MATERIAL DEPOSITION METHOD FOR REPAIRING AERONAUTICAL COMPONENTS

Abstract

A method is disclosed for repairing an aeronautical component comprising a nickel-based alloy. An aeronautical component is disclosed comprising a nickel-based alloy and one or more of the following elements: tungsten, cobalt, chromium, aluminum, molybdenum, tantalum, titanium, hafnium, carbon, boron, and zirconium.

Claims

1. A method for repairing an aeronautical component comprising a nickel-based alloy, the method comprising the following steps: a) performing surface preparation on a damaged portion of the component; and b) depositing a plurality of layers on the damaged portion of the component by cold spraying a powdered material by a non-oxidizing carrier gas, wherein the powdered material which is deposited is the same alloy as the alloy comprised in the component to be repaired or an alloy from the same family, and deposition is performed with pressure conditions between 40 and 80 bar and temperature conditions between 850 and 1100° C. of the carrier gas.

2. The method according to claim 1, wherein the nickel-based alloy further comprises one or more of the following elements: tungsten, cobalt, chromium, aluminum, molybdenum, tantalum, titanium, hafnium, carbon, boron, and zirconium.

3. The method according to claim 2, wherein the nickel-based alloy further comprises the one or more of the following elements: tungsten, cobalt, chromium, aluminum, molybdenum, tantalum, titanium, and hafnium.

4. The method according to claim 1, wherein step b) comprises: depositing an anchoring layer with first pressure and temperature conditions of the carrier gas; and depositing at least one growth layer on the anchoring layer with second pressure and temperature conditions of the carrier gas, wherein the first pressure and temperature conditions are different from the second pressure and temperature conditions, or the first pressure and temperature conditions are the same as the second pressure and temperature conditions.

5. The method according to claim 4, further comprising depositing a plurality of growth layers.

6. The method according to claim 5, wherein depositing the plurality of growth layers comprises depositing at least four growth layers.

7. The method according to claim 1, wherein the alloy comprises up to 0.2% carbon, 5-14% chromium, 5-15% cobalt, up to 4% molybdenum, up to 12% tungsten, up to 12% tantalum, up to 1% vanadium, up to 1% niobium, 3-6% aluminum, up to 5% titanium, up to 2% hafnium, up to 0.02% boron, up to 0.06 zirconium, and up to 3.1% rhenium.

8. The method according to claim 1, wherein step a) comprises machining and/or cleaning the surface of the damaged portion of the component and/or providing roughness to said surface.

9. The method according to claim 8, wherein the roughness is provided by abrasion.

10. The method according to claim 1, further comprising applying a protective surface layer for protection against corrosion, wear, and/or impact on at least the damaged portion of the component after step b).

11. The method according to claim 10, wherein the application of the protective surface layer comprises coating at least the damaged portion of the component with aluminum.

12. The method according to claim 1, further comprising performing a heat treatment on the damaged portion of the component in temperature conditions between 800 and 1340° C. after step b).

13. The method according to claim 1, further comprising performing non-destructive tests on the damaged portion of the component to check the condition thereof.

14. The method according to claim 1, wherein in step b) deposition is performed at a distance of 15-40 mm from the surface of the damaged portion of the component.

15. The method according to claim 14, wherein in step b) deposition is performed at the distance of 25 mm from the surface of the damaged portion of the component.

16. The method according to claim 1, wherein the powdered material has a particle size comprised between 10 and 75 μm.

17. The method according to claim 1, wherein in step b) the pressure of the carrier gas is in the range between 40 and 70 bar and the temperature is in the range between 1000 and 1100° C.

18. An aeronautical component formed by a nickel-based alloy and one or more of the following elements: tungsten, cobalt, chromium, aluminum, molybdenum, tantalum, titanium, hafnium, carbon, boron, and zirconium, the component comprising: at least one damaged portion; and a plurality of layers deposited on the damaged portion, wherein the plurality of deposited layers is of the same alloy as the alloy comprised in the aeronautical component or of an alloy from the same family.

19. The aeronautical component according to claim 18, wherein the aeronautical component is formed by the one or more of the following elements: tungsten, cobalt, chromium, aluminum, molybdenum, tantalum, titanium, and hafnium.

20. The method according to claim 10, further comprising performing a heat treatment on the damaged portion of the component in temperature conditions between 800 and 1340° C. after the application of the protective surface layer.

Description

DETAILED DESCRIPTION

[0063] The present disclosure discloses an improved method for repairing aeronautical components with nickel-based alloys having a poor weldability by means of cold spraying, achieving very compact coatings that are free of porosity and defects in the repaired areas.

[0064] Some of the superalloys which are of particular interest for the disclosure as they comprise a nickel base and have a poor weldability are, for example, MARM-247, CM247LC, and CMSX-4.

[0065] The superalloy MARM-247 comprises the following elements expressed as % by weight: 0.13-0.17% carbon, 8-8.80% chromium, 9-11% cobalt, 0.5-0.8% molybdenum, 9.75-10.25% tungsten, 2.8-3.1% tantalum, 5.35-5.65% aluminum, 0.9-1.2% titanium, 0.01-0.02% boron, 0.03-0.06% zirconium, 1.4-1.6% hafnium, incidental impurities and nickel to balance the alloy. In an embodiment, the mentioned incidental impurities comprise the following elements expressed as % or ppm by weight: 0-0.1% copper, 0-0.1% rhenium, 0-0.5% iron, 0-0.1% manganese, 0-0.1% niobium, 0-0.01% phosphorus, 0-0.1% silicon, 0-0.1% vanadium, 0-2 ppm silver, 0-20 ppm arsenic, 0-0.5 ppm bismuth, 0-0.2 ppm cadmium, 0-25 ppm gallium, 0-0.2 ppm indium, 0-50 ppm magnesium, 0-35 ppm nitrogen, 0-20 ppm oxygen, 0-5 ppm lead, 0-0.015 ppm sulfur, 0-5 ppm antimony, 0-3 ppm selenium, 0-20 ppm tin, 0-1 ppm tellurium, 0-0.5 ppm thallium, and 0-5 ppm zinc.

[0066] In an example, the superalloy MARM-247 comprises the following elements expressed as % by weight: 0.15% carbon, 8% chromium, 10% cobalt, 0.6% molybdenum, 10% tungsten, 3.0% tantalum, 5.5% aluminum, 1.0% titanium, 0.015% boron, 0.03% zirconium, 1.5% hafnium, incidental impurities and nickel to balance the alloy.

[0067] The superalloy CM247LC comprises the following elements expressed as % by weight: 0.07-0.08% carbon, 8-8.5% chromium, 9.0-9.4% cobalt, 0.4-0.6% molybdenum, 9.3-10% tungsten, 3.1-3.3% tantalum, 5.4-5.7% aluminum, 0.6-0.9% titanium, 0.010-0.018% boron, 0.006-0.015% zirconium, 1.3-1.6% hafnium, incidental impurities and nickel to balance the alloy. In an embodiment, the incidental impurities comprise the following elements expressed as % or ppm by weight: 0-0.01% copper, 0-0.15% iron, 0-0.05% manganese, 0-0.1% niobium, 0-0.005% phosphorus, 0-0.03% silicon, 0-0.1% vanadium, 0-3 ppm silver, 0-20 ppm arsenic, 0-0.2 ppm bismuth, 0-1 ppm cadmium, 0-15 ppm gallium, 0-0.2 ppm indium, 0-80 ppm magnesium, 0-15 ppm nitrogen, 0-15 ppm oxygen, 0-2 ppm lead, 0-10 ppm sulfur, 0-5 ppm antimony, 0-2 ppm selenium, 0-15 ppm tin, 0-1 ppm tellurium, 0-0.3 ppm thallium, 0-5 ppm zinc.

[0068] In an example, the nickel-based superalloy CM247LC comprises the following elements expressed as % by weight: 0.07% carbon, 8% chromium, 9% cobalt, 0.5% molybdenum, 10% tungsten, 3.2% tantalum, 5.6% aluminum, 0.7% titanium, 0.015% boron, 0.010% zirconium, 1.4% hafnium, incidental impurities and nickel to balance the alloy.

[0069] The superalloy CMSX-4 comprises the following elements expressed as % by weight: 6.0-6.6% chromium, 9.0-10% cobalt, 0.5-0.7% molybdenum, 6.0-6.6% tungsten, 6.3-7.0% tantalum, 5.40-5.75% aluminum, 0.9-1.1% titanium, 0.07-0.12% hafnium, 2.8-3.1% rhenium, incidental impurities and nickel to balance the alloy. In an embodiment, the incidental impurities comprise the following elements expressed as % or ppm by weight: 0-100 ppm carbon, 0-20 ppm magnesium, 0-150 ppm zirconium.

[0070] In an example, the nickel-based superalloy CMSX-4 comprises the following elements expressed as % by weight: 6% chromium, 9% cobalt, 0.6% molybdenum, 6% tungsten, 7% tantalum, 5.6% aluminum, 1.0% titanium, and 0.1% hafnium, incidental impurities and nickel to balance the alloy.

[0071] A particular example of the method for repairing an aeronautical component made up of the MARM-247 alloy is described below. This method comprises the following two steps:

[0072] a) performing surface preparation on a damaged portion of the component,

[0073] b) depositing a plurality of layers on the damaged portion of the component by means of cold spraying a powdered material by means of a non-oxidizing carrier gas.

[0074] Before starting the repair, the region of the component where the damaged portion is located is identified. Once this damaged portion is located, the method is started with step a), in which surface preparation of said damaged portion of the component is performed. In one embodiment, in this first step, machining by means of chip removal, which removes part of the material of this component, is carried out to thereby ensure a proper impact of the powder particles during the subsequent deposition step, and to in turn facilitate adherence between the elements. In turn, cleaning of the surface is performed, removing possible oxides or dirt that may be found on the surface of the damaged portion to be treated so as to ensure that the repair is effective. Alternatively, roughness can also be provided on the actual surface of the damaged portion by means of abrasion (for example by means of sanding, abrasive blasting, etc.), thereby improving the mechanical anchoring between the layers of deposited material and the treated surface in the next step.

[0075] Once the surface preparation of the damaged portion has been performed, deposition step b) is started. Deposition is performed by means of cold spraying the MARM-247 alloy (the same alloy forming the damaged component to be repaired) on the damaged portion of the component in the form of layers that are gradually stacked one on top of another on the portion in question. This cold spraying is carried out by means of the action of a non-oxidizing carrier gas in specific pressure and temperature conditions. In one embodiment, MARM-247 powder particles having a size between 10 and 75 μm are cold sprayed.

[0076] In this particular example, deposition is carried out by first depositing a first anchoring layer and a plurality of growth layers are then sequentially deposited on this anchoring layer until reaching at least 4 growth layers. The pressure and temperature conditions for the deposition of the anchoring layer and growth layers are different. In particular, in one embodiment, the anchoring layer is deposited with a pressure and a temperature of the carrier gas that are lower than the pressure and temperature used for the growth layers.

[0077] In this deposition step b), the powder particles are sprayed at a distance of 25 mm from the surface on which these particles will be deposited, forming the anchoring layer and the growth layers.

[0078] Once the deposition step has been completed, in a particular embodiment, a coating type protective surface layer protecting the damaged portion that has been repaired against corrosion, wear, or impact is applied. This protective coating covers all the layers deposited on the damaged portion. The surface condition of the component after repair is thereby assured. In a particular embodiment, the coating applied is an aluminum coating or an intermetallic coating rich in this element.

[0079] Furthermore, to favor bonding between the deposited layers and the treated surface, in one embodiment, a heat treatment is carried out in temperature conditions between 800 and 1340° C. This treatment can be carried out directly on the damaged portion of the component after step b) has been completed or on the protective surface layer. This heat treatment in turn relieves the stresses between the layers of deposited material and the surface of the damaged portion and also favors atomic diffusion either between the anchoring layer and the surface of the damaged portion or else between the layers of deposited material.

[0080] Several repair tests have been performed on damaged portions in aeronautical components according to the conditions of the particular example above, in which the MARM-247 alloy has been used. Successful outcomes with very compact coatings, a porosity below 1%, and an excellent adherence at about 65-70 MPa before heat treatment, are achieved by means of the present method.

[0081] Although the example above refers to a component manufactured from the MARM-247 alloy, the present disclosure is applicable to components manufactured from other materials. Furthermore, although a component manufactured from MAR-M247 has been repaired in the described example using powder of the same alloy, in other embodiments a different alloy can be used for repair, for example, MAR-M247 powder can be used to repair a component manufactured from MAR-M002, since the latter is part of the same family as the MAR-M247 alloy.

[0082] The present method also contemplates the option of performing non-destructive tests on the damaged portion of the component to check its condition. These tests can be carried out during the repair process to ensure the absence of defects or indications of damage such as cracks, pores, or inclusions.