Patent Application
20230374906
The present disclosure relates to a high-temperature part which is allowed to be exposed to a combustion gas and a gas turbine including the same.
Priority is claimed on Japanese Patent Application No. 2022-082124, filed on May 19, 2022, the content of which is incorporated herein by reference.
A gas turbine includes a compressor which compresses air to produce compressed air, a combustor which bus fuel in the compressed air to produce a combustion gas, a turbine which is driven by the combustion gas, and an intermediate casing. The compressor includes a compressor rotor which rotates about an axis and a compressor casing which covers the compressor rotor. The combustor includes a burner which injects fuel and a transition piece (or a combustion cylinder) which sends the combustion gas produced by the combustion of the fuel to the turbine. The turbine includes a turbine rotor which rotates about an axis, a turbine casing which covers the turbine rotor, and a plurality of stator vane rows. The turbine rotor includes a rotor shaft which rotates about an axis and a plurality of rotor blade rows which are attached to the rotor shaft. The plurality of rotor blade rows are arranged in an axial direction in which an axis extends. Each rotor blade row includes a plurality of rotor blades which are arranged in a circumferential direction around an axis. The plurality of stator vane rows are arranged in an axial direction and are attached to an inner peripheral side of the turbine casing. Each of the plurality of stator vane rows is disposed on an axial upstream side of any one rotor blade row of the plurality of rotor blade rows. Each stator vane row includes a plurality of stator vanes which are arranged in a circumferential direction around an axis. The turbine casing includes a split ring. The split ring is axially disposed between the plurality of stator vane rows and defines an outer peripheral side of a combustion gas flow path through which the combustion gas flows in the turbine.
The compressor casing and the turbine casing are connected through the intermediate casing. The combustor is attached to the intermediate casing. The transition piece of the combustor is disposed inside the intermediate casing. The compressed air from the compressor is discharged into the intermediate casing. The compressed air flows into the combustor and is used for the combustion of the fuel.
An outlet flange of the transition piece and a shroud of the first stage stator vane constituting the stator vane row on the most axial upstream side in the plurality of stator vane rows are connected by an outlet seal (or a combustion cylinder seal).
All of the transition piece, the outlet seal, the stator vane, the split ring, and the rotor blade in the above-described component parts of the gas turbine are the high-temperature parts exposed to the combustion gas.
Japanese Unexamined Patent Application No. 2021-131041 discloses an outlet seal which is a kind of high-temperature part. The outlet seal includes a body portion which defines a part of a combustion gas flow path, a transition piece connection portion to which an outlet flange of a transition piece is connected, and a stator vane connection portion to which a shroud of a first stage stator vane is connected. The transition piece connection portion is provided on the axial upstream side of the body portion. The stator vane connection portion is provided on the axial downstream side of the body portion. The outlet seal includes a base and a heat insulating coat covering a part of the surface of the base. The heat insulating coat is formed on a surface of a flow path forming portion which is a portion forming the body portion in the base.
The high-temperature part such as the outlet seal may come into contact with other adjacent parts during an assembly step of the gas turbine and the heat insulating coat of the high-temperature part may be damaged.
Here, an object of the present disclosure is to provide a high-temperature part capable of suppressing damage of a gas turbine in an assembly step.
A high-temperature part of an aspect according to the invention for achieving the above-described object is a high-temperature part of a gas turbine exposed to a combustion gas including: a high-temperature part body which includes a base and a coating layer formed on a part of a surface of the base; and a protective layer which is formed on at least a part of a surface of the coating layer. The protective layer is formed of a material which is allowed to disappear from the surface of the coating layer under an operation environment of the gas turbine.
In this aspect, since the coating layer is protected by the protective layer when the gas turbine is assembled by using the high-temperature part, it is possible to suppress damage of the coating layer. Further, in this aspect, when the combustion gas contacts the high-temperature pan by operation of the gas turbine, the protective layer in the high-temperature part disappears from the surface of the coating layer due to the heat of the combustion gas. Therefore, even when the high-temperature part of this aspect includes a protective layer, there is no influence on the performance of the gas turbine.
A gas turbine according to the invention for achieving the above-described object includes: a high-temperature part of the gas turbine according to the above-described aspect; and a plurality of other parts not exposed to a combustion gas. The plurality of other parts include all parts constituting a compressor of the gas turbine and parts forming an outer shape of a turbine of the gas turbine.
A method of operating a high-temperature pan of a gas turbine according to the invention for achieving the above-described object is a method of operating a high-temperature part of a gas turbine exposed to a combustion gas, the method including: a preparation step of preparing a high-temperature part including a high-temperature part body and a protective layer formed on at least a part of a surface of the high-temperature part body; an assembly step of assembling a gas turbine by using the high-temperature part and a plurality of other parts; and an operation step of producing a combustion gas by supplying fuel to the gas turbine. The high-temperature part body includes a base and a coating layer formed on a part of a surface of the base. The protective layer is formed on at least a part of a surface of the coating layer. The protective layer is formed of a material which is allowed to disappear from the surface of the coating layer under an operation environment of the gas turbine. The protective layer disappears from the surface of the coating layer due to the influence of heat of the combustion gas during the operation step.
In this aspect, it is possible to suppress damage of the high-temperature part body during the assembly step. Further, when the combustion gas contacts the high-temperature pan by performing the operation step, the protective layer in the high-temperature part can disappear.
According to one aspect of the present disclosure, it is possible to suppress of the high-temperature part when assembling the gas turbine using the high-temperature part.
Hereinafter, an embodiment according to the present disclosure will be described in detail with reference to the drawings.
An embodiment of a gas turbine will be described with reference to
As shown in
The compressor 20 includes a compressor rotor 21 which rotates about an axis Ar, a compressor casing 25 which covers the compressor rotor 21, and a plurality of stator vane rows 26. The turbine 40 includes a turbine rotor 41 which rotates about the axis Ar, a turbine casing 45 which covers the turbine rotor 41, and a plurality of stator vane rows 44. Additionally, hereinafter, the extension direction of the axis Ar is referred to as the axial direction Da, the circumferential direction about the axis Ar is simply referred to as the circumferential direction Dc. and the direction perpendicular to the axis Ar is referred to as the radial direction Dr. Further, one side of the axial direction Dra is referred to as the axial upstream side Dau and the opposite side is referred to as the axial downstream side Dad. Further, the side closer to the axis Ar in the radial direction Dr is referred to as the radial inside Dri and the opposite side is referred to as the radial outside Dro.
The compressor 20 is disposed on the axial upstream side Dau with respect to the turbine 40.
The compressor rotor 21 and the turbine rotor 41 are located on the same axis Ar and are connected to each other to form a gas turbine rotor 11. For example, a rotor of a generator GEN is connected to the gas turbine rotor 11. The gas turbine 10 further includes an intermediate casing 16. This intermediate casing 16 is disposed between the compressor casing 25 and the turbine casing 45 in the axial direction Da. The compressor casing 25, the intermediate casing 16, and the turbine casing 45 are connected to each other to form a gas turbine casing 15.
As shown in
The turbine rotor 41 includes a rotor shaft 42 which extends in the axial direction Da around the axis Ar and a plurality of rotor blade rows 43 which are attached to the rotor shaft 42. The plurality of rotor blade rows 43 are arranged in the axial direction Da. Each rotor blade row 43 includes a plurality of rotor blades 43a which are arranged in the circumferential direction Dc. Any one stator vane row 44 of the plurality of stator vane rows 44 is disposed on each axial upstream side Dau of the plurality of rotor blade rows 43. Each stator vane row 44 is provided inside the turbine casing 45. Each stator vane row 44 includes a plurality of stator vanes 44a which are arranged in the circumferential direction Dc.
An annular space which is located between the outer peripheral side of the rotor shaft 42 and the inner peripheral side of the turbine casing 45 and in which the stator vane 44a and the rotor blade 43a are arranged in the axial direction Da forms a combustion gas flow path 49p through which a combustion gas G flows from the combustor 30. This combustion gas flow path 49p has an annular shape centered on the axis Ar and is elongated in the axial direction Da.
The turbine casing 45 includes a plurality of split rings 46, a plurality of heat insulating rings 47, a Vane ring 48, and a turbine casing body 49. The split ring 46 is located on the radial outside Dro of the rotor blade row 43 and faces the rotor blade row 43 in the radial direction Dr. The Vane ring 48 has an annular shape centered on the axis Ar and is located on the radial outside Dro of the plurality of split rings 46 or the stator vane 44a. One heat insulating ring 47 of the plurality of heat insulating rings 47 is located between the split ring 46 and the Vane ring 48 in the radial direction Dr and connects the split ring 46 and the Vane ring 48. Further, the remaining heat insulating rings 47 of the plurality of heat insulating rings 47 are located between the stator vane 44a and the Vane ring 48 in the radial direction Dr and connect the stator vane 44a and the Vane ring 48. The Vane ring 48 is fixed to the inner peripheral side of the turbine casing body 49.
The plurality of combustors 30 are arranged in the circumferential direction Dc around the axis Ar and are attached to the intermediate casing 16. The combustor 30 includes a transition piece (or a combustion cylinder) 32 in which the fuel F is combusted and a plurality of burners 31 which inject fuel into the transition piece 32. The inner peripheral side of the transition piece 32 forms a combustion space (or a combustion gas flow path) 39p. The transition piece 32 extends in a direction including a directional component of the axial downstream side Dad while the combustor 30 is attached to the intermediate casing 16.
As shown in
The stator vane 50 includes a Vane body 51 which has an airfoil-shape cross-section and a shroud 52 which is provided on both sides of the Vane body 51 in the height direction of the vane. Additionally, the shroud 52 which is provided on one side of the Vane body 51 in the height direction of the vane is an inner shroud 52i and the shroud 52 which is provided on the other side of the Vane body 51 in the height direction of the vane is an outer shroud 52o. Both the inner shroud 52i and the outer shroud 52o spread in a direction perpendicular to the height direction of the vane. In a state in which the stator vane 50 is attached to the turbine casing 45, the height direction of the vane is the radial direction Dr. Further, one side of the height direction of the vane is the radial outside Dro and the other side of the height direction of the vane is the radial inside Dri. Thus, the inner shroud 52i is provided on the radial inside Dri of the Vane body 51 and the outer shroud 52o is provided on the radial outside Dro of the Vane body 51. The inner shroud 52i defines a part of an edge on the radial inside Dri of the combustion gas flow path 49p. The outer shroud 52o defines an edge on the radial outside Dro of the combustion gas flow path 49p. The Vane body 51 which is located between the inner shroud 52i and the outer shroud 52o in the radial direction Dr is located in the combustion gas flow path 49p through which the combustion gas G passes.
As shown in
As shown in
As shown in
As shown in
The split ring 60 also includes a base 65 and a coating layer 67 which is formed on a part of the surface of the base 65 similarly to the stator vane 50 or the transition piece 32. A portion forming the split ring body 61 in the base 65 forms the flow path forming portion 66.
As shown in
The outlet seal 70 also includes a base 75 and a coating layer 77 which is formed on a part of the surface of the base 75 similarly to the stator vane 50 and the like. A portion forming the body portion 71 in the base 75 forms a flow path forming portion 76.
The flow path forming portion 76 includes a flow path defining surface 76m which defines a part of the combustion gas flow path 79p, a downstream end surface 76d which is connected to an edge on the axial downstream side Dad of the flow path defining surface 76m, and an upstream end surface 76u which is connected to an edge on the axial upstream side Dau of the flow path defining surface 76m.
The coating layer 77 includes a main coating portion 77m which is formed on the flow path defining surface 76m, a downstream coating portion 77d which is formed on the downstream end surface 76d to be connected to the main coating portion 77m, and an upstream coating portion 77u which is formed on the upstream end surface 76u to be connected to the main coating portion 77m.
An embodiment of the high-temperature part will be described with reference to
As shown in
The protective layer 85 includes an end surface protection portion 85e and a main protection portion 85m. The end surface protection portion 85e is formed on the surface of the downstream coating portion 774 in the outlet seal 70 which is the high-temperature part body 81. The main protection portion 85m is formed on the surface of the main coating portion 77m to be connected to the end surface protection portion 85e.
The protective layer 85 has a property that the protective layer adheres to the surface of the high-temperature part 80 and disappears due to the temperature by heat or the combustion by heat during a normal operation of the gas turbine 10. Further, the protective layer 85 preferably has a certain level of elasticity. Examples of such a protective layer 85 include the following protective layer forming materials.
All of the protective layer forming materials provided as exemplary examples above are resins or materials containing resins as main components. In the protective layer forming materials containing resins as main components, for example, the residual component may be metal powder or the like.
Since the coating layer 77 formed on the surface of the outlet seal 70 which is the high-temperature part body 81 is very hard, the coating is vulnerable to impact. Therefore, for example, if the outlet seal 70 collides with the stator vane 50 when the outlet seal 70 is assembled to the stator vane 50, the coating layer 77 may be damaged.
When the outlet seal 70 is assembled to the stator vane 50, there is a high possibility that the main coating portion 77m and the downstream coating portion 77d of the outlet seal 70 and the periphery of the corner collide with the shroud 52 of the stator vane 50. Here, in this embodiment, in order to protect the main coating portion 77m and the downstream coating portion 77d of the outlet seal 70 and the periphery of the corner, the protective layer 85 is formed in the periphery of the corner.
Next, a method of operating the high-temperature part 80 will be described according to the flowchart shown in
First, the above-described high-temperature part 80 is prepared (preparation step S1). Next, the gas turbine 10 is assembled by using the high-temperature part 80 and a plurality of other parts not exposed to the combustion gas G (assembly step S2). Additionally, the plurality of other parts not exposed to the combustion gas G are all parts constituting the compressor 20, the turbine casing body 49 corresponding to a part forming the outer shape of the turbine 40 among the parts constituting the turbine 40, and the like.
As described above, since the coating layer 77 in the high-temperature part 80 is protected by the protective layer 85, it is possible to suppress damage of the coating layer 77 in the high-temperature part 80 in the assembly step S2.
Next, the fuel F is supplied to the gas turbine 10 to produce the combustion gas G (operation step S3).
When the combustion gas 0 is once caused to contact the high-temperature part 80 by performing the operation step S3, the protective layer 85 in the high-temperature part 80 disappears from the surface of the coating layer 77 due to the heat of the combustion gas G.
Thus, even when the high-temperature part 80 includes the protective layer 85, there is no influence on the performance of the gas turbine 10.
The gas turbine 10 includes the high-temperature part. 80 at, the ending time point of the assembly step S2. However, since the protective layer 85 in the high-temperature part 80 disappears when the operation step S3 is first performed, the gas turbine 10 includes the high-temperature part body 81, but does not include the high-temperature part 80.
As described above, the protective layer 85 is formed in the periphery of the corner between the downstream coating portion 77d and the main coating portion 77m of the outlet seal 70 which is the high-temperature part body 81. However, the protective layer 85 may be formed in the periphery of the corner between the upstream coating portion 77u and the main coating portion 77m of the outlet seal 70 and the protective layer 85 may be formed in the periphery of both corners.
The high-temperature part 80 of the above-described embodiment is a pan including the outlet seal 70 as the high-temperature part body 81. However, the high-temperature part may use the transition piece 32, the stator vane 50, and the split ring 60 described above as the high-temperature part bodies. These high-temperature part bodies also include the bases 35, 55, and 65 and the mating layers 37, 57, and 67 similarly to the above-described outlet seal 70.
As described above, the base 35 of the transition piece 32, the base 55 of the stator vane 50, and the base 65 of the split ring 60 also include the flow path forming portions 36, 56, and 66 defining a part of the combustion gas flow paths 39p and 49p through which the combustion gas flows. These flow path funning portions 36, 56, and 66 also include the flow path defining surface, the downstream end surface, and the upstream end surface defining a part of the combustion gas flow paths 39p and 49p similarly to the flow path funning portion 76 of the outlet seal 70.
The coating layers 37, 57, and 67 are formed on a part of the surfaces of these flow path forming portions 36, 56, and 66. The coating layers 37, 57, and 67 include the main coating portion which is formed on the flow path defining surface, the downstream coating portion formed on the downstream end surface to be connected to the main coating portion, and the upstream coating portion formed on the upstream end surface to be connected to the main coating portion.
The high-temperature part can be made by forming the protective layer on the surface of the above-described high-temperature part body. The protective layers of these high-temperature parts also include the end surface protection portion and the main protection portion similarly to the above-described outlet seal 70. The end surface protection portion is formed any one of the downstream coating portion and the upstream coating portion. The main protection portion is formed on the main protection portion 85m to be connected to the end surface protection portion.
Even in the above-described high-temperature part, since the coating layers 37, 57, and 67 of the high-temperature part are protected by the protective layer, it is possible to suppress damage of the coating layers 37, 57, and 67 of the high-temperature pan during the assembly step S2.
The high-temperature part 80 of the gas turbine 10 of the above-described embodiment is understood, for example, a below.
In this aspect, since the coating layer 77 is protected by the protective layer 85 when the gas turbine 10 is assembled by using the high-temperature part 80, damage of the coating layer 77 can be suppressed. Further, in this aspect, when the combustion gas G is once caused to contact the high-temperature part 80 by operation of the gas turbine 10, the protective layer 85 in the high-temperature part 80 disappears from the surface of the coating layer 77 due to the heat of the combustion gas G. Therefore, even when the high-temperature part 80 of this aspect includes the protective layer 85, there is no influence on the performance of the gas turbine 10.
In this aspect, since the material forming the protective layer 85 is a resin or a material containing a resin as a main component, the protective layer 85 can have a certain level of elasticity. Therefore, in this aspect, it is possible to further suppress damage of the coating layer 77 when the gas turbine 10 is assembled by using the high-temperature part 80. Further, in this aspect, the protective layer 85 can be combusted or sublimated at a relatively low temperature.
In this aspect, when the gas turbine 10 is assembled by using the high-temperature part 80, it is possible to suppress damage in the periphery of the corner between the main coating portion 77m and the downstream coating portion 77d or the corner between the main coating portion 77m and the upstream coating portion 77u which is most likely to be impacted in the high-temperature part 80.
The gas turbine 10 of the above-described embodiment is understood, for example, as below.
The method of operating the high-temperature part 80 of the gas turbine 10 of the above-described embodiment is understood, for example, as below.
In this aspect, it is possible to suppress damage of the high-temperature part body 81 during the assembly step S2. Further, when the combustion gas G is once caused to contact the high-temperature part 80 by performing the operation step S3, the protective layer 85 in the high-temperature part 80 can disappear.
| Number | Date | Country | Kind |
|---|---|---|---|
| 2022-082124 | May 2022 | JP | national |
