Coated nickel superalloy

Abstract

The present invention relates to an aircraft turbine component (100) comprising: - a substrate (10) made of a nickel superalloy, the nickel superalloy being composed of a γ phase and a γ' phase; - a coating (20) covering the substrate, the coating being made of an alloy containing more than 95% by volume of the γ' phase and containing 2% to 5% of chromium atoms and 18% to 25% of aluminum atoms and nickel comprising the majority of the alloy; and - a thermal barrier layer (30); wherein the coating is disposed between the substrate and the thermal barrier layer.

Description

Technical Field

[0001] This disclosure relates to the aerospace industry, and more specifically to metallic alloys used in that field, particularly to protective coatings for such alloys. Background Technology

[0002] The efficiency of an aircraft turbine depends on its operating temperature.

[0003] In recent decades, it has become possible to increase the operating temperature of turbines, especially through the use of nickel-based superalloys.

[0004] These alloys exhibit good creep resistance at high temperatures, while also possessing corrosion resistance and oxidation resistance suitable for use in aerospace turbines.

[0005] One of the materials used in this application is an alloy commercially available under the name AM1.

[0006] Since the success of this alloy, new superalloy compositions have been proposed to further enhance the mechanical properties at high temperatures, thereby enabling further increases in turbine temperatures.

[0007] In this process, the chromium content in the superalloy composition was significantly reduced compared to AM1, which in turn made the new alloy more sensitive to the environment of aero-turbine turbines.

[0008] It has been proposed to compensate for this sensitivity by developing novel protective coatings that improve the oxidation or corrosion resistance of the underlying substrate and provide excellent integrity for thermal barriers that are typically located on the outside of turbine components to protect them from temperature effects.

[0009] However, at the operating temperature of the turbine, diffusion causes atoms to migrate, so phase transitions are often observed in the coating.

[0010] These phase transitions thus create surface undulations between the substrate and the coating, thereby promoting the exfoliation of the thermal barrier.

[0011] The term "wrinkling" is used to describe the appearance of these surface undulations between the substrate and the coating, which cause the thermal barrier set on the coating to lose adhesion.

[0012] For the latest coated superalloys (which have a higher rhenium content and a lower chromium content than AM1), secondary reaction zones (SRZs) have been found to appear in addition to wrinkling, thus reducing the service life of the entire component.

[0013] Such regions are caused by the interdiffusion of elements between the coating and the substrate. More precisely, some elements in the coating migrate into the substrate, and some elements in the substrate migrate into the coating, thus forming a region between the substrate and the coating with a phase different from both the substrate and the coating. Such regions significantly reduce the mechanical properties of the coated superalloy.

[0014] More specifically, these regions are associated with the degradation of the microstructure, which in turn leads to a decline in mechanical strength properties, particularly poor crack resistance, ultimately resulting in poor thermal barrier integrity and premature wear of components in operation.

[0015] Therefore, there is still a need for a material with the mechanical properties and oxidation resistance required to improve the operating temperature of aero-turbines compared to AM1. Summary of the Invention

[0016] The present invention aims to provide a solution to the above-mentioned problems.

[0017] For this purpose, a first aspect of the present invention relates to an aircraft turbine component comprising: - Nickel-based superalloy substrate composed of γ phase and γ' phase; - A coating covering the substrate, the coating being made of an alloy containing greater than 95% γ' phase by volume and comprising 2% to 5% chromium by atomic percentage, 18% to 25% aluminum by atomic percentage, and nickel comprising the majority of the alloy; and - Thermal barrier layer; The coating is disposed between the substrate and the thermal barrier layer.

[0018] The coating contains a γ' phase volume fraction greater than 95%.

[0019] This avoids surface undulations (sometimes referred to as "wrinkling" in the literature) between the coating and the substrate, because the coating does not undergo microstructural changes during aging. The absence of these surface undulations results in excellent thermal barrier integrity, extending the overall service life of the component.

[0020] Therefore, this coating not only meets environmental and mechanical requirements, but also provides excellent corrosion and oxidation protection, and offers mechanical properties suitable for the desired application.

[0021] In one embodiment, the substrate contains less than or equal to 7% by mass of chromium and greater than or equal to 2.5% by mass or even greater than or equal to 4.0% by mass of rhenium. Such contents are characteristic of the latest generation of nickel-based superalloys, which contain less chromium and more rhenium than AM1 generation alloys.

[0022] In these new-generation alloys, the secondary reaction zone that appears with the existing technology and coating is very obvious, while the coating as described above is more effective.

[0023] Using the coating avoids these secondary reaction zones, and in addition to the advantages already described, the coating significantly improves the mechanical properties of the resulting parts.

[0024] In one embodiment, the substrate may be selected from commercially available nickel-based superalloys under the names CMSX-4 PLUS, MCNG, and CMSX-10.

[0025] In one embodiment, the coating is disposed in direct contact with the substrate.

[0026] In one embodiment, the turbine component does not include any layer other than the substrate, the coating disposed in direct contact with the substrate, and the thermal barrier disposed in direct contact with the coating.

[0027] This method is particularly advantageous because it allows for the easy acquisition of turbine components that are very simple to assemble.

[0028] In one embodiment, the coating may contain 1.0% to 6.0% platinum content in atomic percentages.

[0029] This method can further extend the service life of the alloy by increasing the integrity of the thermal barrier.

[0030] In one embodiment, the coating may contain an aluminum content of 18% to 23% by atomic percentage.

[0031] In one embodiment, the coating may also contain a non-zero hafnium content of 0.0% to 1.0% in atomic percentage.

[0032] In one embodiment, the coating may also contain a non-zero silicon content of 0.0% to 1.0% in atomic percentage.

[0033] In one embodiment, the coating may also contain a non-zero yttrium content of 0.0% to 2.0% in atomic percentage.

[0034] In one embodiment, the coating may also contain a non-zero zirconium content of 0.0% to 2.0% in atomic percentage.

[0035] More specifically, the inventors believe that these four elements play a positive role in improving the oxidation and corrosion resistance imparted by the coating. Furthermore, these amounts of elements ensure that the coating maintains a structure containing a γ' phase volume fraction greater than 95%.

[0036] Finally, the inventors have verified that these compositions provide excellent thermal barrier integrity in turbine components.

[0037] The coating prevents thermal barrier delamination while ensuring excellent protection of the substrate from corrosion and oxidation, even for substrates that are inherently more sensitive to corrosion and oxidation than AM1.

[0038] In one embodiment, the substrate comprises 25% to 35% γ phase volume fraction and complementary γ' phase volume fraction.

[0039] The term "complementary volume fraction" should be understood to mean that the sum of the volume fractions of the γ phase and the γ' phase is equal to 100%, because the substrate is composed of the γ phase and the γ' phase.

[0040] The phase content is known to vary with temperature. The given values ​​apply to the efficient operating temperature of the component, such as 1000°C to 1200°C.

[0041] The coating contains a γ' phase volume fraction greater than 95% by volume, which ensures that the coating does not undergo microstructural changes throughout its entire service life.

[0042] More specifically, the change of the γ' phase towards γ over time does not produce surface undulations (wrinkling) and is also extremely slow because the coating is close to equilibrium with the substrate under the conditions encountered by the turbine.

[0043] By ensuring that the initial volume fraction of the γ' phase is greater than 95% of the coating, the influence of changes in the microstructure over time on the coating structure is very small, thus limiting the delamination of the thermal barrier.

[0044] In one embodiment, the coating contains a γ' phase volume fraction greater than 99% by volume, or is even composed of the γ' phase.

[0045] As mentioned above, the more γ' phase present, the less we need to worry about the microstructure changes of the coating over time.

[0046] In one embodiment, the coating does not contain any element other than nickel, chromium, aluminum, and optionally platinum, hafnium, zirconium, silicon, or yttrium, in an amount greater than or equal to 0.1 atomic%.

[0047] This method ensures that the coating achieves the desired beneficial effects while maintaining a γ' phase volume fraction of ≥95%.

[0048] In one embodiment, the coating thickness is 5.0 µm to 50 µm, or even 10 µm to 30 µm.

[0049] This thickness is less than that of existing technology coatings with comparable performance, thereby enabling weight reduction for components.

[0050] In one implementation, the average composition of the coating is uniform across its entire thickness after deposition. This means that the coating does not contain compositional gradients.

[0051] Typically, the thermal barrier layer may include a yttrium oxide zirconium oxide layer.

[0052] In one embodiment, the thermal barrier layer may comprise a gadozirconate (Gd2Zr2O7) layer.

[0053] In one embodiment, the thermal barrier layer may comprise a yttrium oxide zirconium oxide layer and a gadozirconate layer disposed in direct contact with each other.

[0054] In one embodiment, the turbine component is a turbine blade or guide vane, preferably a hot-section turbine blade or guide vane.

[0055] This invention is particularly advantageous for such components because it enables a reduction in coating thickness while maintaining equivalent strength performance, and turbine blades or guide vanes whose geometry is constrained due to aerodynamic reasons particularly benefit from this advantage.

[0056] According to another aspect, the present invention also relates to a method for obtaining the aircraft turbine component as described above.

[0057] The method includes: - A step of coating the outer surface of a substrate made of a nickel-based superalloy, the nickel-based superalloy being composed of a γ phase and a γ' phase, the coating being made of an alloy containing more than 95% γ' phase by volume and containing 2% to 5% chromium by atomic percentage and 18% to 25% aluminum by atomic percentage, and nickel constituting the majority of the alloy; and - The step of setting a thermal barrier on the coating.

[0058] In one embodiment, the coating step can be performed by physical vapor deposition (PVD) (e.g., electron beam physical vapor deposition (EB-PVD)), by electric arc or magnetron, by high-speed oxygen fuel spraying (HVOF), or by plasma process.

[0059] All these methods enable coatings to be prepared directly from a γ' phase volume fraction exceeding 95%, thus limiting interdiffusion.

[0060] The extremely low interdiffusion between the substrate and the coating ensures the absence of secondary reaction zones, which ensures the achievement of the aforementioned coating's technical effects, particularly the better time-dependent stability of the thermal barrier set on the coating compared to existing coatings.

[0061] Furthermore, these methods ensure that the coating is deposited, rather than obtained by the diffusion of elements from the substrate to the coating itself, thus ensuring the absence of secondary reaction zones.

[0062] In one embodiment, the method may further include a homogenization diffusion heat treatment at 1000°C to 1200°C for 1 hour to 6 hours in a vacuum or argon or oxygen atmosphere between the coating step and the step of providing a thermal barrier.

[0063] Homogenization (if applicable) ensures excellent reproducibility of the resulting parts. This step is not strictly necessary but can be implemented as a precaution to ensure that all parts produced by this method have the same thermal history and therefore uniform composition.

[0064] Thermal barriers can be deposited onto coatings using methods known per se, preferably the same methods used for depositing coatings.

[0065] Unlike existing methods, the method proposed in this paper enables coatings to be obtained in a single deposition step, performed in a single machine.

[0066] Typically, existing deposition methods involve an initial step of depositing electrolytic platinum, followed by an aluminization step via chemical vapor deposition, which requires changing machines between the two steps.

[0067] Compared with existing methods, this implementation saves time and improves the ease of deposition.

[0068] The properties of the coating (including its composition and microstructure) ensure that the coating and the thermal barrier placed on it maintain excellent integrity over time.

[0069] According to another aspect, the present invention relates to an aircraft turbine comprising the components described above, wherein the components are blades or guide vanes.

[0070] Such turbines can operate at higher temperatures than those used in existing technologies, making them more efficient.

[0071] More specifically, compared to blades or guide vanes of the prior art, the coated blades or guide vanes as described above offer better oxidation resistance and better temperature resistance, which enables the turbine to operate at higher temperatures overall.

[0072] Brief description of the attached figures

[0073] [ Figure 1 ] Figure 1 This is a schematic diagram of a turbine.

[0074] [ Figure 2 ] Figure 2 A coated turbine blade or guide vane according to one embodiment of the present invention is shown. Detailed Implementation

[0075] The invention will now be described with reference to the accompanying drawings, which are intended to illustrate certain embodiments of the invention and should not be construed as limiting the invention.

[0076] Figure 1 A vertical cross-section through the main shaft A of the turbofan engine 1 is shown. It includes a fan 2, a low-pressure compressor 3, a high-pressure compressor 4, a combustion chamber 5, a high-pressure turbine 6, and a low-pressure turbine 7, flowing along the airflow from upstream to downstream.

[0077] In this application, relative positional terms such as “upstream,” “downstream,” “inner,” and “outer” are to be understood as relative to the casing horizontal shaft A that defines the axial direction, which passes through the flow direction of the turbine’s main and secondary airflows.

[0078] Therefore, the "upstream" element will be passed before the "downstream" element, and the "inner" element will be closer to the A-axis than the "outer" element.

[0079] Figure 2 The moving blades are shown, which may belong to the low-pressure compressor 3, the high-pressure compressor 4, the high-pressure turbine 6, or the low-pressure turbine.

[0080] More precisely, blade 100 includes: - A substrate 10 made of a nickel-based superalloy, said nickel-based superalloy being composed of a γ phase and a γ' phase; and - A coating 20 covering the substrate 10, the coating being made of an alloy containing more than 95% γ' phase volume fraction and containing 2% to 5% chromium content and 18% to 25% aluminum content and nickel comprising the majority of the alloy; - Thermal barrier layer 30.

[0081] As mentioned above, such blades or guide vanes offer better temperature resistance and better thermal barrier adhesion, resulting in a longer component life compared to existing turbine components.

[0082] Figure 2 This further clarifies the meaning of the “thickness” e1 of coating 20.

[0083] The term is used in its conventional sense in this document and application, referring to the shortest distance required to cover the coating from one side to the other.

[0084] In this application, the γ phase and γ' phase of nickel-based superalloys have the meanings conventionally used in the art.

[0085] Nickel-based superalloys can consist of: a γ-Ni face-centered cubic austenitic γ-phase (or matrix) optionally containing α-substituted solid solution additives (Co, Cr, W, Mo); and a γ'-Ni3X γ' phase (or precipitate), where X = Al, Ti, or Ta. The γ' phase has an ordered L... 12 The structure is derived from the face-centered cubic structure and has a coherent relationship with the matrix, meaning it has an atomic lattice that is very similar to that of the matrix.

[0086] The amount of a phase in a given substrate or coating can be determined by methods known to those skilled in the art, such as by statistical determination of a sample observed by scanning electron microscopy, or by sequentially performing an EDS microscopy analysis step to determine the chemical composition and then simulating the precise chemical composition determined by the EDS analysis. Simulations are performed using tools known per se, such as THERMOCALC® software.

[0087] In one embodiment, the substrate may be selected from commercially available nickel-based superalloys under the names CMSX-4 PLUS, MCNG, and CMSX-10.

[0088] As described above, coating 20 is a nickel-based alloy and contains a γ' phase volume fraction greater than 95%.

[0089] The blade or guide vane 100 also includes a thermal barrier layer 30 disposed on the coating.

[0090] Preferably, the thermal barrier layer 30 is the outer surface of the blade or guide vane 100.

[0091] The thermal barrier layer is selected according to the needs of the invention and general practice in the field.

[0092] In one embodiment, the thermal barrier layer may comprise a yttrium oxide-zirconia layer.

[0093] In one embodiment, the thermal barrier layer may comprise a gadozirconate (Gd2Zr2O7) layer.

[0094] In one embodiment, the thermal barrier layer may comprise a yttrium oxide zirconium oxide layer and a gadozirconate layer disposed in direct contact with each other.

[0095] For illustrative purposes, the thermal barrier layer is shown here as a single layer 30.

[0096] Figure 2 The thickness e2 of the thermal barrier 30 is shown.

[0097] In one embodiment, the thickness e2 of the thermal barrier 30 can be from 100 µm to 200 µm.

[0098] Preferably, the thermal barrier layer 30 is the outer surface of the blade or guide vane 100.

Claims

1. An aircraft turbine component (100), comprising: - A substrate (10) made of a nickel-based superalloy, the nickel-based superalloy being composed of a γ phase and a γ' phase, - A coating (20) covering the substrate, the coating being made of an alloy consisting of a γ' phase and containing 2% to 5% chromium and 18% to 25% aluminum by atomic percentage, and nickel comprising the majority of the alloy; as well as - Thermal barrier layer (30); The coating is disposed between the substrate and the thermal barrier layer.

2. The aircraft turbine component (100) as claimed in claim 1, wherein, The substrate (10) contains 25% to 35% γ phase volume fraction and complementary γ' phase volume fraction.

3. The aircraft turbine component (100) as claimed in claim 1 or claim 2, which does not include any layer other than the substrate (10), the coating (20) disposed in direct contact with the substrate, and the thermal barrier (30) disposed in direct contact with the coating.

4. The aircraft turbine component (100) as claimed in any one of claims 1 to 3, wherein, The coating (20) contains 1.0% to 6.0% platinum by atomic percentage.

5. The aircraft turbine component (100) as claimed in any one of claims 1 to 4, wherein, The coating (20) further comprises: a non-zero hafnium content of 0.0% to 1.0% by atomic percentage; and a non-zero silicon content of 0.0% to 1.0% by atomic percentage; The non-zero zirconium content is 0.0% to 2.0% by atomic percentage; and / or the non-zero yttrium content is 0.0% to 2.0% by atomic percentage.

6. The aircraft turbine component (100) as claimed in any one of claims 1 to 5, wherein, The coating (20) does not contain any element with a content greater than or equal to 0.1 atomic percent other than nickel, chromium, aluminum and optionally platinum, hafnium, zirconium, silicon and yttrium.

7. The aircraft turbine component (100) as claimed in any one of claims 1 to 6, wherein, The thickness (e1) of the coating (20) can be from 5.0 µm to 50 µm, preferably from 10 µm to 30 µm.

8. A method for obtaining an aircraft turbine component (100) as claimed in any one of claims 1 to 7, the method comprising: - A step of coating the outer surface of a substrate (10) made of a nickel-based superalloy with a coating (20), the nickel-based superalloy being composed of a γ phase and a γ' phase, the coating (20) being made of an alloy containing more than 95% γ' phase volume fraction and containing 2% to 5% chromium content by atomic percentage and 18% to 25% aluminum content by atomic percentage and nickel comprising the majority of the alloy; as well as - The step of setting a thermal barrier (30) on the coating.

9. The method of claim 8, wherein, The coating step is performed by physical vapor deposition (PVD), optionally by electron beam physical vapor deposition (EB-PVD); by electric arc or magnetron; by high-speed oxygen fuel spraying (HVOF); or by plasma process.

10. An aircraft turbine comprising the components as claimed in any one of claims 1 to 7, wherein, The component is a blade or guide vane.

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