JP2010043351A - Thermal barrier coating and method for production thereof - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a thermal insulating ceramic layer for use on metal alloy components exposed to hostile thermal and chemical environment such as a gas turbine. <P>SOLUTION: The thermal barrier coating includes cracks made by a series of steps, including a step of applying a shockwave onto at least a portion of the thermal barrier coating so as to form microcracks in the thermal barrier coating without substantially deforming a substrate. <P>COPYRIGHT: (C)2010,JPO&INPIT

Description

  The present invention broadly relates to protective coatings for alloy parts that are exposed to hot gas environments and harsh operating conditions, such as the working parts of a power generating gas turbine engine. Specifically, the present invention relates to a thermal barrier coating (TBC) for use in a gas turbine engine and a method for manufacturing a TBC.

  The operating conditions to which the gas turbine hardware components are exposed can be severely thermally and chemically. Metal substrate surfaces used in turbines, combustors and augmentor parts should exhibit above average mechanical strength, durability and erosion resistance in very harsh hot gas environments. “Erosion” generally refers to a process in which contaminating particles of high energy sufficient to release other particles from the surface (erosion) collide with the surface (especially metal), leading to degradation and cracking of the substrate. .

  Recently, great progress has been achieved by utilizing heat resistant alloys in gas turbine systems by introducing iron, nickel and cobalt based superalloys into the coatings applied to the bases of critical turbine components. There are generally two purposes for effective surface coatings. First, the coating should form a protective and adhesive layer that protects the underlying substrate from oxidation, corrosion and degradation. Second, the coating should have a low thermal conductivity with respect to the substrate. As superalloy compositions become more complex, it is difficult to achieve the required high level of strength and sufficient corrosion and oxidation resistance simultaneously (especially at high gas turbine operating temperatures). Due to the trend of higher gas turbine combustion temperatures, the problems of oxidation, corrosion and degradation are becoming more difficult. Therefore, despite recent improvements in thermal barrier coatings, many alloy components are not economical and effective because they cannot withstand long-term service exposure and repeated cycles found in typical gas turbine environments, In addition, there remains a great need for heat resistant coatings that are less prone to degradation.

  Many of the known prior art coatings for gas turbine components contain aluminide and ceramic components. Typically, the ceramic coating was formed from an oxidation resistant alloy such as MCrAlY (where M is iron, cobalt and / or nickel) or from a diffusion aluminide or platinum aluminide that forms an oxidation resistant intermetallic compound. Used with bond coats. In high temperature applications, these bond coats form an oxide layer (also referred to as “scale”) that chemically bonds to the ceramic layer to form the final bond coat.

It is also known to use zirconia (ZrO ₂ ) partially or wholly stabilized with yttria (Y ₂ O ₃ ), magnesia (MgO) or other oxides as the main component of the ceramic layer. Yttria-stabilized zirconia (YSZ) exhibits a favorable thermal cycle fatigue property and is therefore diverse as a ceramic layer for thermal barrier coatings. That is, YSZ is far superior to other known coatings in the ability to withstand stress and fatigue when the temperature rises or falls from the start to the stop of the gas turbine. Typically, the deposition of YSZ on a metal substrate is a known method such as atmospheric plasma spray (“APS”) or low pressure plasma spray (“LPPS”), as well as electron beam physical vapor deposition (“EBPVD”). By physical vapor deposition (“PVD”). In particular, YSZ deposited by EBPVD is characterized by a stress-resistant columnar grain structure that allows the substrate to expand and contract without causing harmful stresses that cause spalling. The stress resistance of such systems is known. For example, see US Pat. No. 6,730,413 for a known thermal barrier coating system.

  The formation of longitudinal cracks in the manufacturing environment is difficult or problematic. In one aspect, the present invention generally relates to a method that includes inducing cracks or microcracks after film deposition. As a result, the thermal barrier coating can be densely applied, but it is easier to apply. After application, the coating can be selectively cracked using, for example, shock wave exposure.

  Laser peening is well known. For example, laser peening has been used to create a compressive stress protection layer on the outer surface of the workpiece, and such a process can be applied to the workpiece against fatigue failure as disclosed in US Pat. No. 4,937,421. It is known to greatly increase resistance. Laser shock peening has also been used to create deep compressive residual stresses in turbine blades, as disclosed in US Pat.

US Pat. No. 6,730,413 US Pat. No. 4,937,421 US Pat. No. 5,591,009

  In one aspect, the invention relates to a method for cracking a thermal barrier coating applied to a gas turbine component. The method includes depositing a bond coat comprising MCrAlY (wherein M is iron, cobalt, and / or nickel) on a metal substrate forming a gas turbine component, and including yttria stabilized zirconia on the bond coat. Depositing a thermal barrier coating, subjecting at least a portion of the thermal barrier coating to a shock wave to form microcracks in the thermal barrier coating without substantially deforming the metal substrate.

  In one aspect, the present invention relates to a method for forming a crack in a ceramic-based coating. The method includes depositing a ceramic-based coating including a thermal barrier coating on a metal-based substrate, subjecting at least a portion of the ceramic-based coating to a shock wave, and without substantially deforming the metal-based substrate. Forming a microcrack in the ceramic coating.

FIG. 1 is a cross-sectional view of a metal substrate such as a high-pressure gas turbine blade, showing a thermal barrier coating applied to the blade using the laser shock method according to an embodiment of the present invention. FIG. 2 outlines the amount of energy required to induce cracks in the thermal barrier coating.

  As mentioned above, the thermal barrier coating of the present invention can be applied to a variety of alloy parts (so-called “superalloys”) that must be protected from thermally and chemically harsh environments. Examples of such components include nozzles, buckets, shrouds, airfoils and other hardware found in most gas turbine engines.

  The coating may be any known TBC composition, for example from a thermal barrier ceramic layer having a composition that when deposited increases the erosion resistance of turbine components significantly while maintaining a spalling resistance equal to or better than conventional coatings. It may be. A coating composition may be deposited and cracked after deposition.

  High pressure turbine blades are a prime example of a substrate on which the coating of the present invention can be applied. Typically, turbine blades have airfoils and platforms, and hot combustion gases flow toward them during gas turbine operation. Therefore, the airfoil surface is subject to oxidation, corrosion, and erosion. The airfoil is typically implanted in the turbine disk with a dovetail formed at the root of the blade.

  FIG. 1 shows a thermal barrier coating according to the present invention provided on a substrate. The coating 10 includes a thermal barrier ceramic layer 12 on a bond coat 14, which is provided on an alloy substrate 16 that forms the base material of the turbine blade. Suitable materials for the substrate include iron-based, nickel-based and / or cobalt-based superalloys. The bond coat is oxidation resistant and forms an alumina layer 18 on the surface of the bond coat upon high temperature exposure of the coated blade. The alumina layer protects the underlying superalloy substrate 16 from oxidation and provides a surface to which the ceramic layer adheres.

  Layer 12 has longitudinal cracks formed to increase and / or cause stress resistance. By inducing cracks by exposure to shock waves, cracks can be introduced at a desired density into a material, particularly a desired region of the material. For the formation of cracks, a shock wave may be induced in the material using coupled ablation. Coupled ablation can be achieved by using a pulsed laser in a manner similar to the laser shock peening method in which a shock wave is generated by irradiating a laser pulse through the coupling material and into the ablation material.

  In the prior art, laser shock peening is used for densification of materials. However, in the case of TBC, the resulting shock wave can induce microcracks in the coating to provide stress resistance. Other shock wave exposure means may be used. Other coupled ablation means are also possible.

  In an exemplary embodiment, the stress resistant TBC can be formed using a laser shock peening method. Thermal barrier coatings can be applied to metal substrates using atmospheric plasma spraying. The bond coat is MCrAlY (where M is iron, cobalt and / or nickel) and the TBC may be a ceramic based coating used as a thermal barrier coating for 8% yttria stabilized zirconia and other turbine components. What is necessary is just to laser-shock-peen TBC after constructing to a base material.

  The energy used to induce microcracks in the TBC should preferably not substantially deform the substrate. For example, the energy should be relatively low since the coating can be very thin. In order to substantially deform the base material, it is necessary that the energy of the shock wave is sufficient to apply a stress equal to or higher than the plastic yield point of the base material, but less than its compressive strength. In contrast, the energy used to induce microcracks in the TBC should be sufficient to apply a stress that exceeds the compressive strength of the TBC. Since metal substrates are ductile and ceramic TBCs are brittle, there will be a certain level of energy that can be selected or determined.

  FIG. 2 outlines the amount of energy required to induce cracks in the TBC. The amount of energy (per unit area) to break the material is expressed as the area under the stress / strain curve. FIG. 2 shows a typical porous TBC coating. The porosity reduces the “effective” cross-sectional area, thus reducing the force required for failure (since energy is a function of force rather than pressure or stress). Thus, the area under the curve is significantly reduced.

  In a preferred embodiment, the thermal barrier coating receives a shock wave (eg, by laser ablation) and breaks. The energy required may depend on the shock wave source (such as laser ablation) and / or the properties of the material causing the crack.

  Thus, in an embodiment, the process (eg, laser ablation or laser shock peening) can produce microstructural features (eg, longitudinal cracks). This achieves improved durability of the turbine components and / or reduced manufacturing costs. For example, a simple dense coating may be applied to a part to induce a vertical crack in an area where a vertical crack is required. That is, it is not necessary to introduce cracks throughout the coating depending on the process parameters.

  Although the present invention has been described with respect to the most practical and preferred embodiments at the present time, the present invention is not limited to the disclosed embodiments, and the technical ideas and techniques described in the claims. Various modifications and equivalents belonging to the scope are included.

Claims (9)

A method of forming a crack in a thermal barrier coating applied to a gas turbine component,
Depositing a bond coat comprising MCrAlY (wherein M is iron, cobalt and / or nickel) on a metal substrate forming a gas turbine component;
Depositing a thermal barrier coating comprising yttria stabilized zirconia on the bond coat;
Subjecting at least a portion of the thermal barrier coating to a shock wave to form microcracks in the thermal barrier coating without substantially deforming the metal substrate. The method of claim 1, wherein the gas turbine component comprises a blade. The method of claim 1, wherein the thermal barrier coating comprises 8% yttria stabilized zirconia. The method of claim 1, wherein subjecting at least a portion of the thermal barrier coating to a shock wave comprises laser ablation or laser shock peening. The method of claim 1, wherein subjecting at least a portion of the thermal barrier coating to a shock wave does not include subjecting the entire thermal barrier coating to the shock wave. The method of claim 1, wherein depositing a bond coat on the metal substrate comprises atmospheric plasma spraying the bond coat on the metal substrate. A method of forming cracks in a ceramic coating,
Depositing a ceramic-based coating including a thermal barrier coating on a metal-based substrate;
Subjecting at least a portion of the ceramic coating to a shock wave to form microcracks in the ceramic coating without substantially deforming the metal substrate. The method of claim 7, wherein the ceramic-based coating comprises yttria stabilized zirconia. The method of claim 7, wherein the metal-based substrate comprises a superalloy comprising one or more of iron, nickel, or cobalt.

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