PROCESSING METHOD

Abstract

The present disclosure relates to a method for processing a component formed by an ALM method using a -strengthened superalloy having a solvus temperature. The processing method comprises heating the component to a treatment temperature at or above the solvus temperature at a rate equal to or greater than 50 C./min and then cooling the component at a rate of greater than 60 C./min.

Claims

1. A method for processing a component formed by an ALM method using a -strengthened superalloy having a solvus temperature, the processing method comprising: heating the component to a treatment temperature at or above the solvus temperature at a rate equal to or greater than 50 C./min; and cooling the component at a rate equal to or greater than 60 C./min.

2. A method according to claim 1 wherein the treatment temperature is below the solidus temperature of the superalloy.

3. A method according to claim 1 wherein the component is maintained at or above the treatment temperature for a hold time of between 0.5-4 hours.

4. A method according to claim 1 wherein the component is cooled by gas fan quenching.

5. A method according to claim 1 wherein the component is cooled at a rate between 60 and 150 C./min.

6. A method according to claim 1 wherein the component is cooled from the treatment temperature to around 600 C. at a rate equal to or greater than 60 C./min.

7. A method according to claim 6 wherein the component is subsequently cooled from around 600 C. to room temperature at a lower cooling rate.

8. A method according to claim 1 where the component is heated using induction heating.

9. A method according to claim 1 wherein the component is inserted into a pre-heated chamber/furnace.

10. A method according to claim 1 wherein the heating and/or cooling is carried out at atmospheric/ambient pressure.

11. A method according to claim 1 wherein the method is carried out on a component which has not been subjected to hot isostatic pressing.

12. A method according to claim 1 wherein the method further comprises subjecting the component to an aging step for an aging time after cooling.

13. A method according to claim 1 wherein the component is a turbine or compressor component for use in a gas turbine aero-engine.

14. A method of manufacturing a component comprising: manufacturing the component using an ALM method comprising: depositing a layer of powdered material comprising -strengthened superalloy having a solvus temperature on a base plate and fusing at least a portion of said layer of powdered material using an energy beam to form a first fused layer of the component; depositing a second layer of powdered material comprising -strengthened superalloy on the first fused layer and fusing at least a portion of said second layer of powdered material using the energy beam to form a second fused layer onto the first fused layer; and depositing further layers of powdered material comprising -strengthened superalloy on the second/subsequent fused layers and fusing at least a portion of each of said further layers of powdered material using the energy beam to form third and subsequent fused layers of the component until the desired three dimensional component is obtained; and processing the component using the method according to any one of the preceding claims.

15. A method according to claim 14 wherein the -strengthened superalloy is a nickel superalloy.

16. A method according to claim 14 wherein the powdered material has a particle size of between 15 and 60 microns.

Description

DESCRIPTION OF THE DRAWINGS

[0048] Embodiments of the invention will now be described by way of example with reference to the accompanying figures in which:

[0049] FIG. 1 shows a schematic representation of a method of manufacturing a component;

[0050] FIG. 2 shows a schematic representation of a first embodiment of the processing method;

[0051] FIG. 3 shows a section of an additively made Haynes 282 material post additive deposition (step 1 FIG. 1);

[0052] FIG. 4 shows a section of the additively made Haynes 282 material post heat treatment (rapid heating/cooling); and

[0053] FIG. 5 shows a sectional view of a gas turbine engine.

DETAILED DESCRIPTION

[0054] A turbine or compressor component such as a blade or stator for use in a gas turbine aero-engine is manufactured using an ALM method in which a layer of powdered CM247LC having a particle size of between 15 and 60 microns is deposited on a base plate and fused into a 2D first fused layer using a scanning laser beam to melt and fuse the powdered CM247LC. Next a second layer of powdered CM247LC is deposited on the first fused layer and fused into a 2D second fused layer using the scanning laser beam, the second fused layer being fused to the first fused layer.

[0055] The deposition and fusing of powdered CM247LC is repeated until the desired 3D component is formed from the 2D layers. This is step 1 shown in FIG. 1

[0056] The component is first processed using a surface finishing step which is applied to reduce the extent of surface asperities which arise as a result of semi-fused powdered CM247LC. The surface finishing step also acts to clean the component of loose powdered CM247LC.

[0057] The surface finishing comprises grit blasting using alumina particles. Any portions of the component surface that are not easily accessible e.g. internal bores may be surface finished by abrasive flow machining.

[0058] Next, as shown in FIG. 2 and step 2 of FIG. 1, the surface-finished component is subjected to heating in an argon or nitrogen atmosphere at ambient pressure at a treatment temperature of 1260 C. ata rate greater than 100 C./min. The component is held at 1260 C. for a hold time of 2 hours followed by rapid cooling to 600 C. at a rate greater than 90 C./min. The component is subsequently cooled to room temperature at a lower cooling rate.

[0059] The heating/cooling of the component is followed by aging at an aging temperature for an aging time shown as step 3 in FIG. 1. The aging temperature is 870 C. and the aging time is around 16 hours.

[0060] Finally, as shown as step 4 in FIG. 1, the process comprises finishing by grit blasting, linishing, milling or machining, for example.

[0061] FIG. 3 shows a section of an additively made Haynes 282 material post additive deposition (step 1 FIG. 1). The image has been taken with optical microscope using dark field illumination to highlight cracks and pores.

[0062] FIG. 4 shows a section of an additively made Haynes 282 material post heat treatment (rapid heating/cooling) as described herein (step 2 FIG. 1). The image has been taken with optical microscope using dark field illumination to highlight cracks and pores. It can be seen that the cracks and pores have been closed by the heat treatment used.

[0063] With reference to FIG. 5, a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16 comprising a plurality of combustor tiles manufactured as described above, a high-pressure turbine 17, an intermediate pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

[0064] The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to 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 combusted. The combustion equipment 16 typically comprises an annular combustion chamber which is lined with the plurality of combustor tiles which can be manufactured according to the method described herein.

[0066] The resultant hot combustion products generated within the combustion equipment 16 then expand through, and thereby drive the high, intermediate and low-pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high 17, intermediate 18 and low 19 pressure turbines drive respectively the high pressure compressor 15, intermediate pressure compressor 14 and fan 13, each by suitable interconnecting shaft.

[0067] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.

[0068] 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. 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 scope of the invention.