the inner walls 401, through the impingement holes 601, past the pedestals 402 and through the effusion holes may result in a transpiration-like film being formed on the outer surface of the blade 701 and good cooling properties. Compared to known blades, the required cooling air consumption may be reduced and efficiency may be increased.

[0102] FIG. 8 is a flowchart of a method of manufacturing a double walled section of an aerofoil of a gas turbine engine according to the present disclosure.

[0103] In 801, the method begins.

[0104] In 802, a plurality of columns are formed on an outer surface of a first structure.

[0105] In 803, a plurality of columns are formed on a surface of each of a plurality of parts of a second structure.

[0106] In 804, a double walled section of an aerofoil is formed by attaching the ends of columns on each part of the second structure to the ends of columns on the first structure such that the first structure is an inner wall of the section of the aerofoil and the second structure is an outer wall of the section of the aerofoil.

[0107] In 805, the method ends.

[0108] Embodiments include a number of modifications and variations to the techniques as described above.

[0109] For example, in addition to the outer wall 501, the inner wall 401 may also be initially formed as a plurality of separate parts that are later attached together. The pedestals 402 and parts of the outer wall 501 may be joined together by other techniques, such as welding.

[0110] The imprinting processes may alternatively be performed by other techniques than EDM. For example, any of CNC milling, casting, abrasive jet machining, water jet machining, laser cutting or electronic beam machining may alternatively be used for the imprinting processes.

[0111] The end surfaces of the pedestals 402 may alternatively be joined together by other techniques than diffusion bonding or brazing. For example, friction welding or adhesive bonding may be used to join the end surfaces of the pedestals 402.

[0112] It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.