Turbine engine casing fabrication and heat treatment

Abstract

A method of manufacturing a conical component for an engine, the method comprising the steps of: forming a hollow truncated conical body from a plurality of segments of a first nickel-based alloy by w

Description

Background The present invention is concerned with a method for manufacturing a gas turbine engine component. Specifically, but not exclusively, an invention described herein relates to a method of manufacturing a casing for a low pressure turbine (LPT) of a gas turbine engine. Such components are known in the art as ‘LPT Casings’. It will be recognised from the teaching provided herein that the method may be applied to other components operating at the high temperatures associated with gas turbine machinery or the like. The turbine end of a gas turbine engine experiences extremely high temperatures during engine operation owing to the exhaust gases that leave the engine combustors and which cause the turbine (housed within the casing) to rotate. The specific operation of a gas turbine engine will be understood by a person skilled in the art, and will not therefore be described in detail herein. Temperatures during engine operation can reach in excess of 750 °C and may be sustained for long periods of time during flight. Thermal cycling is also a complication for the engine components as the engine heats and cools during operation and shut-down. Furthermore, the static and dynamic loads on the engine are substantial, in particular the load paths that are transmitted to the aircraft fuselage and main body from the engines. These unique combinations of thermal and loading conditions mean very specific materials need to be selected for engine components operating in modern gas turbine engines. As engine cores become smaller it is expected that engine operating temperatures will increase requiring new materials to be used. Conventionally, components such as the engine exhaust component may be manufactured using materials such as Haynes® 282® alloy. This is extremely resilient to temperatures and does not exhibit micro-cracking, fatigue or the like thereby making it inherently suitable for gas turbine applications. The materials are formed and welded together into the desired shape before being heat treated to create components which can satisfactorily meet the above referenced requirements for current engine needs. Conventionally, depending on material requirements, and in turn design requirements which dictate material requirements, components such as the LPT case are made from INCONEL® alloy 718 or Waspaloy® or a combination of these. A drawback with the conventional material selection for LPT cases is that the materials are extremely expensive and that the forging process itself is expensive. Waspaloy® exhibits superior properties and is used for the skins and hooks of typical LPT cases. The significant cost of Waspaloy® and particularly of forging makes the overall manufacturing cost of an LPT case very high. The inventors have established that Waspaloy® can be prone to Strain Age Cracking with a narrow forgability window which makes it very challenging to make large parts crack free. Additionally, since fine features have to be built in; the buy to fly ratio can be close to 10 making it very unsustainable and a poor business case. The present inventors have established an alternative and counterintuitive approach to LPT manufacture in which a more cost effective Ni-Cr-Co-Mo-AI-Ti alloy may be used. For example, a Ni-20Cr-10Co-8.5Mo-2.ITi-l.5AI alloy may be used which can be purchased using the trade name Haynes® 282® alloy or an alloy similar thereto can be used which matches the forging properties of Waspaloy®. The specific Haynes® 282® alloy has composition ranges as shown in the below table on page 4 of this application. Specifically the inventors have devised a method of heat treatment that can dramatically reduce manufacturing costs through a change in material whilst still maintaining the performance characteristics needed for LPT (and other high temperatures) applications. Summary of the Invention Aspects of the invention are set out in the accompanying claims. Viewed from a first aspect there is provided a method of heat treatment for a component comprising a first portion formed of a first nickel-based alloy. The method comprises: heating the component to a temperature from 920°C to 1120°C for less than about 0.75 hours. The method may additionally comprise, after the heating, performing carbide stabilisation by heating the component to a temperature from 825°C to 843°C for about 4 hours. Viewed from a second aspect there is provided a method of heat treatment for a component comprising a first portion formed of a first nickel-based alloy. The method comprises performing carbide stabilisation by heating the component to a temperature from 825°C to 843°C for about 4 hours. The method may additionally comprise, before performing carbide stabilisation, heating the component to a temperature from 920°C to 1120°C for less than about 0.75 hours. Thus, according to these two aspects a heat treatment method is provided which enables components to be produced which are suitable for use in high temperature operations and during thermal cycling. A material is therefore provided which can be used for turbine casings to match a material such as Waspaloy® forging properties that currently serves for both skin and hooks (inwardly extending projections) for a fabricated turbine case. Specifically, an invention described herein allows a component to be provided where the properties of Tensile strength (including Ultimate Tensile Strength, Yield Strength and Elongation), the Modulus of Toughness (or Containment Factor), the low-cycle fatigue (LCF) and creep of the material are optimised. Furthermore, good weldability of a non-forged product with respect to a skin can also be realised. Creep and LCF are critical properties for the inwardly extending projections or ‘hooks’. The first nickel-based alloy may be a Ni-Cr-Co-Mo-AI-Ti alloy. The Ni-Cr-Co-Mo-AI-Ti alloy may have the following composition: Ni-20Cr-10Co-8.5Mo-2.ITi-l.5AI. Alternatively, the alloy may have composition ranges as shown in the below table: Element First nickel-based alloy compositions Co 9-11 Fe 1.5 max Cr 18.5-20.5 Mo 8-9 Al 1.38-1.65 Ti 1.9-2.3 C 0.04-0.08 Mn 0.3 max Si 0.15 max B 0.003-0.010 Zr 0.02 max P 0.015 max Cu 0.1 max Pb 1.5 max Se - Bi - S 0.015 max Ni Remainder Although in this aspect the carbide stabilisation is performed for about 4 hours, it may be performed for longer or shorter than this. For example, this step may be carried out for between 3 and 5 hours or between 3.5 and 4.5 hours. During the heating step, the component may be heated to a temperature of from 970°C to 1070°C and preferably approximately 1020°C for less than 0.75 hours and optionally less than about 0.5 hours. The carbide stabilisation may be performed by heating the component to a temperature from 825°C to 841°C and preferably from 825°C to 840°C and more preferably from 830°C to 840°C and still more preferably from 835°C to 840°C for about 4 hours. The method may additionally comprise performing a welding operation to weld a second portion to the first portion before heating the component, wherein the heating step may be a post weld heat treatment. The second portion may be formed of a second nickel-based alloy. For example, the nickel based alloy may be Waspaloy® or another similar nickel-based alloy. The second nickel-based alloy may have composition ranges as shown in the below table: Element Second nickel-based alloy compositions Co 12-15 Fe 2 max Cr 18-21 Mo 3.5-5 Al 1.2-1.6 Ti 2.75-3.25 C 0.02-0.10 Mn 0.1 max Si 0.15 max B 0.003-0.10 Zr 0.02-0.08 P 0.015 max Cu 0.1 max Pb 0.0005 max Se 0.0003 max Bi 0.00003 max S 0.015 max Ni Reminder The method may additionally comprise performing a first stress relief step for the component through heat treatment by heating the first portion of the component to between 1065° C and 1093° C for about 1 hour; and performing a machining or manufacturing operation to the first portion of the component and a second stress relief step through solution heat treatment by heating the component to between 1065° C and 1093° C for about 1 hour after the machining or manufacturing operation. This step may be carried out for longer or shorter than about 1 hour. For example, it may be carried out for between about 0.5 and 1.5 hours. The method may additionally comprise performing an additive manufacturing process to the component using a powder or wire formed of the first nickel-based alloy and performing a solution heat treatment process by heating the body to a temperature from 1005°C to 1205°C for about 1 hour. This step may be carried out for longer or shorter than 1 hour. For example, it may be carried out for between about 0.5 and 1.5 hours. The body may be heated to a temperature from 1055°C to 1155°C and preferably approximately 1105°C for about 1 hour. This step may be carried out for longer or shorter than 1 hour. For example, it may be carried out for between about 0.5 and 1.5 hours. The method may additionally comprise performing carbide stabilisation of the first nickel-based alloy before the heating step by heating the component to a temperature from 910°C to 1110°C for 2 hours. This step may be carried out for longer or shorter than about 2 hours. For example, it may be carried out for between about 1 and 3 or 1.5 and about 2.5 hours. The component may be heated to a temperature of from 960°C to 1060°C and preferably approximately 1010°C for about 2 hours. This step may be carried out for longer or shorter than 2 hours. For example, it may be carried out for between about 1 and 3 or about 1.5 and 2.5 hours. The method may additionally comprise after performing carbide stabilisation by heating the component to a temperature from 825°C to 843°C for approximately 4 hours performing precipitation by heating the component to a temperature from 660°C to 860°C for 16 hours. This step may be carried out for longer or shorter than about 16 hours. For example, it may be carried out for between about 15 and 17 or about 15.5 and 16.5 hours. During precipitation the component may be heated to a temperature of from 710°C to 810°C and preferably approximately 760°C for about 16 hours. This step may be carried out for longer or shorter than 16 hours. For example, it may be carried out for between about 15 and 17 or about 15.5 and 16.5 hours. The component may be a conical component for a gas turbine engine formed of a hollow truncated conical body and wherein the second portion may be a concentric flange configured to be added to one or both ends of the hollow truncated conical body. Viewed from a third aspect there is provided a gas turbine case formed according to a method as set out in the first or second aspects of the invention. The gas turbine case may be a low pressure turbine case of a gas turbine engine. Viewed from a fourth aspect there is provided a gas turbine engine comprising a gas turbine case as set out in the second aspect of the invention. Viewed from a fifth aspect there is provided a method of manufacturing a conical component for an engine, the method comprising the steps of: (A) Forming a hollow truncated conical body from a plurality of segments of a first nickel-based alloy by welding a plurality of segments of the first nickel-based alloy together along their edges to form the conical body, wherein the weld lines extend along an axial direction of the conical component. Advantageously this allows plate material to be used to form the desired component. The term ‘edges’ is intended to mean the straight sides of the segments before the segments are welded together. The segments may be formed of a rolled plate of the first nickel-based alloy which have each been formed into a predetermined curved profile such that on alignment the conical body is defined. Precise conical shapes can thereby be formed from the plate stock material allowing a component to be formed for aero-engine applications, and other applications such as Industrial Gas Turbines. The plurality of segments may comprise at least two segments, and preferably three segments. It is desirable to use as few segments as possible to avoid additional joins. However, there are difficulties associated with forming conical parts of more than 180 degrees. Forming the conical component from three segments can be achieved with the available plate material whilst minimising the number of joins required. In other examples, up to five to ten segments may be used. The method may additionally comprise: (B) Performing a first stress relief step for the hollow truncated conical body through heat treatment; (C) Forming a plurality of inwardly extending projections extending from an inner surface of the hollow truncated conical body by an additive manufacture process; and (D) Performing a second stress relief step through solution heat treatment. The first stress relief step may be performed by heating the truncated conical body to between 1065° C and 1093° C for about 1 hour. This step may be carried out for longer or shorter than about 1 hour. For example, it may be carried out for between about 0.5 and 1.5 hours. The use of heat treatment improves the properties of the weld lines between the plurality of segments. The inventors have established that these temperatures and durations advantageously improve the properties of the component as referenced above. The step (B) of performing the first stress relief step may be followed by the steps of: (B2) an expansion step in which the hollow truncated conical body is expanded radially outwards until a predetermined radius of the conical body is reached; and (B3) a second stress relief step of the expanded hollow truncated conical body through heat treatment. The step of performing the second stress relief step of the expanded hollow truncated conical body may be performed by heating the body to between 1065° C and 1093° C for at least about 1 hour. This step may be carried out for longer or shorter than at least 1 hour. For example, it may be carried out for at least between about 0.5 and 1.5 hours. The plurality of inwardly extending projections extending from the inner surface of the truncated conical body may be formed by an additive manufacture process to form a series of concentric circumferentially and inwardly extending surfaces. The additive manufacturing process may be a laser metal deposition process using a wire or powder formed of the first nickel-based alloy. The additive manufacture process may be selected from: Direct metal laser sintering (DMLS), Electron beam melting (EBM), Selective laser melting (SLM), Selective laser sintering (SLS), Direct metal wire deposition or Direct metal powder deposition. Other additive manufacture processes may also advantageously be used. The method may additionally comprise solution heat treating the hollow truncated conical body after the inwardly extending projections have been formed by heating the body to a temperature from 1005°C to 1205°C for at least about 1 hour. This step may be carried out for longer or shorter than 1 hour. For example, it may be carried out for at least between about 0.5 and 1.5 hours. The body may be heated to a temperature from 1055°C to 1155°C and preferably approximately 1105°C for 1 hour. This step may be carried out for longer or shorter than about 1 hour. For example, it may be carried out for between about 0.5 and 1.5 hours. The method may additionally or alternatively comprise the step of performing carbide stabilisation of the first nickel-based alloy by heating the hollow truncated conical body to a temperature from 910°C to 1110°C for at least about 2 hours. This step may be carried out for longer or shorter than 2 hours. For example, it may be carried out for between about 1 and 3 or about 1.5 and 2.5 hours. The component may be heated to a temperature of from 960°C to 1060°C and preferably approximately 1010°C for 2 hours. The method may additionally comprise welding a concentric flange to one or both ends of the hollow truncated conical body wherein the concentric flange may be formed of a second nickel-based alloy to form a welded connection between the first nickel-based alloy and the second nickel-based alloy. Thus, the first nickel-based alloy and second nickel-based alloy may advantageously be structurally joined together. The second nickel-based alloy may be Waspaloy® or another similar nickel-based alloy. The method may additionally comprise the step of post weld heat treating the welded connection of the first nickel-based alloy and the second nickel-based alloy by heating the hollow truncated conical body to a temperature from 920°C to 1120°C and preferably from 970°C to 1070°C and more preferably a temperature of approximately 1020°C for less than about 1 hour. Alternatively, this step may be carried out for less than 0.75 hours. This step may be carried out for longer or shorter than 0.75 hours. For example, it may be carried out for between 0.25 and 1.25 hours. The method may additionally comprise a second step of carbide stabilisation by heating the hollow truncated conical body to between 825°C to 830°C for about 4 hours. This step may be performed for longer or shorter than about 4 hours. For example, this step may be carried out for between about 3 and 5 hours or between about 3.5 and 4.5 hours. The carbide stabilisation may be performed by heating the component to a temperature from 825°C to 843°C and preferably from 825°C to 841 °C and preferably from 825°C to 840°C and more preferably from 830°C to 840°C and still more preferably from 835°C to 840°C for about 4 hours. The method may additionally comprise the step of performing precipitation by heating the hollow truncated conical body to a temperature from 660°C to 860°C for up to about 16 hours. This step may be carried out for longer or shorter than about 16 hours. For example, it may be carried out for between about 15 and 17 or about 15.5 and 16.5 hours. During precipitation the component may be heated to a temperature of from 710°C to 810°C and preferably approximately 760°C for 16 hours. Viewed from a sixth aspect there is provided a conical component for a gas turbine engine, the conical component comprising a hollow truncated conical body formed from a plurality of segments, wherein the segments are welded together along weld lines to form the conical body, wherein the weld lines extend along an axial direction of the conical component. The component may have a radius at its widest part of between 500 and 1200 mm, preferably between 800 and 1200 mm. The component may have an axial length between 350 and 900 mm, preferably between 350 and 750 mm, and more preferably between 450 and 700 mm. In other examples, the component may be smaller or larger than this. The component may be a low pressure turbine case. It will be appreciated that the specific temperatures and times are approximate. For example, temperatures may be altered by + / -14°C and time periods may be varied by up to 10%. Thus, according to an invention described herein a material, process and heat treatment combination for casings to match Waspaloy® forging properties that currently serves for both Skin and hooks in fabricated turbine cases can be realised. Specifically, the properties of interest are Tensile (all:- Ultimate Tensile Strength, Yield Strength and Elongation) for Modulus of Toughness or Containment Factor; strength additionally is critical for payload; LCF is very important, so is weldability for non-forged product consideration with respect to skin. Creep and LCF are critical properties for the inwardly extending projections or ‘hooks’. The target requirement for each of these attributes over range of temperatures of operation i.e. from room temperature (RT) to at least 760°C with high stresses up to 517 MPa (75ksi). The invention disclosed herein results in an improvement of containment factor of up to 85% for a Room Temperature tensile test and 48% for 704°C tensile test with ASTM E8M IASTM E21 Test standard. The inventors have established that the methods and process described here allow a material, process and heat treatment combination for casings to match Waspaloy® forging properties that currently serves for both Skin and hooks in fabricated cases for existing gas turbine engines can be provided. Such a process has not previously been provided to meet aeroengine requirements. Drawings Aspects of the invention will now be described, by way of example only, with reference to the accompanying figures in which: Figure 1 shows the main components of a gas turbine engine; Figure 2 shows a turbine casing; Figure 3 shows an alternative view of a low pressure turbine casing; Figure 4 shows 3 segments of the component formed of the Ni-Cr-Co-Mo-AI-Ti alloy rolled material for forming the conical turbine casing profile; Figure 5 shows the conical casing shape once the 3 segments of Figure 4 have been welded together; Figures 6 shows schematically how the conical body may be formed into a complex geometrical turbine casing; Figures 7A to 7C illustrate the internal conical body structure post AM processing; Figures 8A and 8B show the addition of welded nickel-based alloy flanges to the ends of the conical body; Figure 9A illustrates schematically the steps of the process described herein; Figure 9B shows the corresponding metallurgical steps that provides the improvements in performance; Figure 9C shows the specific temperature and timing cycles that should be applied at each of the steps shown in figure 9B; Figure 10 shows a graph indicating how coefficient of thermal expansion varies with temperature. While the invention is susceptible to various modifications and alternative forms, specific embodiments are shown by way of example in the drawings and are herein described in detail. It should be understood however that the drawings and detailed description attached hereto are not intended to limit the invention to the particular form disclosed but rather the intention is to cover all modifications, equivalents and alternatives falling within the spirit and scope of the claimed invention. Any reference to prior art documents in this specification is not to be considered an admission that such prior art is widely known or forms part of the common general knowledge in the field. As used in this specification, the words “comprises”, “comprising”, and similar words, are not to be interpreted in an exclusive or exhaustive sense. In other words, they are intended to mean “including, but not limited to”. The invention is further described with reference to the following examples. It will be appreciated that the invention as claimed is not intended to be limited in any way by these examples. It will also be recognised that the invention covers not only individual embodiments but also combination of the embodiments described herein. The various embodiments described herein are presented only to assist in understanding and teaching the claimed features. These embodiments are provided as a representative sample of embodiments only and are not exhaustive and / or exclusive. It is to be understood that advantages, embodiments, examples, functions, features, structures, and / or other aspects described herein are not to be considered limitations on the scope of the invention as defined by the claims or limitations on equivalents to the claims, and that other embodiments may be utilised and modifications may be made without departing from the spirit and scope of the claimed invention. Various embodiments of the invention may suitably comprise, consist of, or consist essentially of, appropriate combinations of the disclosed elements, components, features, parts, steps, means, etc, other than those specifically described herein. In addition, this disclosure may include other inventions not presently claimed, but which may be claimed in future. It will be recognised that the features of the aspects of the invention(s) described herein can conveniently and interchangeably be used in any suitable combination Detailed Description Figure 1 shows the main components of a gas turbine engine 1. The operation of a gas turbine engine will be understood by a person skilled in the art and will not therefore be explained herein in detail. The engine 1 comprises a series of fan blades 2 which are driven by the central shaft 3. The fan blades 2 provide the primary driving thrust of the engine 1. Air is compressed first by the low pressure compressor 4 and then the high pressure compressor 5. Compressed air is released from the aft end 6 of the high pressure compressor 5 and into the combustors 7 where fuel is introduced and combustion occurs. Exhaust gas is then driven through the turbines, which are contained with the turbine casing 8 before leaving the rear of the engine through exhaust 9. The turbines are arranged as a high pressure turbine and a downstream or aft low pressure turbine. The turbine structures are extremely hot in operation owing to the high temperature exhaust gas that is driven through them towards the exhaust 9. As shown in figure 1, the general outer shape of the turbine casing 8 is a hollow conical body into which the turbine blades and central shaft are located. Figure 2 illustrates turbine casing 8 in isolation. As shown, the turbine casing 8, in this case a low pressure turbine casing 8, comprises a hollow body which is generally conical is shape. The inner surface of the casing body comprises a plurality of concentric rings which are known in the art as ‘hooks’ 10. The hooks are used to locate the turbine blades and vanes within the turbine structure. Conventionally the hooks may be formed by machining. However, in the invention and teaching described herein the hooks, or radially inwardly extending projections, are formed using an additive manufacturing technique. This is described in more detail below. Figure 3 shows an alternative view of a low pressure turbine casing 8 as described herein illustrating the projections of hooks 10. The casing 8 is formed of a plurality of segments. For example, the casing 8 may be formed of three segments 11 A, 11B and 11C as shown in Figure 4. Each segment is formed from a plate formed of a first nickel-based alloy. For example, a Ni-20Cr-10Co-8.5Mo-2.ITi-l.5AI alloy which can be purchased using the trade name Haynes® 282® alloy (UNS N07208) or an alloy similar thereto may be used. Haynes® 282® alloy is a material specifically designed super alloy for high temperature applications such as gas turbine engines and will be well known to the person skilled in the art. According to an invention described herein sections of plate of a first nickel-based alloy may be cut and rolled into segments with a predetermined profile as show in figure 4. The profiles are such that when brought into abutment with each other, the segments form the hollow conical profile of the turbine casing 8. Figure 4 shows 3 segments but it will be recognised that more or less segments may be used to form the casing profile. Figure 5 shows the conical casing shape once the 3 segments have been welded together. As shown, the segments 11A, 11B and 11C have been welded together along weld lines 12A, 12B and 12C to form the conical casing shape. Any suitable welding operation may be used. In one example the segment formed of a first nickel-based alloy may be welded together by a welding process such as an electron beam weld (EBW). It will be recognised from Figure 5 that the conical profile is geometrically quite simple and that the geometry of the desired turbine casing 8 (as illustrated in figures 1 and 2) is far more complex. This additional geometric complexity is achieved by expanding the conical profile radially outwards using suitable tooling, for example, a hydraulic heated mandrel. This is illustrated schematically in figure 6 where, using a mandrel, opposing forces F+, F-(which may be rotated around the central axis of the conical body) can create the desired radial displacement and profile along the axis of rotation of the body. The radial profile 13 of the casing 8 can be adapted to follow the desired inner profile of the turbine blades and stators within the casing 8. It will be recognised that other manufacturing techniques may be used to create the desired geometry. Once the desired outer casing geometry has been achieved the inner surface of the conical body can be adapted to create the desired radially inwardly extending projections or hooks for the internal components of the turbine casing 8. The hooks are deposited with Direct Energy Deposition (DED), for example using DED wire. In other examples, DED powder can be used. According to an invention described herein these features can be conveniently formed using an additive manufacturing (AM) technique. This will be described with reference to figure 7A to 7C which illustrate the internal conical body structure post AM processing. Figure 7A shows the expanded conical turbine casing 8 after the additive manufacturing process. AM techniques include a variety of processes in which geometries are built up on a substrate structure. In this case the inner surface of the conical body. AM technique are very well known in the art and include, amongst others: Direct metal laser sintering (DMLS) Electron beam melting (EBM) Selective laser melting (SLM) Selective laser sintering (SLS) Direct metal wire deposition (MLD) Direct metal powder deposition (MPD) According to an invention described herein the inner geometry of the conical body can be formed using metal laser deposition. Specifically the inner geometry can be creating using powder formed of a first nickel-based alloy i.e. the same material as the segments. This provide a homogeneous inward projection from the inner surface of the conical body. The specifics of the laser metal deposition process will be well understood in the art. The plate formed of the first nickel-based alloy in combination with hooks deposited with Additive Manufacturing process (DED) from Ni-Cr-Co-Mo-AI-Ti alloy powder (and or wire deposition) has been identified by the inventors as providing for the manufacture of an affordable and viable product. Advantageously since the volume fraction of the strengthening phase (y - that contributes most to the strength among other factors) is lower in the first nickel-based alloy compared to Waspaloy®; efforts are made towards identifying a window of opportunity with regards to selection of Heat Treatment to bring out the best of the first nickel-based alloy Plate properties (Tensile, LCF and weldability) along with the Creep and LCF in the hooks (DED of the first nickel-based alloy) and forged flanges formed of a second nickel-based alloy. Grain size of ASTM 4.5 or finer is selected in delivery condition to sustain similar or finer grain size after the full manufacturing cycle. Returning to figure 7A the inner geometry of the conical body is shown. Figure 7B shows a cut away section of the conical body with the inwardly extending projections 15 extending radially inward from the inner surface of the conical body. Figure 7C shows an expanded view of the inwardly extending projections or ‘hooks’ 15 shown in figure 7B. As shown the AM process has been used to build the projections using the powder material formed of the first nickel-based alloy creating a homogeneous conical body and projection. Figures 8A and 8B show the final stage of forming the turbine casing 8 with the welding of a flange 16 formed of the second nickel-based alloy to either end of the conical body. The inventor has established that for each of the steps of the method described herein, a specific sequence of heat treatment, carbide stabilisation and precipitation step can advantageously be used to obtain optimal material performance using cost effective materials. Furthermore, the specific parameters the inventor has identified are inconsistent with standard industry parameters and are thus counterintuitive in their selection. These will now be described with reference to figures 9A to 9C. The steps of the process are shown schematically in figure 9A from the forming of the segments formed of the first nickel-based alloy to the final machining. Figure 9B shows the corresponding metallurgical steps that provides the improvements in performance. Finally, Figure 9C shows the specific temperature and timing cycles that should be applied at each of the steps shown in figure 9B. It will be recognised by the person skilled in the art that at each step the conical body is positioned in autoclave or oven of the type known in the art with the autoclave / oven set to the parameters defined in figure 9C. The inventor has established that the Solution temperature close to or lower than a delivery condition (-1121 °C) has maximum influence on grain growth. In realistic conditions of applied stress, stress relief / annealing helps to bring back the material in a relaxed state. Specifically, the inventor has established that the grain size of a plate is a function of the solution temperature and thus the grain size increases with solution temperature. According to an invention described herein, weldability is important and thus, LCF is a critical property. These improve with low grain size. In addition, grain size can also improve strength (particularly Yield Strength) to a limited extent. Strength also has an influence on the containment which is an important consideration for aero-engine components and gas turbine engines. The inventor has also established that based on the manufacturing needs, Stress Relief operation may be required after cold forming i.e. stress relief 1. The discussion above explains the formation of the conical turbine case body and the formation steps. The discussion below explains specifics of the metallurgy. Specifically, as well as the above, the inventor has also established that deviation from the well established AMS (Aerospace Material Specification) guidelines actually improves the performance of the turbine cases, as set out below. As described above, the hooks are deposited additively using, for example, laser metal deposition. According to the invention this is followed by a non - AMS 5951 Solution Heat Treatment i.e. 1105°C instead of the normal 1121-1177°C. This may then be followed by a standard Carbide Stabilization step for first nickel-based alloy. Next, a post weld Heat Treatment is required after the second nickel-based alloy flanges have been welded to the turbine casing 8. This is followed by Carbide stabilization of the second nickel-based alloy and precipitation step of the second nickel-based alloy, as shown in figures 9B and 9C. Importantly, AMS 5951 does not have a step similar to Carbide precipitation of the second nickel-based alloy according to an invention described herein i.e. between 825°C to 843°C for 4hours - hence its effect (beneficial or deleterious) on the first nickel-based alloy is surprising but significant. Typically Superalloy Heat Treatment has 3 Steps: Solution Treatment followed by Carbide Stabilization and then Ageing. In one example Carbide Stabilization of the first nickel-based alloy is 1010 °C for 2h followed by Ageing at 788 °C for 8h. It has been established that any additional step in between may worsen the Tensile and other properties. Furthermore, Haynes® 282® alloy precipitation is normally 788°C for 8h. According to the invention described herein the precipitation step is instead 760°C for 16h which is again a deviation from accepted processes. The lower than standard AMS 5951 Solution step (1105°C) with post weld stress relief (close to the Carbide Stabilization) and 760°C 16h precipitation brings out an improvement in strength (both yield and Ultimate) as well as retaining finer grains in the structure. However, this has the effect of reducing elongation hence reducing the containment factor significantly. The Carbide stabilization step of the first nickel-based alloy actually improves thin carbide formation in grain boundaries of the first nickel-based alloy, hence further improving the strength in the first nickel-based alloy. Still further, the diffusion hence intra-grain composition aided by this between 825°C to 843°C for approximately 4 hour step is such that elongation is improved significantly. This enormous improvement in elongation helps in otherwise unexpected improvement in containment factor. According to the method described herein, whatever the form of the first nickel-based alloy material which is to be used to form the turbine cases, be it Plate, Sheet, Forging, bar or Cast; in applications that require the best properties of the first nickel-based alloy at higher temperature of service (up to 760°C) - it has been established according to an invention described herein that it can be performed with non AMS specified Solution Heat Treatment, specifically ~ 15-20°C below recommended. This can be followed by: Standard Carbide Stabilization: heating component to 1010°C for 2 hours Post weld stress relief: heating component to 1020°C for up to 0.75 hours Carbide Stabilization Step of second nickel-based alloy: heating component to 825-843°C for 4 hours Precipitation: heating component to 760°C for 16 hours or to 788°C for 8 hours Figure 10 shows how coefficient of thermal expansion (CTE) varies with temperature. The top line, ‘Desired’ shows how CTE varies with temperature for a second nickel-based alloy such as Waspaloy. The bottom line, ‘Standard’ shows how CTE varies with temperature for a Ni-Cr-Co-Mo-AI-Ti alloy such as Haynes® 282® alloy. The heat treatment method according to the present invention achieves a CTE similar to that shown by the ‘Desired’ line. CLAUSES 1. A method of heat treatment for a component comprising a first portion formed of a first nickel-based alloy, the method comprising: Heating the component to a temperature from 920°C to 1120°C for less than about 0.75 hours. 2. A method as set out in clause 1, additionally comprising after the heating, performing carbide stabilisation by heating the component to a temperature from 825°C to 843°C for about 4 hours. 3. A method of heat treatment for a component comprising a first portion formed of a first nickel-based alloy, the method comprising: Performing carbide stabilisation by heating the component to a temperature from 825°C to 843°C for about 4 hours. 4. A method as set out in clause 3, additionally comprising, before performing carbide stabilisation, heating the component to a temperature from 920°C to 1120°C for less than about 0.75 hours. 5. A method as set out in clause 1,2 or 4, wherein during the heating step, the component is heated to a temperature of from 970°C to 1070°C and preferably approximately 1020°C for less than about 0.75 hours and optionally less than about 0.5 hours. 6. A method as set out in clause 2, 3 or 4, wherein the carbide stabilisation is performed by heating the component to a temperature from 825°C to 841 °C and preferably from 825°C to 840°C and more preferably from 830°C to 840°C and still more preferably from 835°C to 840°C for about 4 hours. 7. A method as set out in clause 1, 2 or 4, additionally comprising: Performing a welding operation to weld a second portion to the first portion before heating the component, wherein the heating step is a post weld heat treatment. 8. A method as set out in clause 7, wherein the second portion is formed of a second nickel- based alloy. 9. A method as set out in any of clauses 1 to 8, additionally comprising: Performing a first stress relief step for the component through heat treatment by heating the first portion of the component to between 1065° C and 1093° C for about 1 hour; and Performing a machining or manufacturing operation to the first portion of the component and a second stress relief step through solution heat treatment by heating the component to between 1065° C and 1093° C for about 1 hour after the machining or manufacturing operation. 10. A method as set out in any of clauses 1 to 9, additionally comprising: Performing an additive manufacturing process to the component using a powder or wire formed of the first nickel-based alloy and performing a solution heat treatment process by heating the body to a temperature from 1005°C to 1205°C for about 1 hour. 11. A method as set out in clause 10, wherein the body is heated to a temperature from 1055°C to 1155°C and preferably approximately 1105°C for about 1 hour. 12. A method as set out in any of clauses 1 to 11, additionally comprising: Performing carbide stabilisation of the first nickel-based alloy before the heating step by heating the component to a temperature from 910°C to 1110°C for about 2 hours. 13. A method as set out in clause 12, wherein the component is heated to a temperature of from 960°C to 1060°C and preferably approximately 1010°C for about 2 hours. 14. A method as set out in any of clauses 2 to 13, additionally comprising after performing carbide stabilisation by heating the component to a temperature from 825°C to 843°C for about 4 hours: Performing precipitation by heating the component to a temperature from 660°C to 860°C for about 16 hours. 15. A method as set out in clause 14, wherein during precipitation the component is heated to a temperature of from 710°C to 810°C and preferably approximately 760°C for about 16 hours. 16. A method as set out in clause 7 or 8, wherein the component is a conical component for a gas turbine engine formed of a hollow truncated conical body. 17. A method as set out in clause 16 wherein the second portion is a concentric flange configured to be added to one or both ends of the hollow truncated conical body. 18. A gas turbine case formed according to a method as set out in any of clauses 1 to 17. 19. A gas turbine case as set out in clause 18, wherein the turbine case is a low pressure turbine case of a gas turbine engine. 20. A gas turbine engine comprising a gas turbine case as set out in clause 18 or 19. 21. A conical component for a gas turbine engine, the conical component comprising a hollow truncated conical body formed from a plurality of segments, wherein the segments are welded together along weld lines to form the conical body, wherein the weld lines extend along an axial direction of the conical component. 22. The component of clause 21, wherein the component is a low pressure turbine case.

Claims

1. A method of manufacturing a conical component for an engine, the method comprising the steps of:5 (A) Forming a hollow truncated conical body from a plurality of segments of a first nickelbased alloy by welding the plurality of segments of the first nickel-based alloy together along their edges to form the conical body, wherein the weld lines extend along an axial direction of the conical component;(B) Performing a first stress relief step for the hollow truncated conical body through heat 10 treatment;(C) Forming a plurality of inwardly extending projections extending from an inner surface of the hollow truncated conical body by an additive manufacture process; and(D) Performing a second stress relief step through solution heat treatment.15 2. A method as claimed in claim 1, wherein the segments are formed of a rolled plate of firstnickel-based alloy which have each been formed into a predetermined curved profile such that on alignment the conical body is defined.

3. A method as claimed in any of claims 1 to 2, wherein the plurality of segments comprise at 20 least two segments, and preferably three segments.

4. A method as claimed in claim 1, wherein the first stress relief step is performed by heating the truncated conical body to between 1065° C and 1093° C for about 1 hour.25 5. A method as claimed in claim 1 or 4, wherein the step (B) of performing the first stress reliefstep is followed by the steps of:(B2) an expansion step in which the hollow truncated conical body is expanded radially outwards until a predetermined radius of the conical body is reached; and(B3) a second stress relief step of the expanded hollow truncated conical body through30 heat treatment.

6. A method of claim 5, wherein the step of performing the second stress relief step of the expanded hollow truncated conical body is performed by heating the body to between 1065° C and 1093° C for at least about 1 hour.3519 02 267. A method as claimed in any of claims 1 or 4 to 6, wherein the plurality of inwardly extending projections extending from the inner surface of the truncated conical body are formed by an additive manufacture process to form a series of concentric circumferentially and inwardly extending surfaces.

58. A method as claimed in any of claims 1 or 4 to 7, wherein the additive manufacturing process is a laser metal deposition process using a wire or powder formed of the first nickelbased alloy.10 9. A method as claims in any of claims 1 or 4 to 8, wherein the additive manufacture processis selected from: Direct metal laser sintering (DMLS), Electron beam melting (EBM), Selective laser melting (SLM), Selective laser sintering (SLS), Direct metal wire deposition or Direct metal powder deposition.15 10. A method as claimed in any of claims 1 to 9, further comprising the step of solution heattreating the hollow truncated conical body after the inwardly extending projections have been formed by heating the body to a temperature from 1005°C to 1205°C for at least about 1 hour.

11. A method as claimed in claim 10, wherein the body is heated to a temperature from 20 1055°C to 1155°C and preferably approximately 1105°C for about 1 hour.

12. A method as claimed in any of claims 10 to 11 further comprising the step of performing carbide stabilisation of the first nickel-based alloy by heating the hollow truncated conical body to a temperature from 910°C to 1110°C for at least about 2 hours.2513. A method as claimed in claim 12, wherein the component is heated to a temperature of from 960°C to 1060°C and preferably approximately 1010°C for about 2 hours.

14. A method as claimed in any of claims 10 to 13, further comprising welding a concentric 30 flange to one or both ends of the hollow truncated conical body wherein the concentric flange is formed of a second nickel-based alloy to form a welded connection between the first nickelbased alloy and the second nickel-based alloy.

15. A method of claim 14, further comprising the step of post weld heat treating the welded 35 connection of the first nickel-based alloy and the second nickel-based alloy by heating the hollow truncated conical body to a temperature from 920°C to 1120°C and preferably from19 02 26970°C to 1070°C and more preferably a temperature of approximately 1020°C for less than about 1 hour.

16. A method as claimed in claim 14, wherein the hollow truncated conical body is heated to 5 a temperature from 920°C to 1120°C and preferably from 970°C to 1070°C and more preferably a temperature of approximately 1020°C for less than about 0.75 hours.

17. A method as claimed in any of claims 1 or 16, further comprising a second step of carbide stabilisation by heating the hollow truncated conical body to between 825°C to 830°C for about 10 4 hours.

18. A method as claimed in claim 17, wherein the carbide stabilisation is performed by heating the component to a temperature from 825°C to 843°C and preferably from 825°C to 841°C and preferably from 825°C to 840°C and more preferably from 830°C to 840°C and still 15 more preferably from 835°C to 840°C for about 4 hours.

19. A method as claimed in any of claims 1 to 18, further comprising the step of performing precipitation by heating the hollow truncated conical body to a temperature from 660°C to 860°C for up to about 16 hours.2020. A method as claimed in claim 19, wherein during precipitation the component is heated to a temperature of from 710°C to 810°C and preferably approximately 760°C for about 16 hours.25T +44(0)30 0300 2000A

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