WO2016129375A1 - Gas turbine component, intermediate structure of gas turbine component, gas turbine, method of manufacturing gas turbine component, and method of repairing gas turbine component - Google Patents

Abstract

This gas turbine component comprises a metal intermediate structure (110) and a ceramic layer (130). A gas path surface (102) which faces toward a combustion gas flow path (GP) side in a radial direction (Dr), and a pair of side surfaces (103) which oppose one another in a circumferential direction Dc are formed in the intermediate structure (110). The ceramic layer (130) is formed on the gas path surface (102) of the intermediate structure (110). However, the ceramic layer (130) is not formed on a rotationally upstream corner portion (115u) between the rotationally upstream side surface (103u), on the upstream side in the direction of rotation of a rotor, from among the pair of side surfaces (103), and the gas path surface (102).

Description

Gas turbine component, intermediate structure of gas turbine component, gas turbine, method for manufacturing gas turbine component, and method for repairing gas turbine component

The present invention relates to a gas turbine component that defines a combustion gas flow path in a gas turbine, an intermediate structure of the gas turbine component, a gas turbine, a method for manufacturing the gas turbine component, and a method for repairing the gas turbine component.
This application claims priority based on Japanese Patent Application No. 2015-026468 filed in Japan on February 13, 2015, the contents of which are incorporated herein by reference.

The gas turbine includes a compressor that generates compressed air, a combustor that generates combustion gas by burning fuel in the compressed air, and a turbine that is driven by the combustion gas. The turbine includes a turbine rotor that rotates about an axis, and a turbine casing that covers the turbine rotor. The turbine rotor has a rotor shaft extending in the axial direction in which the axis extends with the axis as a center, and a plurality of blade rows fixed to the rotor shaft. The plurality of moving blade rows have a plurality of moving blades that are spaced apart from each other in the axial direction and fixed to the rotor shaft and arranged in the circumferential direction with reference to the axis. The turbine further includes a stationary blade row disposed on the upstream side of each blade row, and a split ring that forms a ring around the axis and faces the blade row in the radial direction.

Of the components that make up the turbine, the moving blades, stationary blades, and split ring are all high-temperature components that are exposed to high-temperature combustion gas. Each of these high-temperature components has a cooling means for reducing damage caused by high-temperature combustion gas. For example, Patent Document 1 discloses a split ring in which a cooling passage is formed.

In addition, the above-mentioned high-temperature component has a thermal barrier coating (Thermal Barrier Coating®: TBC) layer on the surface of the metal base material. This thermal barrier coating layer is described in Patent Document 2, for example. This thermal barrier coating layer has a bond coat layer formed of a metal such as CoNiCrAlY and a ceramic layer formed of a ZrO 2 ceramic. The bond coat layer is formed on the surface of the base material. The ceramic layer is formed on the surface of this bond coat layer. The bond coat layer mainly plays a role of relaxing the difference between the thermal expansion amount of the ceramic layer and the thermal expansion amount of the base material and ensuring the adhesion between the base material and the ceramic layer.

JP 2010-031753 A JP 2007-270245 A

Here, foreign substances such as fine metal powder flow in the flow path through which the combustion gas flows together with the combustion gas. When this foreign matter collides with a high-temperature part, a thinning due to erosion occurs in a part of the ceramic layer in the thermal barrier coating layer forming the surface layer of the high-temperature part. That is, a part of the ceramic layer is peeled off. Once a part of the ceramic layer is peeled off, thinning by erosion proceeds from the peeled part. For this reason, it is necessary to suppress peeling of the ceramic layer.

Therefore, an object of the present invention is to suppress delamination of a ceramic layer in a gas turbine component that defines a combustion gas flow path in a gas turbine.

A gas turbine component as a first aspect according to the invention for achieving the above object is as follows:
In a gas turbine, in a gas turbine component that defines an annular combustion gas flow path centered on an axis, and forms an annular ring body with respect to the axis, a gas path surface facing the combustion gas flow path side, and the axis A metal intermediate structure in which a pair of side surfaces facing each other in the circumferential direction as a reference and a pair of axial end surfaces facing each other in the direction of the axis are formed, and the axis of the pair of side surfaces Of the pair of axial end surfaces upstream of the combustion gas flow path, and a rotation upstream corner portion composed of a rotation upstream side surface and the gas path surface on the upstream side in the rotation direction of the rotor of the gas turbine rotating about A ceramic layer covering the gas path surface of the intermediate structure, leaving at least one of the shaft upstream corner portion formed of the shaft upstream end surface on the side and the gas path surface; Provided.

The erosion speed of a member formed of ceramic, which is a brittle material, gradually increases as the collision angle of foreign matter on the surface approaches 90 °. On the other hand, the erosion speed of the member formed of a metal that is a ductile material gradually decreases as the collision angle of foreign matter or the like with the surface approaches 90 °. For this reason, when the collision angle is close to 90 °, the erosion speed of the metal member is much lower than that of the ceramic member.

In gas turbine parts, foreign matter in combustion gas or the like collides with the angle of rotation close to 90 ° at the angle of rotation upstream between the rotation upstream side surface and the gas path surface, or at the angle upstream angle portion between the shaft upstream end surface and the gas path surface. To do. In the gas turbine component, the ceramic layer is not formed on the surface of the corner portion of the intermediate structure that becomes the corner portion of the gas turbine component, and the metal is exposed. For this reason, in the gas turbine component, erosion at the corners can be suppressed. Furthermore, in the gas turbine component, the peeling of the ceramic layer due to erosion can be suppressed.

A gas turbine component as a second aspect according to the invention for achieving the above-described object is:
In the gas turbine component of the first aspect, the ceramic layer covers at least the rotation upstream corner portion and covers the gas path surface, and the rotation upstream corner portion rotates toward the combustion gas flow path side. A rotating upstream taper surface that is inclined downstream in the direction may be formed.

A gas turbine component as a third aspect according to the invention for achieving the above object is as follows:
In the gas turbine component of the first or second aspect, the ceramic layer covers at least the shaft upstream corner portion and covers the gas path surface, and the shaft upstream corner portion is disposed on the combustion gas flow path side. An axial upstream taper surface that inclines toward the downstream side of the combustion gas flow path in the direction of the axis as it goes may be formed.

The taper surface in the corner of the gas turbine part is a metal surface. A foreign object or the like collides with the tapered surface at a collision angle close to 90 °. As described above, the erosion speed of the member formed of a metal that is a ductile material gradually decreases as the collision angle of foreign matter or the like with the surface approaches 90 °. For this reason, in the gas turbine component, erosion at the corners can be suppressed.

A gas turbine component as a fourth aspect according to the invention for achieving the above-described object is:
In the gas turbine component of the first or second aspect, the ceramic layer may cover the gas path surface while leaving the shaft upstream corner and the rotation upstream corner.

A gas turbine component as a fifth aspect according to the invention for achieving the object is as follows:
The gas turbine component according to any one of the first to fourth aspects, further comprising a metal bond coat layer that relaxes a difference in thermal expansion amount of the ceramic layer with respect to the intermediate structure, and the rotation upstream corner portion and the rotation The bond coat layer may be formed on at least one of the upstream corners.

The bond coat layer is formed of a metal that relaxes the difference in thermal expansion of the ceramic layer relative to the intermediate structure. This metal is a metal with high oxidation resistance. For this reason, in the gas turbine component, oxidation of the corner portion where the metal is exposed can be suppressed.

A gas turbine component as a sixth aspect according to the invention for achieving the above-described object is:
In the gas turbine component according to any one of the first to fifth aspects, the intermediate structure has a cooling air hole for ejecting air into the combustion gas flow path from the rotation upstream corner or the shaft upstream corner. May be formed.

In the gas turbine component, the corner can be cooled by the cooling air flowing through the cooling air hole. Moreover, in the gas turbine component, the collision of the foreign matter to this portion can be suppressed by the cooling air ejected from the cooling air hole, and the erosion of this portion can be suppressed.

A gas turbine component as a seventh aspect according to the invention for achieving the object is as follows:
In the gas turbine component according to any one of the first to sixth aspects, a split ring that radially faces a moving blade of the gas turbine and defines an outer peripheral side of the annular combustion gas passage may be formed. Good.

A gas turbine component as an eighth aspect according to the invention for achieving the above-described object is:
In the gas turbine component according to any one of the first to sixth aspects, an outer peripheral side in the annular combustion gas flow path may be defined to form an outer shroud in a stationary blade of the gas turbine.

A gas turbine as a ninth aspect according to the invention for achieving the above-described object is:
The gas turbine component according to any one of the first to eighth aspects, and the rotor.

An intermediate structure of a gas turbine component as a tenth aspect according to the invention for achieving the above-described object,
In the gas turbine, an intermediate structure that defines an annular combustion gas flow path centered on an axis and forms an annular ring body with the axis as a reference, a gas path surface that is formed of metal and faces the combustion gas flow path side A pair of side surfaces facing each other in the circumferential direction with respect to the axis, and a pair of axial end surfaces facing each other in the direction of the axis, the gas path surface being an edge side gas path surface not covered with a ceramic layer A rotary upstream side surface on the upstream side in the rotational direction of the rotor of the gas turbine rotating about the axis, and the gas path surface, of the pair of side surfaces, A rotary upstream corner portion, and, of the pair of axial end surfaces, an axial upstream corner portion composed of an upstream shaft end surface on the upstream side of the combustion gas flow path and the gas path surface; Among them, the gas path surface included in at least one corner portion forms the edge gas path surface, the remaining gas path surface forms the main gas path surface, and the edge gas path surface is closer to the axis than the main gas path surface. It is located in the said combustion gas flow path side in the radial direction with respect to.

A gas turbine component manufacturing method as an eleventh aspect according to the invention for achieving the above-described object is as follows:
In the gas turbine, in the method for manufacturing a gas turbine component in which an annular combustion gas flow path is defined with an axis as a center, and an annular ring body is configured with the axis as a reference, the combustion gas flow path side in a radial direction with respect to the axis A metal intermediate structure in which a gas path surface facing each other, a pair of side surfaces facing each other in the circumferential direction with respect to the axis, and a pair of axial end surfaces facing each other in the direction of the axis is manufactured An intermediate structure manufacturing step, a gas path surface, a corner surface of the pair of side surfaces and the gas path surface, and a corner surface of the pair of axial end surfaces and the gas path surface. A covering step of covering with a gas, and a gas upstream surface and a gas upstream surface of the pair of side surfaces that are upstream of the rotation direction of the rotor of the gas turbine rotating about the axis. At least one corner portion of the rolling upstream corner portion and the shaft upstream corner portion formed of the gas upstream end surface and the gas path surface on the upstream side of the combustion gas flow path among the pair of axial end surfaces. A partial removal step of removing the ceramic layer.

A gas turbine component repair method as a twelfth aspect according to the invention for achieving the above-described object is as follows.
In a gas turbine, in a method for repairing a gas turbine component in which an annular combustion gas flow path is defined with an axis as a center and an annular ring body is configured with the axis as a reference, a ceramic layer is formed facing the combustion gas flow path side A metal intermediate structure in which a gas path surface, a pair of side surfaces facing each other in a circumferential direction with respect to the axis, and a pair of axial end surfaces facing each other in the direction of the axis are formed , A first removal step of removing the ceramic layer, the gas path surface, the corner surfaces of the pair of side surfaces and the gas path surface, and the corner surfaces of the pair of axial end surfaces and the gas path surface A covering step of covering with a ceramic layer, and a rotation upstream side surface on the upstream side in the rotation direction of the rotor of the gas turbine rotating about the axis among the pair of side surfaces; A rotary upstream corner portion composed of the gas path surface, and an axial upstream corner portion composed of the gas path surface and the upstream shaft end surface on the upstream side of the combustion gas flow path among the pair of axial end surfaces. A second removal step of removing the ceramic layer at at least one corner.

According to one aspect of the present invention, the peeling of the ceramic layer in the gas turbine component can be suppressed.

It is a principal part notch whole side view of the gas turbine in one Embodiment concerning this invention. It is principal part sectional drawing of the gas turbine in one Embodiment which concerns on this invention. It is a perspective view of a division ring as a gas turbine part in a first embodiment concerning the present invention. It is the figure which looked at the division | segmentation ring in 1st embodiment which concerns on this invention from radial direction inner side. FIG. 5 is a cross-sectional view taken along line VV in FIG. 4. FIG. 5 is a sectional view taken along line VI-VI in FIG. 4. It is a flowchart which shows the manufacture procedure of the split ring as gas turbine components in 1st embodiment which concerns on this invention. It is a flowchart which shows the repair procedure of the division | segmentation ring as gas turbine components in 1st embodiment which concerns on this invention. It is a graph which shows the relationship between the collision angle of a foreign material etc., and the erosion speed. It is a perspective view of the outer side shroud as a gas turbine component in 2nd embodiment which concerns on this invention. It is the figure which looked at the outer side shroud in 2nd embodiment which concerns on this invention from radial direction inner side. It is principal part sectional drawing of the outer side shroud in 2nd embodiment which concerns on this invention. It is principal part sectional drawing of the gas turbine components in the 1st modification which concerns on this invention. It is principal part sectional drawing of the gas turbine components in the 2nd modification which concerns on this invention. It is principal part sectional drawing of the gas turbine components in the 3rd modification which concerns on this invention. It is principal part sectional drawing of the gas turbine components in the 4th modification which concerns on this invention.

Hereinafter, various embodiments according to the present invention will be described in detail with reference to the drawings.

"Embodiment of gas turbine"
An embodiment of a gas turbine according to the present invention will be described with reference to FIGS. 1 and 2.

As shown in FIG. 1, the gas turbine according to the present embodiment includes a compressor 10 that compresses air, a combustor 20 that generates combustion gas by burning fuel in the compressed air compressed by the compressor 10, and And a turbine 30 driven by combustion gas.

The compressor 10 includes a compressor rotor 11 that rotates about an axis Ar, and a compressor casing 15 that covers the compressor rotor 11. The turbine 30 includes a turbine rotor 31 that rotates about an axis Ar and a turbine casing 35 that covers the turbine rotor 31. The combustor 20 is fixed to the turbine casing 35.

The compressor rotor 11 and the turbine rotor 31 are located on the same axis Ar and connected to each other to form the gas turbine rotor 1. The compressor casing 15 and the turbine casing 35 are connected to each other to form the gas turbine casing 5. In the following description, the direction in which the axis Ar extends is the axial direction Da, the side where the compressor 10 is present with respect to the turbine 30 in the axial direction Da is the shaft upstream side Dau, and the opposite side is the shaft downstream side Dad. Further, in the circumferential direction Dc with respect to the axis Ar, the upstream side in the rotational direction of the rotating gas turbine rotor 1 is defined as a rotational upstream side Dcu, and the opposite side is defined as a rotational downstream side Dcd. Further, the radial direction with respect to the axis Ar is simply referred to as a radial direction Dr.

As shown in FIG. 2, the turbine rotor 31 has a rotor shaft 32 extending in the axial direction Da around the axis line Ar, and a plurality of rotor blade rows 33 attached to the rotor shaft 32. The rotor blade rows 33 are attached to the rotor shaft 32 so as to be separated from each other in the axial direction Da. Each of the blade rows 33 is composed of a plurality of blades 34 that are separated from each other in the circumferential direction Dc. The moving blade 34 has a wing body 34a extending in the radial direction Dr, and a platform 34b provided inside the wing body 34a in the radial direction Dr.

The turbine 30 further includes a stationary blade row 43 disposed on the axial upstream side Dau of each rotor blade row 33, and a plurality of split rings 53 facing the rotor blade row 33 in the radial direction Dr. The stationary blade row 43 has a plurality of stationary blades 44 that are separated from each other in the circumferential direction Dc. The stationary blade 44 includes a blade body 44a extending in the radial direction Dr, an inner shroud 44b provided on the inner side in the radial direction Dr of the blade body 44a, and an outer shroud provided on the outer side in the radial direction Dr of the blade body 44a. 44c. The inner shrouds 44b and the outer shrouds 44c of the plurality of stationary blades 44 are arranged in the circumferential direction Dc to form an annular ring body around the axis Da. The plurality of split rings 53 are also arranged in the circumferential direction Dc to form an annular ring centered on the axis Da.

In the turbine casing 35, an annular combustion gas flow path GP is formed around the axis Ar. The annular combustion gas flow path GP is defined by the outer shroud 44 c of the stationary blade 44 and the split ring 53 on the outer peripheral side, and by the inner shroud 44 b of the stationary blade 44 and the platform 34 b of the moving blade 34 on the inner peripheral side. The combustion gas G from the combustor 20 flows through the combustion gas flow path GP. The moving blades 34, the stationary blades 44, and the split ring 53 are all gas turbine components that are exposed to the high-temperature combustion gas G.

“First Embodiment of Gas Turbine Parts”
A first embodiment of a gas turbine component according to the present invention will be described with reference to FIGS. The gas turbine component of this embodiment shown in FIG. 3 is the above-described split ring.

As shown in FIG. 3, the split ring 100 of the present embodiment is configured to fix the flow path forming body 101 spreading in the axial direction Da and the circumferential direction Dc, and the flow path forming body 101 to the inner peripheral side of the turbine casing 35. And a hook 105. The hook 105 is provided outside the flow path forming body 101 in the radial direction Dr.

As shown in FIGS. 5 and 6, the split ring 100 includes an intermediate structure 110 formed of a base material such as a nickel-based alloy, and a metal bond coat layer 120 that covers a part of the surface of the intermediate structure 110. And a ceramic layer 130 that covers a part of the surface of the bond coat layer 120. The bond coat layer 120 and the ceramic layer 130 covering a part of the surface of the intermediate structure 110 form a thermal barrier coating layer. The bond coat layer 120 has high oxidation resistance and reduces the difference between the thermal expansion amount of the ceramic layer 130 and the thermal expansion amount of the intermediate structure 110, and the adhesion between the intermediate structure 110 and the ceramic layer 130. It is made of a material that can secure the properties. For this reason, the bond coat layer 120 is formed of, for example, an MCrAlY alloy. Here, M is a metal containing at least one metal element among Ni, Co, and Fe. The ceramic layer 130 is made of, for example, a ZrO 2 -based ceramic. In FIG. 3 to FIG. 6, the region with a plurality of dots is the region where the ceramic layer 130 is formed.

As shown in FIGS. 3 and 4, the intermediate structure 110 includes a pair of gas path surfaces 102 facing the combustion gas flow path GP in the radial direction Dr, that is, facing the inner side in the radial direction Dr, and a pair facing each other in the circumferential direction Dc. A side surface 103 and a pair of axial end surfaces 104 facing each other in the axial direction Da are formed. The bond coat layer 120 is formed on the gas path surface 102 and is formed on the gas path surface 102 side in the radial direction Dr within the pair of side surfaces 103 and the pair of axial end surfaces 104.

Here, of the pair of side surfaces 103 of the intermediate structure 110, the side surface 103 located on the rotation upstream side Dcu is defined as a rotation upstream side surface 103u, and the opposite side is defined as a rotation downstream side surface 103d. Of the pair of axial end surfaces 104 of the intermediate structure 110, the axial end surface 104 located on the axial upstream side Dau is defined as the axial upstream end surface 104u, and the opposite side is defined as the axial downstream end surface 104d. Further, the corner between the gas path surface 102 and the rotation upstream side surface 103u is the rotation upstream corner portion 115u, the corner between the gas path surface 102 and the rotation downstream side surface 103d is the rotation downstream corner portion 115d, and the gas path surface 102 and the shaft upstream end surface 104u are The corner is defined as the shaft upstream corner 116u, and the corner between the gas path surface 102 and the shaft downstream end surface 104d is defined as the shaft downstream corner 116d.

As shown in FIGS. 5 and 6, the gas path surface 102 of the intermediate structure 110 includes an edge side gas path surface 112b included in the rotation upstream corner portion 115u and the shaft upstream corner portion 116u, and a gas path surface 102 excluding the edge side gas path surface 112b. And a main gas path surface 112a. The edge side gas path surface 112b is located closer to the combustion gas flow path GP than the main gas path surface 112a, that is, inside the radial direction Dr. More specifically, in the present embodiment, the edge side gas path surface 112b of the intermediate structure 110 is positioned on the inner side in the radial direction Dr by the thickness Tc of the ceramic layer 130 from the main gas path surface 112a.

Here, the gas path surface 102 and the corners formed on each surface adjacent to the gas path surface 102 will be described.

As shown in FIG. 5, the rotational upstream corner 115u of the present embodiment is defined with reference to the intersection line I between the edge side gas path surface 112b forming the corner 115u and the rotational upstream side 103u adjacent thereto. . The rotation upstream corner portion 115u is on the gas path surface 102, the portion from the intersection line I to the position of the predetermined distance Lc, and the rotation upstream side surface 103u adjacent to the gas path surface 102, and the intersection line I To the position of the distance Lc. Similarly, as shown in FIG. 6, the axial upstream corner portion 116u of the present embodiment is defined on the basis of the line of intersection between the edge gas path surface 112b and the axial upstream end surface 104u forming the corner portion 116u. That is, the axial upstream corner portion 116u of the present embodiment includes a portion from this intersection line to a position at a predetermined distance on the gas path surface 102, and a portion from this intersection line to a position at a predetermined distance on the shaft upstream end surface 104u. It is a part formed by.

The bond coat layer 120 and the ceramic layer 130 are formed on the entire main gas path surface 112a of the intermediate structure 110. As shown in FIG. 5, the ceramic layer 130 is not formed on the surface of the rotation upstream corner portion 115u of the intermediate structure 110, and the bond coat layer 120 is formed. That is, the bond coat layer 120 is exposed on the edge side gas path surface 112b and the rotation upstream side surface 103u included in the rotation upstream corner portion 115u. An opening of a cooling air hole 119 is formed on the surface of the rotary upstream corner portion 115u. The cooling air hole 119 extends in a direction having a radial Dr component, and jets cooling air existing outside the radial ring Dr of the split ring 100 into the combustion gas flow path GP from the aforementioned opening.

Further, as shown in FIG. 6, the ceramic layer 130 is not formed on the surface of the axial upstream corner portion 116u of the intermediate structure 110, and the bond coat layer 120 is formed. That is, the bond coat layer 120 is exposed on the edge side gas path surface 112b and the shaft upstream end surface 104u included in the shaft upstream corner portion 116u. On the other hand, a bond coat layer 120 and a ceramic layer 130 are formed on the surface of the shaft downstream corner portion 116d and the surface of the rotating downstream corner portion 115d of the intermediate structure 110.

In this embodiment, the second embodiment to be described later, and various modifications, a ceramic layer is not formed on the axial upstream corner and the rotational upstream corner of the intermediate structure. Therefore, the shaft upstream corner and the rotation upstream corner of the intermediate structure in the present embodiment, the second embodiment to be described later, and various modifications are the shaft upstream corner or the rotation upstream corner of the gas turbine component that is a finished product. But there is. Therefore, a metal member is exposed on the surface of the shaft upstream corner or the rotary upstream corner.

The thickness Tb2 of the portion of the bond coat layer 120 whose surface is not covered with the ceramic layer 130 may be thicker than the thickness Tb1 of the portion of the bond coat layer 120 whose surface is covered with the ceramic layer 130.

Next, the manufacturing procedure of the split ring 100 described above will be described according to the flowchart shown in FIG.

First, the intermediate structure 110 is manufactured (S1: intermediate structure manufacturing process). In this intermediate structure manufacturing step (S1), for example, the base material such as a nickel-base alloy is formed by casting or the like so that it has a substantially desired shape. Subsequently, it is machined as necessary to adjust the shape. Next, the bond coat layer 120 is formed on the gas path surface 102, and the bond coat layer 120 is formed on the gas path surface 102 side in the radial direction Dr within the pair of side surfaces 103 and the pair of axial end surfaces 104 (S2: Bond coat layer forming step). The bond coat layer 120 is formed by spraying the sprayed powder such as the aforementioned MCrAlY alloy on the surface of the intermediate structure 110. The bond coat layer 120 may be formed by welding an MCrAlY alloy or the like to the surface of the intermediate structure 110.

Next, the above-described ceramic spray powder such as ZrO 2 is sprayed on the surface of the intermediate structure 110 including the portion where the ceramic layer 130 is formed (S3: coating step). At this time, the ceramic spray powder is sprayed so that the thickness of the ceramic layer 130 is slightly larger than the target thickness.

Next, the surface layer portion of the ceramic layer 130 formed in the covering step (S3) is removed by grinding or polishing (S4: partial removal step). In the covering step (S3), the ceramic layer is not only formed on the main gas path surface 112a that is the surface on which the ceramic layer 130 is to be formed among the surfaces of the intermediate structure 110, but also on the edge side gas path surface 112b on which the ceramic layer 130 is not formed. 130 is formed. Therefore, here, the surface layer portion of the ceramic layer 130 formed in the coating step (S3) is ground or polished until the surface of the bond coat layer 120 forming the edge gas path surface 112b is exposed.

As described above, in the present embodiment, after the ceramic layer 130 is formed to be thicker than the target thickness in the covering step (S3), the bond coat layer 120 that forms the edge side gas path surface 112b in the partial removal step (S4). The surface layer portion of the ceramic layer 130 formed in the coating step (S3) is ground or polished until the surface is exposed. For this reason, in this embodiment, the surface of the ceramic layer 130 on the main gas path surface 112a and the surface of the bond coat layer 120 forming the edge side gas path surface 112b can be easily made flush.

Next, the repair procedure of the split ring 100 will be described according to the flowchart shown in FIG.

First, the surface of the split ring 100 to be repaired is ground or polished to remove at least the ceramic layer 130 (S11: first removal step). In the first removal step (S11), all or part of the bond coat layer 120 may be removed along with the removal of the ceramic layer 130.

Next, the bond coat layer 120 is additionally formed on the surface of the split ring 100 from which the ceramic layer 130 has been removed, and the ceramic layer 130 is formed thereon (S12: coating step). When the ceramic layer 130 is formed, the ceramic spray powder is sprayed so that the thickness of the ceramic layer 130 is slightly larger than the target thickness, as in the above-described coating step (S3).

Next, the surface layer portion of the ceramic layer 130 formed in the covering step (S12) is removed by grinding or polishing (S13: second removal step). In the second removal step (S13), the ceramic layer 130 formed in the covering step (S12) is exposed until the surface of the bond coat layer 120 forming the edge gas path surface 112b is exposed, as in the partial removal step (S4) described above. Grind or polish the surface layer. Therefore, the surface of the ceramic layer 130 on the main gas path surface 112a and the surface of the bond coat layer 120 on the edge side gas path surface 112b can be easily flushed.

Next, the function and effect of the split ring 100 described above will be described.

As shown in FIG. 4, in the combustion gas flow path GP, the flow G1 of the combustion gas G flowing from the stationary blade 44 to the moving blade 34 is a flow toward the rotating downstream side Dcd while moving toward the axial downstream side Dad. For this reason, when the foreign gas such as metal powder is included in the combustion gas G, the foreign material collides with the portion on the rotary upstream side Dcu of the split ring 100 at an angle close to 90 °.

Further, there is an outer shroud 44c of the stationary blade 44 disposed on the shaft upstream side Dau of the split ring 100 that ejects cooling air to the shaft downstream side Dad. The cooling air may also contain foreign matters such as metal powder. In this case, the foreign matter in the cooling air collides with the portion upstream of the axis Dau of the split ring 100 at an angle close to 90 °.

In addition, there is another split ring 100 disposed on the rotary upstream side Dcu of the split ring 100 that ejects cooling air from the rotary downstream side surface 103d to the rotary downstream side Dcd. The cooling air may also contain foreign matters such as metal powder. In this case, the foreign matter in the cooling air collides with the rotation upstream side Dcu of the split ring 100 at an angle close to 90 °.

For this reason, if the ceramic layer 130 is formed on the axial upstream side Dau portion and the rotational upstream side Dcu portion of the split ring 100, a part of the ceramic layer 130 is peeled off by the foreign matter colliding with these portions. The possibility to do increases. Once a part of the ceramic layer 130 is peeled off, thinning by erosion proceeds from the peeled part.

Therefore, in the present embodiment, the axial upstream side Dau portion and the rotational upstream side Dcu portion of the split ring 100, more specifically, the surface of the shaft upstream corner portion 116u in the intermediate structure 110 of the split ring 100, and this The metal bond coat layer 120 is exposed without forming the ceramic layer 130 on the surface of the rotation upstream corner portion 115u of the intermediate structure 110.

As shown in FIG. 9, the erosion speed of the member formed of ceramics, which is a brittle material, gradually increases as the collision angle of foreign matter on the surface approaches 90 °. For this reason, if the ceramic layer 130 is formed on the surface of the shaft upstream corner portion 116u and the surface of the rotating upstream corner portion 115u in the split ring 100 where the collision angle of the foreign matter is close to 90 °, the erosion speed of this portion is increased. It becomes higher than the erosion speed of other parts. On the other hand, a member formed of a metal that is a ductile material has a high erosion speed when the collision angle of foreign matter on the surface is about 20 °, but gradually increases as the collision angle approaches from 90 ° to 90 °. The erosion speed decreases. For this reason, when the collision angle is close to 90 °, the erosion speed of the metal member is much lower than that of the ceramic member.

Therefore, in the present embodiment, as described above, the ceramic layer 130 is not formed on the surface of the shaft upstream corner portion 116u and the surface of the rotating upstream corner portion 115u in the split ring 100, and a metal bond coat is formed. Since the layer 120 is exposed, erosion of this portion can be suppressed. Furthermore, in this embodiment, peeling of the ceramic layer 130 due to erosion can be suppressed.

In the present embodiment, the bond coat layer 120 exposed on the surface of the shaft upstream corner portion 116u and the surface of the rotating upstream corner portion 115u in the split ring 100 is formed of a metal having high oxidation resistance. , Oxidation at this portion can be suppressed. Moreover, in this embodiment, the thickness Tb2 of the bond coat layer 120 exposed on the surface of the shaft upstream corner portion 116u and the surface of the rotation upstream corner portion 115u is a bond in which the ceramic layer 130 is formed on the surface. Since it is thicker than the thickness Tb1 of the coat layer 120, the oxidation resistance can be further enhanced.

Furthermore, in this embodiment, the opening of the cooling air hole 119 is formed on the surface of the rotation upstream corner portion 115u, and the cooling air is ejected from this opening, so that this portion where the ceramic layer 130 is not formed can be cooled. it can. In this embodiment, the opening of the cooling air hole 119 is formed on the surface of the rotating upstream corner portion 115u. However, the opening of the cooling air hole may be formed on the surface of the shaft upstream corner portion 116u. In this embodiment, the ceramic layer 130 is not formed on the entire axial direction Da in the rotational upstream corner portion 115u, and the bond coat layer 120 is exposed, but the axial direction Da in the rotational upstream corner portion 115u is exposed. The bond coat layer 120 may be exposed on the upstream side, and the ceramic layer 130 may be formed on the downstream side.

Here, the combustion gas flow G1 and the combustion gas flow G2 shown in FIG. 4 will be described. The speed of the circumferential direction Dc component of the flow G2 of the combustion gas flowing from the moving blade 34 to the stationary blade 44 is smaller than the speed of the circumferential direction Dc component of the flow G1 of the combustion gas G flowing from the stationary blade 44 to the moving blade 34. For this reason, the collision speed of the foreign object with respect to the area | region of the shaft downstream side Dad in the rotation upstream corner part 115u is smaller than the collision speed of the foreign object with respect to the area | region of the shaft upstream side Dau in this rotation upstream corner part 115u. Therefore, as described above, the ceramic layer 130 may be applied to the shaft downstream side Dad in the rotation upstream corner portion 115u.

Further, in this embodiment, the ceramic layer is not formed at both corners of the shaft upstream corner and the rotation upstream corner, and the metal that is a ductile material is exposed, but the shaft upstream corner and the rotation upstream corner are exposed. Among the corners, a ceramic layer may be formed at one corner, and a metal that is a ductile material may be exposed at the other corner. For example, a ceramic layer may be formed at the shaft upstream corner, and a metal that is a ductile material may be exposed at the rotation upstream corner.

“Second Embodiment of Gas Turbine Parts”
A second embodiment of the gas turbine component according to the present invention will be described with reference to FIGS. The gas turbine component of this embodiment is an outer shroud of a stationary blade.

As shown in FIG. 10, the outer shroud 200 of the present embodiment includes a flow path forming body 201 extending in the axial direction Da and the circumferential direction Dc, and a hook for fixing the outer shroud 200 to the inner peripheral side of the turbine casing 35. 205. The hook 205 is provided outside the flow path forming body 201 in the radial direction Dr. The vane body 44 a of the stationary blade is provided inside the flow path forming body 201 in the radial direction Dr.

Similarly to the above-described split ring 100, the outer shroud 200 is also made of an intermediate structure 210 made of metal and a metal bond coat layer covering a part of the surface of the intermediate structure 210, as shown in FIG. 120 and a ceramic layer 130 covering a part of the surface of the intermediate structure 210. The bond coat layer 120 and the ceramic layer 130 covering a part of the surface of the intermediate structure 210 form a thermal barrier coating layer.

The intermediate structure 210 includes a gas path surface 202 that faces the combustion gas flow path GP in the radial direction Dr, that is, an inner side in the radial direction Dr, a pair of side surfaces 203 that face each other in the circumferential direction Dc, and that faces each other in the axial direction Da. And a pair of axial end faces 204 are formed. The bond coat layer 120 is formed on the gas path surface 202 and is formed on the gas path surface 202 side in the radial direction Dr in the pair of side surfaces 203 and the pair of axial end surfaces 204.

Here, of the pair of side surfaces 203 of the intermediate structure 210, the side surface 203 located on the rotation upstream side Dcu is defined as a rotation upstream side surface 203u, and the opposite side is defined as a rotation downstream side surface 203d. Of the pair of axial end surfaces 204 of the intermediate structure 210, the axial end surface 204 located on the axial upstream side Dau is defined as the axial upstream end surface 204u, and the opposite side is defined as the axial downstream end surface 204d. Further, the corner between the gas path surface 202 and the rotation upstream side surface 203u is the rotation upstream corner 215u, the corner between the gas path surface 202 and the rotation downstream side 203d is the rotation downstream corner 215d, and the gas path surface 202 and the shaft upstream end surface 204u are The corner is defined as the shaft upstream corner 216u, and the corner between the gas path surface 202 and the shaft downstream end surface 204d is defined as the shaft downstream corner 216d.

The gas path surface 202 of the intermediate structure 210 includes an edge side gas path surface 212b included in the axial upstream corner portion 216u and the rotational upstream corner portion 215u, and a main gas path surface 212a that is the gas path surface 202 excluding the edge side gas path surface 212b. . The edge-side gas path surface 212b is positioned on the combustion gas flow path GP side, that is, on the radial direction Dr side from the main gas path surface 212a by approximately the thickness Tc of the ceramic layer 130.

The bond coat layer 120 and the ceramic layer 130 are formed on the entire main gas path surface 212a of the intermediate structure 210. The ceramic layer 130 is not formed on the surface of the rotation upstream corner portion 215u of the intermediate structure 210, and the bond coat layer 120 is formed. That is, the bond coat layer 120 is exposed on the edge side gas path surface 212b and the rotation upstream side surface 203u included in the rotation upstream corner portion 215u. The bond coat layer 120 is also exposed on the surface of the axial upstream corner 216u of the intermediate structure 210. That is, the ceramic layer 130 is not formed on the edge side gas path surface 212b and the shaft upstream end surface 204u included in the shaft upstream corner portion 216u, and the bond coat layer 120 is formed. On the other hand, a bond coat layer 120 and a ceramic layer 130 are formed on the surface of the shaft downstream corner 216d and the surface of the rotation downstream corner 215d of the intermediate structure 210.

Note that the thickness of the ceramic layer 130 and the thickness of the bond coat layer 120 in this embodiment are the same as those in the first embodiment. The definition of the corner in this embodiment is the same as the definition of the corner in the first embodiment.

The outer shroud 200 described above is manufactured in the same manner as the method for manufacturing the split ring 100 described above. Further, the outer shroud 200 is repaired in the same manner as the repair method of the split ring 100 described above. However, the outer shroud 200 is not manufactured or repaired alone, but is manufactured or repaired integrally with the wing body and the inner shroud.

Next, the function and effect of the outer shroud 200 described above will be described.

As shown in FIG. 11, the combustion gas flow G2 flowing from the moving blade 34 to the stationary blade 44 in the combustion gas flow path GP is a flow in the circumferential direction Dc while moving toward the axial downstream side Dad. For this reason, when the foreign gas such as metal powder is contained in the combustion gas G, the foreign matter collides with the shaft upstream side Dau portion of the outer shroud 200 at an angle close to 90 °. In the combustion gas flow path GP, the combustion gas flow G1 flowing from the stationary blade 44 to the moving blade 34 is a flow toward the rotating downstream side Dcd while moving toward the axial downstream side Dad. For this reason, when the foreign gas such as metal powder is included in the combustion gas G, the foreign matter collides with the portion of the outer shroud 200 on the rotational upstream side Dcu at an angle close to 90 °.

Also, some of the split rings 54 on the shaft upstream side Dau of the outer shroud 200 eject cooling air to the shaft downstream side Dad. The cooling air may also contain foreign matters such as metal powder. In this case, the foreign matter in the cooling air collides with the axial upstream side Dau of the outer shroud 200 at an angle close to 90 °.

In addition, some of the other outer shrouds 200 arranged on the rotation upstream side Dcu of the outer shroud 200 eject cooling air from the rotation downstream side surface 203d to the rotation downstream side Dcd. The cooling air may also contain foreign matters such as metal powder. In this case, the foreign matter in the cooling air collides with the rotation upstream side Dcu of the outer shroud 200 at an angle close to 90 °.

For this reason, if the ceramic layer 130 is formed on the axial upstream side Dau portion and the rotational upstream side Dcu portion of the outer shroud 200, a part of the ceramic layer 130 is peeled off by foreign matter colliding with these portions. The possibility to do increases. Therefore, also in the present embodiment, as in the split ring 100 of the first embodiment, the axial upstream side Dau portion and the rotational upstream side Dcu portion of the outer shroud 200, more specifically, the intermediate structure 210 of the outer shroud 200. The metal bond coat layer 120 is exposed without forming the ceramic layer 130 on the surface of the shaft upstream corner 216u and the surface of the rotation upstream corner 215u of the intermediate structure 210.

Therefore, also in this embodiment, like the split ring 100 of the first embodiment, erosion of the shaft upstream corner 216u in the intermediate structure 210 of the outer shroud 200 and the rotation upstream corner 215u in this intermediate structure 210 is suppressed. Can do. Furthermore, also in this embodiment, peeling of the ceramic layer 130 due to erosion can be suppressed.

Also in this embodiment, the bond coat layer 120 exposed on the surface of the shaft upstream corner portion 216u in the intermediate structure 210 of the outer shroud 200 and the surface of the rotation upstream corner portion 215u in the intermediate structure 210 has an acid resistance. Since it is formed of a metal having high chemical properties, oxidation at this portion can be suppressed.

In this embodiment, the ceramic layer 130 is not formed on the entire axial direction Da in the rotational upstream corner portion 215u and the bond coat layer 120 is exposed, but the axial direction Da in the rotational upstream corner portion 215u is exposed. The bond coat layer 120 may be exposed on the downstream side, and the ceramic layer 130 may be formed on the upstream side. The speed of the circumferential direction Dc component of the combustion gas flow G2 shown in FIG. 11 is smaller than the speed of the circumferential direction Dc component of the combustion gas flow G1. For this reason, the collision speed of the foreign object with respect to the area | region of the shaft upstream side Dau in the rotation upstream corner | angular part 215u is smaller than the collision speed of the foreign object with respect to the area | region of the shaft upstream side Dau in this rotation upstream corner | angular part 215u. Therefore, as described above, the ceramic layer 130 may be formed on the shaft upstream side Dau in the rotation upstream corner portion 215u.

In the present embodiment, an opening for the cooling air hole may be formed on the surface of the rotary upstream corner portion 215u or the surface of the shaft upstream corner portion 116u. In each of the modifications described later, an opening for the cooling air hole may be formed on the surface of the rotary upstream corner or the surface of the shaft upstream corner.

"First modification of gas turbine parts"
A first modification of the gas turbine component in the first and second embodiments described above will be described with reference to FIG. In addition, all the gas turbine components in each modified example described below are a split ring of a gas turbine or an outer shroud of a stationary blade.

As in the above embodiments, the gas turbine component 300 of the present modification is also made of a metal that covers an intermediate structure 110h formed of metal and a part of the surface of the intermediate structure 110h, as shown in FIG. The bond coat layer 120h and the ceramic layer 130 covering a part of the surface of the intermediate structure 110h are formed.

The intermediate structure 110h also includes a gas path surface 102 facing the combustion gas flow path GP in the radial direction Dr, a pair of side surfaces 103 facing each other in the circumferential direction Dc, and a pair of axial end surfaces facing each other in the axial direction Da. 104.

The gas path surface 102 includes an edge side gas path surface 112b included in the rotational upstream corner portion 115u and the shaft upstream corner portion 116u of the intermediate structure 110h, and a main gas path surface 112a that is the gas path surface 102 excluding the edge side gas path surface 112b. . The edge-side gas path surface 112b is positioned on the combustion gas flow path GP side, that is, on the radial direction Dr side from the main gas path surface 112a by approximately the thickness Tc of the ceramic layer 130.

The bond coat layer 120h is formed on the main gas path surface 112a of the intermediate structure 110h, the surface of the shaft downstream corner of the intermediate structure 110h, and the surface of the rotation downstream corner of the intermediate structure 110h. The bond coat layer 120h is not formed on the surface of the axial upstream corner portion 116u of the intermediate structure 110h and the surface of the rotational upstream corner portion 115u of the intermediate structure 110h.

The ceramic layer 130 is formed on the entire surface of the bond coat layer 120h, and is not formed in a region where the bond coat layer 120h is not formed. Therefore, the ceramic layer 130 is not formed on the surface of the shaft upstream corner portion 116u of the intermediate structure 110h where the bond coat layer 120h is not formed and the surface of the rotation upstream corner portion 115u of the intermediate structure 110h. Therefore, in this modification, the shaft upstream corner portion 116u of the intermediate structure 110h forms the shaft upstream corner portion of the gas turbine component 300, and the metal intermediate structure 110h is exposed on this surface. In this modification, the rotation upstream corner 115u of the intermediate structure 110h forms the rotation upstream corner of the gas turbine component 300, and the metal intermediate structure 110h is exposed on this surface.

Therefore, even in this modification, it is possible to suppress the peeling of the ceramic layer 130 due to erosion.

"Second modification of gas turbine parts"
A second modification of the gas turbine component in the first and second embodiments described above will be described with reference to FIG.

In each of the above embodiments, the edge side gas path surface 112b of the intermediate structure 110 is positioned closer to the combustion gas flow path GP than the main gas path surface 112a of the intermediate structure 110. That is, the edge side gas path surface 112 b of the intermediate structure 110 is discontinuous with respect to the main gas path surface 112 a of the intermediate structure 110.

In the gas turbine component 400 of this modification, the edge side gas path surface 112bi of the intermediate structure 110i is flush with the main gas path surface 112ai of the intermediate structure 110i. That is, the edge side gas path surface 112bi is continuous with the main gas path surface 112ai. For this reason, in this modification, the main gas path on the gas path surface 402 of the gas turbine component 400 is obtained by making the thickness of the bond coat layer 120i on the edge side gas path surface 112bi of the intermediate structure 110i substantially the thickness of the ceramic layer 130. The surface 402a and the edge side gas path surface 402b are flush with each other. The main gas path surface 402a of the gas turbine component 400 is a gas path surface on the main gas path surface 112ai of the intermediate structure 110i in the gas path surface 402 of the gas turbine component 400. The edge gas path surface 402b of the gas turbine component 400 is a gas path surface on the edge gas path surface 112bi of the intermediate structure 110i in the gas path surface 402 of the gas turbine component 400.

Therefore, in this modification, it is not necessary to discontinuously form the main gas path surface 112ai and the edge gas path surface 112bi in the process of forming the gas path surface 102i of the intermediate structure 110i, and the intermediate structure 110i can be easily formed. In addition, the formation cost of the intermediate structure 110i can be suppressed.

“Third Modification of Gas Turbine Parts”
A third modification of the gas turbine component in the first and second embodiments described above will be described with reference to FIG.

The gas turbine component 500 of this modification is formed by forming tapered surfaces 117 and 118 at the shaft upstream corner portion 116u and the rotating upstream corner portion 115u of the intermediate structure 110j. A bond coat layer 120j is formed on the surfaces of the axial upstream corner portion 116u and the rotational upstream corner portion 115u of the intermediate structure 110j of the present modification, as in the above embodiments. Therefore, the surfaces of the bond coat layer 120j formed on the tapered surfaces 117 and 118 in the axial upstream corner portion 116u and the rotational upstream corner portion 115u of the intermediate structure 110j also form inclined tapered surfaces 117j and 118j in the same manner.

The taper surface 118 of the shaft upstream corner portion 116u gradually moves toward the shaft downstream side Dad as it goes toward the combustion gas flow path GP in the radial direction Dr. The taper surface 118 has a width Wt in the axial direction Da and a width Wt in the radial direction Dr as follows.
0.5Tc ≦ Wt <Lc
Tc is the thickness of the ceramic layer 130, and Lc is the corner width defined in the first embodiment. Further, the base point of the width of the tapered surface 118 is the intersection line I of the extended surfaces of the two surfaces forming the corner. Further, the width Wt in the axial direction Da and the width Wt in the radial direction Dr of the tapered surface 118 may be different from each other as long as the values satisfy the above formula.

Further, the taper surface 117 of the rotation upstream corner portion 115u gradually moves toward the rotation downstream side Dcd as it goes toward the combustion gas flow path GP in the radial direction Dr. The width Wt in the circumferential direction Dc and the width Wt in the radial direction Dr of the tapered surface 117 are the same as the values shown in the above formula.

As described above with reference to FIG. 9, the erosion speed of the member formed of a metal, which is a ductile material, gradually decreases as the collision angle of foreign matter or the like with respect to the surface approaches from 20 ° to 90 °. In the present modification, as described above, the tapered surfaces 117j and 118j are formed on the shaft upstream corner portion 116u and the rotation upstream corner portion 115u of the metallic intermediate structure 110j. The foreign matter collides with the taper surfaces 117j and 118j made of metal at a collision angle close to 90 °. For this reason, in this modification, the erosion in the shaft upstream corner and the rotation upstream corner of the gas turbine component 500 can be suppressed.

In addition, although this modification is an example applied to the gas turbine component of 1st and 2nd embodiment, you may apply to another modification.

"Fourth modification of gas turbine parts"
A fourth modification of the gas turbine component in the first and second embodiments described above will be described with reference to FIG.

In each of the above embodiments, the gas path surface of the gas turbine part is flush with the main gas path surface of the gas turbine part. Note that the definitions of the edge gas path surface of the gas turbine component and the main gas path surface of the gas turbine component are the same as those described in the second modification of the gas turbine component. In the gas turbine component 600 of the present modified example, the edge gas path surface 602b of the gas turbine component 600 is positioned closer to the combustion gas flow path GP in the radial direction Dr than the main gas path surface 602a of the gas turbine component 600. . A bond coat layer 120k is formed on the surfaces of the axial upstream corner portion 116u and the rotational upstream corner portion 115u of the intermediate structure 110k of the present modification, as in the above embodiments, and the bond coat layer 120k is exposed. Yes.

Foreign matter from the shaft upstream side Dau or the rotational upstream side Dcu hardly collides with a portion near the edge side gas path surface 602b on the main gas path surface 602a of the gas turbine component 600 due to the presence of the edge side gas path surface 602b of the gas turbine component 600. . For this reason, in this modification, the erosion of the part near the edge side gas path surface 602b of the ceramic layer 130 can be suppressed.

By the way, the tip clearance which is the distance in the radial direction Dr between the gas path surface of the split ring and the radial direction Dr outer end of the rotor blade facing the split ring in the radial direction Dr is a dimension that affects the performance of the gas turbine. Because it is, it is managed very strictly.

Here, the gas turbine component 600 of the present modification is a split ring, and the edge side gas path surface 602b of the rotation upstream corner portion 115u in this split ring is closer to the combustion gas flow path GP side in the radial direction Dr than the main gas path surface 602a. Consider the case of positioning.

In the above case, the distance in the radial direction Dr between the edge side gas path surface 602b of the rotating upstream corner 115u and the outer end in the radial direction Dr of the rotor blade in the split ring is the distance between the main gas path surface 602a and the outer end in the radial direction Dr of the rotor blade. It becomes narrower than the interval in the radial direction Dr. In this case, when the distance in the radial direction Dr between the edge side gas path surface 602b and the outer end in the radial direction Dr of the moving blade is treated as the above-described tip clearance, the radial direction Dr between the main gas path surface 602a and the outer end in the radial direction Dr of the moving blade. Is wider than the tip clearance, the performance of the gas turbine is reduced. On the other hand, when the distance in the radial direction Dr between the main gas path surface 602a and the outer end in the radial direction Dr of the moving blade is treated as the above-described tip clearance, the radial direction Dr between the edge side gas path surface 602b and the outer end in the radial direction Dr of the moving blade. Is smaller than the tip clearance, there is a possibility that the moving blade contacts the edge side gas path surface 602b.

Therefore, when the gas turbine component 600 of the present modification is a split ring, only the edge side gas path surface 602b of the region that does not face the radial direction Dr end of the rotor blade is selected from the edge side gas path surface 602b of the split ring. The gas path surface 602a is preferably positioned closer to the combustion gas flow path GP in the radial direction Dr. From this viewpoint, the edge-side gas path surface 602b including the axial upstream corner portion 116u of the split ring may be positioned closer to the combustion gas flow path GP in the radial direction Dr than the main gas path surface 602a. Further, with respect to the edge side gas path surface 602b including the rotation upstream corner portion 115u of the split ring, only the portion of the shaft upstream side Dau and the shaft downstream side Dad that are not opposed to the radial direction Dr end of the rotor blade is larger in diameter than the main gas path surface 602a. It may be positioned on the combustion gas flow path GP side in the direction Dr.

In addition, although this modification is an example applied to the gas turbine component of 1st and 2nd embodiment, you may apply to another modification.

"Other variations of gas turbine parts"

In the gas turbine component of each embodiment and each modification described above, the ceramic layer is not formed on both corners of the shaft upstream corner and the rotation upstream corner, and the metal which is a ductile material is exposed. . However, a ceramic layer may be formed at one corner of the shaft upstream corner and the rotation upstream corner, and a metal that is a ductile material may be exposed at the other corner. For example, a ceramic layer may be formed at the shaft upstream corner, and a metal that is a ductile material may be exposed at the rotation upstream corner. Further, a ceramic layer may be formed at the rotation upstream corner portion, and the ductile material metal may be exposed at the shaft upstream corner portion.

The gas turbine parts of the embodiments and the modifications described above are all applied to the gas turbine parts that define the outer peripheral side of the annular combustion gas flow path GP, that is, the outer shroud and the split ring of the stationary blade. This is an applied example. However, the present invention may be applied to gas turbine components that define the inner peripheral side of the annular combustion gas flow path GP, that is, the inner shroud of the stationary blade and the platform of the moving blade.

According to one aspect of the present invention, the peeling of the ceramic layer in the gas turbine component can be suppressed.

1: gas turbine rotor, 5: gas turbine casing, 10: compressor, 11: compressor rotor, 15: compressor casing, 20: combustor, 30: turbine, 31: turbine rotor, 32: rotor shaft, 34: Rotor blade, 34a: Wing body, 34b: Platform, 35: Turbine casing, 44: Static blade, 44a: Wing body, 44b: Inner shroud, 44c, 200: Outer shroud, 53, 100: Split ring, 110, 110h, 110i, 110j, 110k, 210: intermediate structure, 102, 202, 402, 602: gas path surface, 112a, 212a, 402a, 602a: main gas path surface, 112b, 212b, 402b, 602b: edge side gas path surface, 103, 203 : Side surface, 103u, 203u: rotation upstream side surface, 103d, 203d: rotation downstream Side surface 104, 204: axial end surface, 104u, 204u: axial upstream end surface, 104d, 204d: axial downstream end surface, 115u, 215u: rotational upstream corner, 115d, 215d: rotational downstream corner, 116u, 216u: axial upstream Corner portion, 116d, 216d: shaft downstream corner portion, 117, 117j, 118, 118j: taper surface, 120, 120h, 120i, 120j, 120k: bond coat layer, 130: ceramic layer, 300, 400, 500, 600: Gas turbine parts

Claims (12)

In a gas turbine, a gas turbine component that defines an annular combustion gas flow path centering on an axis and that forms an annular ring body based on the axis,
A metal path formed with a gas path surface facing the combustion gas flow path side, a pair of side surfaces facing each other in the circumferential direction with respect to the axis, and a pair of axial end surfaces facing each other in the direction of the axis An intermediate structure;
Of the pair of side surfaces, a rotation upstream corner portion composed of a rotation upstream side surface on the upstream side in the rotation direction of the rotor of the gas turbine rotating about the axis and the gas path surface, and a pair of axial end surfaces Among them, ceramics that covers at least one corner portion of the shaft upstream end portion composed of the shaft upstream end surface on the upstream side of the combustion gas flow path and the gas path surface, and covers the gas path surface of the intermediate structure Layers,
Gas turbine parts comprising. The gas turbine component according to claim 1,
The ceramic layer leaves at least the rotation upstream corner, covers the gas path surface,
A rotational upstream taper surface that is inclined toward the downstream side in the rotational direction as it goes toward the combustion gas flow path side is formed in the rotational upstream corner portion.
Gas turbine parts. The gas turbine component according to claim 1 or 2,
The ceramic layer leaves at least the shaft upstream corner, covers the gas path surface,
An axial upstream taper surface that is inclined toward the downstream side of the combustion gas flow path in the direction of the axis as it goes toward the combustion gas flow path side is formed at the axial upstream corner portion.
Gas turbine parts. In the gas turbine component according to any one of claims 1 to 3,
The ceramic layer covers the gas path surface, leaving the shaft upstream corner and the rotation upstream corner.
Gas turbine parts. In the gas turbine component according to any one of claims 1 to 4,
A metal bond coat layer that relaxes the difference in thermal expansion of the ceramic layer relative to the intermediate structure,
The bond coat layer is formed on at least one of the rotation upstream corner and the rotation upstream corner.
Gas turbine parts. In the gas turbine component according to any one of claims 1 to 5,
The intermediate structure is formed with cooling air holes for ejecting air from the rotation upstream corner or the shaft upstream corner into the combustion gas flow path.
Gas turbine parts. In the gas turbine component according to any one of claims 1 to 6,
Facing the rotor blade of the gas turbine in the radial direction with respect to the axis, and forming a split ring defining an outer peripheral side in the annular combustion gas flow path;
Gas turbine parts. In the gas turbine component according to any one of claims 1 to 6,
Defining an outer peripheral side in the annular combustion gas flow path and forming an outer shroud in a stationary blade of the gas turbine;
Gas turbine parts. A gas turbine component according to any one of claims 1 to 8,
The rotor;
A gas turbine comprising: In the gas turbine, in an intermediate structure that defines an annular combustion gas flow path centering on an axis, and that forms an annular ring on the basis of the axis,
Formed of metal,
A gas path surface facing the combustion gas flow path side, a pair of side surfaces facing each other in a circumferential direction with respect to the axis, and a pair of axial end surfaces facing each other in the direction of the axis;
The gas path surface has an edge gas path surface not covered with a ceramic layer and a main gas path surface covered with a ceramic layer;
Of the pair of side surfaces, a rotation upstream corner portion composed of a rotation upstream side surface on the upstream side in the rotation direction of the rotor of the gas turbine rotating about the axis and the gas path surface, and a pair of axial end surfaces Among these, the gas path surface included in at least one of the shaft upstream end portion composed of the shaft upstream end surface on the upstream side of the combustion gas flow path and the gas path surface forms the edge gas path surface, The remaining gas path surface forms the main gas path surface,
The edge side gas path surface is located closer to the combustion gas flow path side in the radial direction with respect to the axis than the main gas path surface.
Intermediate structure. In the gas turbine, in the method of manufacturing a gas turbine component that defines an annular combustion gas flow path centering on an axis, and forms an annular ring body based on the axis,
A gas path surface facing the combustion gas flow path side in the radial direction with respect to the axis, a pair of side surfaces facing each other in the circumferential direction with respect to the axis, and a pair of axial end surfaces facing each other in the direction of the axis are formed. An intermediate structure manufacturing process for manufacturing a metal intermediate structure,
A covering step of covering the gas path surface, the surface of the corner portion of the pair of side surfaces and the gas path surface, and the surface of the corner portion of the pair of axial end surfaces and the gas path surface with a ceramic layer,
Of the pair of side surfaces, a rotation upstream corner portion composed of a rotation upstream side surface on the upstream side in the rotation direction of the rotor of the gas turbine rotating about the axis and the gas path surface, and a pair of axial end surfaces Among them, a partial removal step of removing the ceramic layer in at least one corner portion of the shaft upstream corner portion composed of the shaft upstream end surface on the upstream side of the combustion gas flow path and the gas path surface;
A method for manufacturing a gas turbine component. In a gas turbine, a method for repairing a gas turbine component that defines an annular combustion gas flow path centering on an axis and that forms an annular ring with respect to the axis,
A gas path surface on which the ceramic layer is formed facing the combustion gas flow path side, a pair of side surfaces facing each other in a circumferential direction with respect to the axis, and a pair of axial end surfaces facing each other in the direction of the axis A first removal step of removing the ceramic layer from the metal intermediate structure formed;
A covering step of covering the gas path surface, the surface of the corner portion of the pair of side surfaces and the gas path surface, and the surface of the corner portion of the pair of axial end surfaces and the gas path surface with a ceramic layer,
Of the pair of side surfaces, a rotation upstream corner portion composed of a rotation upstream side surface on the upstream side in the rotation direction of the rotor of the gas turbine rotating about the axis and the gas path surface, and a pair of axial end surfaces Among them, a second removal step of removing the ceramic layer in at least one corner portion of the shaft upstream end portion composed of the shaft upstream end surface on the upstream side of the combustion gas flow path and the gas path surface;
Perform gas turbine parts repair method.

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