JP2022080663A - Repair method for gas turbine component - Google Patents

Abstract

To provide a repair method for a gas turbine component such as a high-temperature part like a combustor liner and a transition piece.SOLUTION: A repair method for a gas turbine component includes forming a heat insulation coating including a coupling layer and a ceramic layer formed via the coupling layer on the first surface of the gas turbine component, and forming a metal layer having a higher antioxidation characteristic than the material forming the gas turbine component on a thickness reduction portion of the second surface of the gas turbine component.SELECTED DRAWING: Figure 1

Description

Embodiments of the present invention relate to a method of repairing parts of a gas turbine.

Of the gas turbine parts, especially liners and transition pieces that are exposed to high temperatures due to combustion gas, avoid thinning and wear due to mechanical factors such as vibration and sliding during operation and erosion due to high temperature gas. I can't. Therefore, a method of reinforcing a portion of a gas turbine component that is easily worn or damaged (Patent Document 1) and a method of suppressing wear by forming a hard film (Patent Document 2) are known.

Japanese Patent No. 3851161 Japanese Patent Application No. 9-191554

Gas turbine components, especially combustor liners and transition pieces, are required to have a long operating life under high temperature conditions. In recent years, it has become clear that not only mechanical thinning but also thinning due to chemical factors due to high temperature oxidation occurs in parts that have been operated for a long period of time.

Since wall thinning due to high-temperature oxidation is a factor that determines the final life of high-temperature parts, while repairing such wall thinning due to oxidation, it is possible to suppress the occurrence and progress of re-oxidation thinning after repair. A repair method is desired.

In addition, ceramics with low thermal conductivity called Thermal Barrier Coating (hereinafter, may be referred to as TBC in the present specification) are on the inner surface side of the liner, transition piece, etc., which are the flow paths of high-temperature gas. A heat shield layer may be formed by (for example, zirconia).

The TBC layer has the effect of reducing the temperature of the base material heated by the combustion gas and suppressing oxidative thinning. By increasing the thickness of the TBC, the heat-shielding characteristics can be improved and the oxidative thinning of the base material can be reduced. rice field.

In order to solve the above problems, in the method of repairing a gas turbine component according to the embodiment of the present invention, the thinned portion on the surface of the high temperature component is covered with an oxidation-resistant dense metal layer, and deterioration of the TBC layer is prevented. Or provide a method of repairing a suppressed gas turbine component.

Therefore, the method for repairing gas turbine parts according to the present invention is as follows.
It is a repair method for gas turbine parts.
A heat-shielding coating composed of a bonding layer and a ceramic layer formed via the bonding layer is formed on the first surface of the gas turbine component, and at the same time.
It is characterized in that a metal layer having better oxidation resistance than the material forming the gas turbine member is formed on the thinned portion on the second surface of the gas turbine component.

In the embodiment of the present invention, a gas turbine component that has been used for a long period of time, for example, a combustor liner or a transition piece combustor liner or a transition piece component, or a wall thinning portion is repaired, and a TBC that further improves the durability of the gas turbine component. It is a method of repairing. According to such a repair method, a gas turbine component that can be used again for a long time is provided.

The figure which shows the outline of the repair method by an embodiment. The figure which shows the outline of the repair method by an embodiment. The figure which shows the outline of the repair method by an embodiment. The figure which shows the outline of the repair method by an embodiment. The figure which shows the outline of the repair method by an embodiment. FIG. 6A is a perspective view showing the appearance of the combustor liner before repair, and FIG. 6B is a cross-sectional view of the base material of the combustor liner showing the oxidized thinned portion in FIG. 6A in an enlarged manner.

Hereinafter, a method of repairing a gas turbine component according to an embodiment of the present invention will be described with reference to the accompanying drawings.
As shown in FIG. 6, in general, in a gas turbine component 1 such as a combustor liner or a transition piece that has been operated for a long period of time, the outer surface side of the combustor liner is thinned by oxidation. It is probable that such oxidative thinning occurred because the flow of the cooling gas on the outer surface side of the combustor liner 1 was insufficient and the combustion gas on the inner surface side heated the combustor liner 1 more than the initial setting. The heat shield coating (TBC) 3 on the opposite surface of the oxidative thinning portion 1'progresses to deteriorate rapidly due to high temperature heating, and TBC peeling may be observed.

When the method for repairing a gas turbine component according to the embodiment of the present invention is carried out for such a deteriorated turbine component, the outer surface (second surface) and / or the inner surface (first surface) of the gas turbine component is used. Surface) can be cleaned with an organic or inorganic cleaning solution. For example, scale and the like attached to the outer surface (second surface) of the gas turbine component can be removed. Further, in particular, when the turbine component to be repaired has deteriorated TBC, it is preferable to remove all or part (preferably all) of the deteriorated TBC by chemical treatment or physical treatment. .. When the deteriorated metal layer of the turbine component, particularly the thinned portion, is fragile to the extent that it is not suitable for further use, the fragile portion can be removed as necessary.

When repairing such a deteriorated gas turbine component, the TBC on the inner surface (first surface) of the gas turbine component 1 is repaired, and the wall thickness of the outer surface (second surface) of the gas turbine component 1 is reduced. The part will be repaired.

The gas turbine component repaired according to the embodiment of the present invention has a repaired heat shield coating and a metal layer in which the thinned portion is repaired. As shown in FIGS. 1 to 5, the repaired gas turbine component is a ceramic layer formed on the first surface (inner side surface) of the gas turbine component 1 via the bonding layer 2 and the bonding layer 2. It has a heat shield coating 4 made of 3, and has a metal layer 5 in which a metal having excellent oxidation resistance is overlaid on a wall thinning portion 1'on the second surface (outer surface) of the gas turbine component 1. ing.

<Embodiment (Specific Example)>
FIG. 1 shows an outline of a gas turbine component repaired according to an embodiment of the present invention.
The gas turbine component 1 repaired according to the embodiment of the present invention shown in FIG. 1 is
A heat shield coating 4 made of a bonding layer 2 and a ceramic layer 3 formed via the bonding layer 2 is formed on the first surface (inner surface surface) of the gas turbine component 1.
A metal layer 4 having better oxidation resistance than the material forming the gas turbine member is formed on the thinned portion 1'on the second surface (outer surface) of the gas turbine component 1.
The ceramic layer 3 is composed of a plurality of ceramic layers (specifically, two ceramic layers 31 and 32).
The ceramic layer 31 on the bonding layer 2 side is a ceramic layer 31 having a relatively low porosity, and the ceramic layer 32 on the opposite side of the bonding layer 2 is a ceramic layer 32 having a relatively high porosity. ..

That is, in the ceramic layer 3 shown in FIG. 1, the porosity in the vicinity of the gas turbine component 1 side is lower than the porosity in the vicinity of the surface on the opposite side of the gas turbine component 1. The porosity in the vicinity of the interface on the side of the bonding layer 2 of the ceramic layer 3 is preferably 5 to 20%, and the porosity in the vicinity of the surface on the opposite side of the bonding layer 2 is preferably 15 to 30%.

The ceramic layer 3 shown in FIG. 1 is composed of a plurality of ceramic layers (specifically, two ceramic layers 31 and 32) having different porosities stepwise in the thickness direction. The ceramic layer 31 has a porosity of preferably 5 to 20%, and the ceramic layer 32 has a porosity of preferably 15 to 30%.

The ceramic layer 3 may be composed of three or more ceramic layers having porosities that are gradually different in the thickness direction (not shown).

The bond layer 2 constituting the TBC is preferably a material having excellent oxidation resistance such as NiCr or NiAl, and is represented by MCrAlY (M represents Ni and Co), which is a material having a large amount of Cr and Al and having excellent oxidation resistance. Is desirable. The ceramic material is preferably a material having a low thermal conductivity among ceramic materials such as ZrO ₂ and HfO ₂ . ZrO ₂ is preferable because it contains stabilizing elements such as MgO and _{Y2O3 ,} and by adding an oxide of Dy or La that improves mechanical properties, the thermal conductivity can be further reduced. ..

Examples of the TBC construction method include physical vapor deposition (PVD) using an electron beam and the like, and a thermal spraying method in which a heated and melted powder material is projected onto a base material. Therefore, a thermal spraying method with a high film forming speed is desirable. In particular, TBC can be applied by a plasma spraying method suitable for forming a film of a ceramic material having a high melting point.

In the ceramic layer 3 constituting the TBC, a dense layer having a low porosity is formed on the bonding layer side, and a ceramic layer having a higher porosity can be formed on the upper surface of the dense layer. After the operation, the TBC may be peeled off at the interface between the bonding layer and the ceramic layer, and it is preferable to form a metal layer having excellent adhesion in this portion. Since the dense metal layer having a low porosity is easily peeled off due to the influence of the residual stress generated in the metal layer during film formation, it is difficult to thicken the metal layer to improve the heat shielding performance. Since TBC with increased porosity reduces residual stress during film formation, heat insulation can be improved by thickening the ceramic layer. The porosity of the ceramic layer can be adjusted according to the thermal spraying conditions. In the case of the thermal spraying method, the porosity depends on the thermal spraying conditions such as the input power for generating plasma, the amount of working gas, the thermal spraying distance, and the particle size of the thermal spray powder. The rate can be changed.

The thickness of the entire ceramic layer 3 is preferably about 0.5 to 1.0 mm, more preferably about 1.0 to 2.0 mm in order to suppress oxidative thinning of the base material. The thicknesses of the low porosity layer and the high porosity layer are not particularly limited and can be adjusted according to operating conditions such as the combustion gas temperature and the cooling gas temperature, but the residual stress during metal layer formation is reduced. From this point of view, it is desirable that the ceramic layer having a low porosity is about 5 to 20% of the total thickness of the ceramic layer.

Since erosion due to combustion gas may occur on the uppermost surface of the ceramic layer, the ceramic layer having a high porosity may be mechanically thinned. Therefore, as shown in FIG. 2, erosion can be suppressed by forming the ceramic layer 33, which is extremely dense and has vertical cracks in the metal layer.

It is preferable that the repair of the second surface (outer surface) of the gas turbine component 1 is performed after removing the scale on the surface of the whole or a part of the second surface (for example, including the thinned portion 1').

Any method can be adopted as the method for forming the metal layer 5 on the thinned portion 1'. Preferably, it can be formed by a laser overlay method, a high-speed frame spraying method, a plasma spraying method, or a suspension plasma spraying method.

Among these, as shown in FIG. 3, a metal powder 51 capable of forming a metal layer 5 is supplied to a metal layer forming portion, the metal powder 51 is heated and melted by a laser 6, and the metal powder 51 is laminated. A metal layer forming method using a laser overlay including the metal layer is desirable. Further, it is also preferable to form a metal layer by a high-speed flame spraying method in which a metal layer is formed by projecting molten powder into a high-speed combustion gas. Of these, the metal layer formed by laser overlay becomes more dense, and the oxidation resistance of the metal layer can be improved, which is particularly preferable. The metal layer by the plasma spraying method is inferior in denseness to the metal layer by the high-speed flame spraying method, but it is possible to obtain a metal layer with high oxidation resistance by compacting the surface by laser heating after film formation. You can. Among the plasma spraying methods, the suspension plasma spraying method, which uses ultrafine powder dispersed in a solvent, obtains a metal layer that is extremely dense compared to the ordinary plasma spraying method without post-treatment. It is more preferable because it can be used.

In the metal layer 5 having the inclined composition described above, the content (for example, composition) of the metal powder 51 to be supplied depends on the place where the metal powder is supplied so that the metal layer having such an inclined composition is formed. It can be formed by appropriately determining. For example, when forming the metal layer 5 in a portion close to the base material of a gas turbine component, an alloy powder made of an alloy having a relatively low content of Cr and / or Al is supplied to form a metal thin film. The metal layer 5 having a gradient composition can be formed by accumulating the metal thin films formed by sequentially supplying the alloy powder having an gradually increasing content of Cr and / or Al on the metal thin film.

The so-called high-speed flame spraying method (specifically, a thermal spraying method in which a metal layer is formed by projecting molten powder into a high-speed combustion gas) is also useful, but a laser for laminating powder heated and melted by a laser. It is particularly preferable to adopt the overlay method. According to such a laser overlay, a dense metal layer can be formed, and the oxidation resistance of the metal layer can be further improved. Since the surface of the second surface (outer surface) of the gas turbine component has a temperature of 800 ° C. to 1000 ° C. during operation, the metal layer 5 must be dense in order to protect the base material of the gas turbine component 1. Is essential.

According to the laser overlay method, by setting the metal density to 95% or more, sufficient oxidation resistance can be obtained to protect the base material. In the laser overlay method, the heat input to the base material is small, and the metal layer can be constructed without peeling the TBC installed on the inner surface side of the liner, so that the TBC of the part to be repaired is peeled off. If not, the repair man-hours can be reduced.

The composition of the metal forming the metal layer 5 can be determined mainly in consideration of the oxidation resistance for protecting the base material from oxidation. In the embodiment of the present invention, Cr is 10 to 20 wt% and Al is contained in a composition containing an oxidation-resistant metal element such as Al and Cr, preferably a composition containing an oxidation-resistant metal element such as Al and Cr. By using an alloy composition of Ni and Co containing 5 to 15 wt%, it is possible to obtain an alloy composition having oxidation resistance while maintaining mechanical properties close to those of the base material.

It is also effective to use the composition of the MCrAlY alloy (representing M = Ni, Co, Fe, etc.) used as the oxide-resistant metal layer of the gas turbine. Y, Si, and the like contained in the MCrAlY-based alloy help to form a protective oxide scale formed on the surface of the oxidation-resistant alloy, so that the oxidation resistance of the alloy can be improved.

When the amount of wall thinning is large and the part requiring repair is deep, as shown in FIG. 4, components contributing to oxidation resistance such as Cr and Al are directed from the surface of the oxidation-resistant alloy layer toward the substrate side. It is effective to have a gradient composition such that the abundance ratio of Cr and / or Al increases. Ni and Co alloys containing a large amount of Cr, Al, etc. are harder than superalloys used for high-temperature parts, and therefore cracks are likely to occur when major repairs are required. Therefore, the thinned part is repaired with a composition that is equal to or similar to that of the base material with less Al and Cr, and an oxidation-resistant alloy containing a large amount of Al and Cr is applied to the surface side to satisfy the mechanical properties of the repaired part. It is possible to obtain a metal layer having oxidation resistance on the surface while allowing the metal layer to be formed.

If necessary, a shape with improved heat dissipation can be formed on the surface of the metal layer 5. FIG. 5 shows a fin shape with improved heat dissipation. In the case where such a shape with improved heat dissipation is formed, heat dissipation to the cooling medium is promoted and the temperature of the repaired portion is lowered, so that further oxidative thinning can be suppressed. The fin processing method can be molded together with the repaired part during machining or laser overlaying. Not only the shape of the fin but also the shape of the protrusion can be provided to increase the surface area of the cooling surface and reduce the temperature of the repaired portion.

By repairing the oxidized thinned portion of the high temperature component such as the combustor liner and the transition piece in this way, the temperature of the thinned portion can be reduced and the repaired portion can be prevented from being further oxidized.

As the heat dissipation improving process, it is possible to improve the contact area between the built-up metal layer 5 and the cooling medium, and to rectify the cooling medium to control the flow velocity, contact direction, and contact mode of the cooling medium. Any processing method can be adopted. For example, by forming the heat dissipation improving shape 7 on the metal layer 5 and forming it into an uneven shape (preferably a fin shape), heat dissipation to the cooling medium is promoted and the temperature of the repaired portion is lowered. Oxidation thinning can be suppressed. The fin shape can be molded together with the repaired part during machining or laser overlaying. Not limited to the fin shape, it is possible to take a shape that can increase the surface area of the cooling surface and lower the temperature of the gas turbine component, such as by providing a protrusion shape.

As a result, the heat dissipation of the metal layer 5 is improved, and the temperature of the repaired gas turbine component, particularly the metal metal layer 4 and its peripheral portion, can be lowered more efficiently. This makes it possible to effectively prevent or suppress the occurrence and progression of oxidative thinning.

The heat dissipation improving shape is formed, for example, by forming a metal layer 5 as shown in FIG. 5, and then using a mechanical means such as polishing on the surface of the metal layer 5, as shown in FIG. 5, for example. , Can be processed into a fin shape. Further, when the metal layer 5 is formed, the heat dissipation improved shape 7 can be formed by controlling the supply location of the metal powder, the irradiation range of the laser, the irradiation intensity, and the like.

1: Gas turbine parts, 1': Thinned part, 2: Bonding layer, 3: Ceramic layer, 31: Ceramic layer with relatively low porosity, 32: Ceramic layer with relatively high porosity, 33: Vertical Ceramic layer with cracks 4: Thermal barrier coating layer (TBC), 5: Metal layer, 51: Metal powder, 6: Laser, 7: Heat dissipation improved shape

Claims (11)

It is a repair method for gas turbine parts.
A heat-shielding coating composed of a bonding layer and a ceramic layer formed via the bonding layer is formed on the first surface of the gas turbine component, and at the same time.
A method for repairing a gas turbine component, which comprises forming a metal layer having better oxidation resistance than the material forming the gas turbine member on a thinned portion on the second surface of the gas turbine component. The method for repairing a gas turbine portion according to claim 1, wherein the ceramic layer has a pore ratio in the vicinity of the gas turbine component side lower than the pore ratio in the vicinity of the surface on the opposite side of the gas turbine component. The ceramic layer has a porosity of 5 to 20% in the vicinity of the interface on the bonding layer side and a porosity of 15 to 30% in the vicinity of the surface on the opposite side of the bonding layer. The method for repairing the gas turbine section described in 1. The method for repairing a gas turbine portion according to any one of claims 1 to 3, wherein the ceramic layer is composed of a plurality of ceramic layers having a porosity that is stepwise different in the thickness direction. The method for repairing a gas turbine portion according to any one of claims 1 to 4, wherein the ceramic layer has a crack on a surface in contact with the combustion gas flow in a direction along the direction of the combustion gas flow. The method for repairing a gas turbine component according to any one of claims 1 to 5, wherein the metal layer has a density of 95% or more with respect to the theoretical density of the metal. The gas turbine component according to any one of claims 1 to 6, wherein the metal layer contains 10 to 20 wt% of Cr and 5 to 20 wt% of Al, and the balance is an alloy layer having a composition of Ni or Co or an alloy thereof. Repair method. According to claim 7, the metal layer has an inclined composition in which the abundance ratio of Cr and / or Al increases from the base material side of the gas turbine component toward the surface side of the metal layer after repair. The method of repairing the gas turbine parts described. The method for repairing a gas turbine component according to any one of claims 1 to 8, wherein the metal layer is formed by a laser overlay method, a high-speed frame spraying method, a plasma spraying method, or a suspension plasma spraying method. The method for repairing a gas turbine component according to any one of claims 1 to 9, wherein a shape with improved heat dissipation is formed on the surface of the metal layer. The method for repairing a gas turbine component according to any one of claims 1 to 10, wherein the gas turbine component is any one of a combustor liner, a transition piece, a shroud segment, a moving blade, and a stationary blade.

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