CN111254377A - Repair method for long-life thermal barrier coating of F-grade ground heavy gas turbine blade - Google Patents
Repair method for long-life thermal barrier coating of F-grade ground heavy gas turbine blade Download PDFInfo
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- CN111254377A CN111254377A CN202010075337.4A CN202010075337A CN111254377A CN 111254377 A CN111254377 A CN 111254377A CN 202010075337 A CN202010075337 A CN 202010075337A CN 111254377 A CN111254377 A CN 111254377A
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- C21D—MODIFYING THE PHYSICAL STRUCTURE OF FERROUS METALS; GENERAL DEVICES FOR HEAT TREATMENT OF FERROUS OR NON-FERROUS METALS OR ALLOYS; MAKING METAL MALLEABLE, e.g. BY DECARBURISATION OR TEMPERING
- C21D1/00—General methods or devices for heat treatment, e.g. annealing, hardening, quenching or tempering
- C21D1/74—Methods of treatment in inert gas, controlled atmosphere, vacuum or pulverulent material
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- C23C4/00—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge
- C23C4/04—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the coating material
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- C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
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- C23C4/00—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge
- C23C4/12—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the method of spraying
- C23C4/134—Plasma spraying
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- C—CHEMISTRY; METALLURGY
- C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
- C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
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Abstract
The invention provides a method for repairing a long-life thermal barrier coating of a F-grade ground heavy gas turbine blade, which solves the technical problem that the long-life thermal barrier coating of the turbine blade in the prior art cannot be repaired. It comprises the following steps: (1) stripping an old thermal barrier coating surface layer on the turbine blade; (2) removing the old thermal barrier coating bonding layer on the turbine blade; (3) preparing a new thermal barrier coating bonding layer by adopting supersonic flame spraying; (4) preparing a new thermal barrier coating surface layer by adopting plasma spraying; (5) vacuum pretreatment: (6) aging treatment: (7) and (4) processing the surface of the thermal barrier coating and a cooling hole. The invention designs a special heat treatment system, ensures the performance of the thermal barrier coating while not losing the service life of the blade substrate, and can improve the 870 ℃/370MPa endurance of the blade substrate to 105-110 percent of the original value.
Description
Technical Field
The invention relates to the field of gas turbine maintenance, in particular to a method for repairing a long-life thermal barrier coating of a F-grade ground heavy gas turbine blade.
Background
The inlet temperature of the turbine of a certain F-grade ground heavy gas turbine is up to 1400 ℃, and the operation time is up to tens of thousands of hours. After a maintenance period (16000EOH), thermal barrier coatings on the surfaces of turbine blades (turbine movable blades and turbine fixed blades) are damaged to different degrees (including chipping, burning loss, cracking, degradation and the like), and the operation performance and safety of a unit are seriously affected. In order to meet the service performance requirement of a unit and provide tens of thousands of hours of safe operation guarantee, the damaged thermal barrier coating of the turbine blade needs to be repaired.
At present, the research work of the long-service-life thermal barrier coating of the gas turbine engine in China mainly focuses on the development and test stage of a test piece, and a long-service-life coating repairing case applied to an actual turbine blade is not found.
The applicant has found that the prior art has at least the following technical problems:
1. long-life thermal barrier coatings of prior art turbine blades cannot be repaired.
Disclosure of Invention
The invention aims to provide a method for repairing a long-life thermal barrier coating of a F-grade ground heavy-duty gas turbine blade, which aims to solve the technical problem that the long-life thermal barrier coating of the turbine blade in the prior art cannot be repaired. The technical effects that can be produced by the preferred technical scheme in the technical schemes provided by the invention are described in detail in the following.
In order to achieve the purpose, the invention provides the following technical scheme:
the invention provides a method for repairing a long-life thermal barrier coating of a F-grade ground heavy gas turbine blade, which comprises the following steps:
(1) stripping an old thermal barrier coating surface layer on the turbine blade;
(2) removing the old thermal barrier coating bonding layer on the turbine blade;
(3) preparing a new thermal barrier coating bonding layer by adopting supersonic flame spraying;
(4) preparing a new thermal barrier coating surface layer by adopting plasma spraying;
(5) vacuum pretreatment:
putting the turbine blade sprayed with the new thermal barrier coating surface layer in the step (4) into a vacuum furnace, and keeping the vacuum degree to be less than or equal to 1.0 multiplied by 10-3Pa, keeping the temperature at 1070-1090 ℃ for 2-2.5 h, cooling the temperature to 900 ℃ along with the furnace after the heat preservation is finished, and then cooling the temperature to room temperature;
(6) aging treatment:
putting the turbine blade subjected to vacuum pretreatment in the step (5) into an air furnace, preserving the heat for 16-18 h at the temperature of 840-860 ℃, and air-cooling to room temperature after the heat preservation is finished;
(7) processing the surface of the thermal barrier coating and the cooling hole:
and (4) carrying out surface roughness reduction treatment on the turbine blade subjected to aging treatment in the step (6), and dredging cooling holes in the turbine blade.
Further, the method also comprises a step (8) of detecting the thickness of the coating after the step (7), and specifically comprises the following steps: and detecting whether the thickness of the thermal barrier coating on the turbine blade is 0.5mm-0.6mm by using an eddy current thickness gauge.
Further, after the step (8), the method also comprises a step (9) of water flow-through inspection, specifically: the cooling holes were checked for 100% open flow.
Further, in the step (1), peeling off the old thermal barrier coating surface layer on the turbine blade, specifically: stripping the surface layer of the old thermal barrier coating by adopting a sand blowing method, wherein the white corundum particles are 60 meshes, the sand blowing pressure is 0.3MPa-0.5MPa, and the distance is 140mm-160 mm.
Further, in the step (2), removing the old thermal barrier coating bonding layer on the turbine blade specifically comprises: removing the old thermal barrier coating bonding layer by adopting a hydrochloric acid solution, and immersing the turbine blade stripped of the old thermal barrier coating surface layer in the hydrochloric acid solution with the temperature of 18-22 ℃ for 42-48 h; the turbine blades are then rinsed clean.
Further, in the step (3), a new thermal barrier coating bonding layer is prepared by supersonic flame spraying, which specifically comprises the following steps: the bonding layer material is selected from CoNiCrAlY multi-element alloy powder, and the thickness of the bonding layer is 0.10mm-0.15 mm.
Further, in the step (4), a new thermal barrier coating surface layer is prepared by adopting plasma spraying, and specifically, 8% Y is adopted as a surface layer material2O3Stabilized ZrO2Powder, the thickness of the surface layer is 0.40mm-0.45 mm.
Further, in the step (7), the thermal barrier coating surface and the cooling hole are processed, specifically: the surface of the surface layer of the thermal barrier coating is polished by adopting flexible diamond polishing cloth; the cooling hole is dredged by a diamond grinding head with the phi of 1.0 mm.
Further, the vacuum pretreatment in the step (5) is to place the turbine blade sprayed with the new thermal barrier coating surface layer in the step (4) into a vacuum furnace, and the vacuum degree is 8.0 × 10-4Pa, keeping the temperature at 1080 ℃ for 2h, cooling to 900 ℃ along with the furnace after the heat preservation is finished, and then cooling to room temperature by air.
Further, the aging treatment in the step (6) is that the turbine blade after vacuum pretreatment in the step (5) is put into an air furnace, the temperature is kept for 16h at 850 ℃, and the air cooling is carried out to the room temperature after the temperature keeping is finished.
Based on the technical scheme, the embodiment of the invention can at least produce the following technical effects:
(1) the method for repairing the long-life thermal barrier coating of the F-grade ground heavy gas turbine blade provided by the invention designs a special heat treatment system, so that the performance of the thermal barrier coating is ensured while the service life of a blade substrate is not lost. Firstly, forming a compact thermal oxide growth layer (TGO) between the bonding layer and the surface layer through vacuum pretreatment; and the long-time oxidation resistance of the coating can be enhanced through aging treatment, and the 870 ℃/370MPa durability of the blade substrate can be improved to 105-110% of the original value.
(2) According to the method for repairing the long-life thermal barrier coating of the F-level ground heavy gas turbine blade, the diamond polishing cloth is adopted to process the thermal barrier coating, so that the surface quality of the thermal barrier coating can be improved, the separation of boundary layers is reduced, the microstructure of the thermal barrier coating is not damaged, and the problems that the coating is high in hardness, difficult to process and easy to crack are solved.
(3) The method for repairing the long-life thermal barrier coating of the F-grade ground heavy gas turbine blade combines the vacuum pretreatment and the blade stabilization treatment process (the heat treatment process for keeping the microstructure of the blade stable), and reduces the process period and the cost.
(4) The method for repairing the long-life thermal barrier coating of the F-grade ground heavy gas turbine blade separates an aging treatment process from a vacuum pretreatment process, and can reduce the production cost by replacing a vacuum furnace with an air furnace.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the drawings without creative efforts.
FIG. 1 is a metallographic image of the thermal barrier coating obtained in example 1 of the present invention, taken by means of an optical microscope, on a scale of 100 um;
FIG. 2 is a microstructure of a turbine blade base material before vacuum pretreatment and aging treatment in example 1 of the present invention, which is photographed with a scanning electron microscope and has a scale of 1 um;
FIG. 3 is a microstructure of a turbine blade base material before vacuum pretreatment and aging treatment in example 1 of the present invention, which is photographed by an optical microscope with a 500um scale;
FIG. 4 is a microstructure of a turbine blade base material subjected to vacuum pretreatment and aging treatment in example 1 of the present invention, which was photographed with a scanning electron microscope and measured with a scale of 1 um;
FIG. 5 is a microstructure of a turbine blade base material subjected to vacuum pretreatment and aging treatment in example 1 of the present invention, which was photographed with an optical microscope and measured at 500. mu.m.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, the technical solutions of the present invention will be described in detail below. It is to be understood that the described embodiments are merely exemplary of the invention, and not restrictive of the full scope of the invention. All other embodiments, which can be derived by a person skilled in the art from the examples given herein without any inventive step, are within the scope of the present invention.
Description of raw materials:
CoNiCrAlY multi-component alloy powder (Metco4700) produced by SulShoume was used.
Example 1:
1.1 repair of thermal barrier coating:
after a certain F-grade heavy-duty gas turbine stationary blade is used for 4.8 ten thousand Equivalent Operating Hours (EOH), coating chipping, coating cracking and coating burning loss occur on the blade back, the edge plate and the air inlet edge of the blade. The thermal barrier coating of the turbine stationary blade is repaired, and the method comprises the following specific steps:
(1) stripping the old thermal barrier coating surface layer on the turbine blade: stripping the surface layer of the old thermal barrier coating by adopting a sand blowing method, wherein white corundum particles are 60 meshes, the sand blowing pressure is 0.4MPa, and the distance is 150 mm;
(2) removing an old thermal barrier coating bonding layer on the turbine blade:
removing the old thermal barrier coating bonding layer by adopting a hydrochloric acid solution, and soaking the turbine blade stripped of the old thermal barrier coating surface layer in a hydrochloric acid solution with the temperature of 20 ℃ and the concentration of 30% for 45 hours; then, the turbine blade is washed clean by deionization;
(3) preparing a new thermal barrier coating bonding layer by adopting supersonic flame spraying:
the bonding layer material is selected from CoNiCrAlY multi-element alloy powder, and the thickness of the bonding layer is 0.13 mm;
(4) preparing a new thermal barrier coating surface layer by adopting plasma spraying:
the new thermal barrier coating surface layer is prepared by adopting plasma spraying, and specifically, 8 percent Y is adopted as the surface layer material2O3Stabilized ZrO2Powder, the thickness of the surface layer is 0.43 mm;
(5) vacuum pretreatment:
putting the turbine blade sprayed with the new thermal barrier coating surface layer in the step (4) into a vacuum furnace, and controlling the vacuum degree to be 8.0 multiplied by 10-4Pa, keeping the temperature at 1080 ℃ for 2 hours, cooling the temperature to 900 ℃ along with the furnace after the heat preservation is finished, and then cooling the temperature to room temperature;
(6) aging treatment:
putting the turbine blade subjected to vacuum pretreatment in the step (5) into an air furnace, preserving heat for 16h at the temperature of 850 ℃, and air-cooling to room temperature after heat preservation is finished;
(7) processing the surface of the thermal barrier coating and the cooling hole:
grinding and polishing the surface of the thermal barrier coating surface layer of the turbine blade subjected to the aging treatment in the step (6) by adopting flexible diamond polishing cloth; and a diamond grinding head with the phi of 1.0mm is adopted to dredge the cooling hole on the turbine blade;
(8) and (3) coating thickness detection:
and detecting the thickness of the thermal barrier coating on the turbine blade to be 0.56mm by adopting an eddy current thickness gauge.
(9) And (3) water flow inspection:
the cooling hole was dredged to 100%.
1.2 testing the prepared thermal barrier coating:
① the thickness of the thermal barrier coating of the turbine blade prepared by the method is 0.56mm, and in the range of 0.50mm-0.60mm, the cooling hole is not blocked, and meets the fault inspection standard, and can be installed for use.
② the thermal barrier coating prepared by the above method is examined metallographically, the metallograph is shown in figure 1, the upper structure is a surface layer, and the lower structure is a bonding layer.
③ after vacuum pretreatment and aging treatment, the durability of 870 ℃/370MPa of the turbine blade base material reaches 108% of the original value, and the specific results are shown in Table 1 below.
④ see FIGS. 2-5 for a view of the microstructure of the turbine blade base material before and after vacuum pretreatment plus aging treatment.
TABLE 1 high temperature durability of the turbine blade base material before and after vacuum pretreatment and aging treatment (test conditions: 870 deg.C, 370MPa)
1.3 comparative experiment:
after the turbine blade is installed on site and used for a maintenance period (16000EOH), the thermal barrier coating of the turbine blade is good, and faults such as layer rising, block falling and the like do not exist. The results show that the thermal barrier coating prepared in example 1 of the present invention performs better than the coating performance of the original manufacturer's new product.
Example 2:
2.1 repair of thermal barrier coating:
after a certain F-grade heavy-duty gas turbine stationary blade is used for 4.8 ten thousand Equivalent Operating Hours (EOH), coating chipping, coating cracking and coating burning loss occur on the blade back, the edge plate and the air inlet edge of the blade. The thermal barrier coating of the turbine stationary blade is repaired, and the method comprises the following specific steps:
(1) stripping the old thermal barrier coating surface layer on the turbine blade: stripping the surface layer of the old thermal barrier coating by adopting a sand blowing method, wherein white corundum particles are 60 meshes, the sand blowing pressure is 0.3MPa, and the distance is 140 mm;
(2) removing an old thermal barrier coating bonding layer on the turbine blade:
removing the old thermal barrier coating bonding layer by adopting a hydrochloric acid solution, and soaking the turbine blade stripped of the old thermal barrier coating surface layer in a hydrochloric acid solution with the temperature of 22 ℃ and the concentration of 30% for 42 hours; then, the turbine blade is washed clean by deionization;
(3) preparing a new thermal barrier coating bonding layer by adopting supersonic flame spraying:
the bonding layer material is selected from CoNiCrAlY multi-element alloy powder, and the thickness of the bonding layer is 0.15 mm;
(4) preparing a new thermal barrier coating surface layer by adopting plasma spraying:
the new thermal barrier coating surface layer is prepared by adopting plasma spraying, and specifically, 8 percent Y is adopted as the surface layer material2O3Stabilized ZrO2Powder, the thickness of the surface layer is 0.45 mm;
(5) vacuum pretreatment:
putting the turbine blade sprayed with the new thermal barrier coating surface layer in the step (4) into a vacuum furnace, and controlling the vacuum degree to be 1.0 multiplied by 10-3Pa, preserving heat for 2.5 hours at the temperature of 1090 ℃, cooling to 900 ℃ along with the furnace after heat preservation, and then cooling to room temperature;
(6) aging treatment:
putting the turbine blade subjected to vacuum pretreatment in the step (5) into an air furnace, preserving heat for 18 hours at the temperature of 840 ℃, and air-cooling to room temperature after heat preservation;
(7) processing the surface of the thermal barrier coating and the cooling hole:
grinding and polishing the surface of the thermal barrier coating surface layer of the turbine blade subjected to the aging treatment in the step (6) by adopting flexible diamond polishing cloth; and a diamond grinding head with the phi of 1.0mm is adopted to dredge the cooling hole on the turbine blade;
(8) and (3) coating thickness detection:
and detecting the thickness of the thermal barrier coating on the turbine blade to be 0.60mm by adopting an eddy current thickness gauge.
(9) And (3) water flow inspection:
the cooling hole was dredged to 100%.
2.2 testing the thermal barrier coating prepared again:
① the thickness of the thermal barrier coating of the turbine blade prepared by the method is 0.60mm, and in the range of 0.50mm-0.60mm, the cooling hole is not blocked, and meets the fault inspection standard, and can be installed for use.
② after vacuum pretreatment and aging treatment, the durability of 870 ℃/370MPa of the turbine blade base material reaches 110% of the original value, and the specific results are shown in Table 2 below.
TABLE 2 high temperature durability of the turbine blade base material before and after vacuum pretreatment and aging treatment (test conditions: 870 deg.C, 370MPa)
2.3 comparative experiment:
after the turbine blade is installed on site and used for a maintenance period (16000EOH), the thermal barrier coating of the turbine blade is good, and faults such as layer rising, block falling and the like do not exist. The results show that the thermal barrier coating prepared in example 2 of the present invention has better performance than the coating of the original manufacturer.
Example 3:
3.1 repair of thermal barrier coating:
after a certain F-grade heavy-duty gas turbine stationary blade is used for 4.8 ten thousand Equivalent Operating Hours (EOH), coating chipping, coating cracking and coating burning loss occur on the blade back, the edge plate and the air inlet edge of the blade. The thermal barrier coating of the turbine stationary blade is repaired, and the method comprises the following specific steps:
(1) stripping the old thermal barrier coating surface layer on the turbine blade: stripping the surface layer of the old thermal barrier coating by adopting a sand blowing method, wherein white corundum particles are 60 meshes, the sand blowing pressure is 0.5MPa, and the distance is 160 mm;
(2) removing an old thermal barrier coating bonding layer on the turbine blade:
removing the old thermal barrier coating bonding layer by adopting a hydrochloric acid solution, and soaking the turbine blade stripped of the old thermal barrier coating surface layer in a hydrochloric acid solution with the temperature of 18 ℃ and the concentration of 30% for 48 hours; then, the turbine blade is washed clean by deionization;
(3) preparing a new thermal barrier coating bonding layer by adopting supersonic flame spraying:
the bonding layer material is selected from CoNiCrAlY multi-element alloy powder, and the thickness of the bonding layer is 0.10 mm;
(4) preparing a new thermal barrier coating surface layer by adopting plasma spraying:
the new thermal barrier coating surface layer is prepared by adopting plasma spraying, and specifically, 8 percent Y is adopted as the surface layer material2O3Stabilized ZrO2Powder, the thickness of the surface layer is 0.40 mm;
(5) vacuum pretreatment:
putting the turbine blade sprayed with the new thermal barrier coating surface layer in the step (4) into a vacuum furnace, and controlling the vacuum degree to be 9.0 multiplied by 10-4Pa, keeping the temperature at 1070 ℃ for 2h, cooling the temperature to 900 ℃ along with the furnace after the heat preservation is finished, and then cooling the temperature to room temperature;
(6) aging treatment:
putting the turbine blade subjected to vacuum pretreatment in the step (5) into an air furnace, preserving heat for 18 hours at the temperature of 860 ℃, and air-cooling to room temperature after heat preservation is finished;
(7) processing the surface of the thermal barrier coating and the cooling hole:
grinding and polishing the surface of the thermal barrier coating surface layer of the turbine blade subjected to the aging treatment in the step (6) by adopting flexible diamond polishing cloth; and a diamond grinding head with the phi of 1.0mm is adopted to dredge the cooling hole on the turbine blade;
(8) and (3) coating thickness detection:
and detecting the thickness of the thermal barrier coating on the turbine blade to be 0.50mm by adopting an eddy current thickness gauge.
(9) And (3) water flow inspection:
the cooling hole was dredged to 100%.
3.2 testing the thermal barrier coating prepared again:
① the thickness of the thermal barrier coating of the turbine blade prepared by the method is 0.50mm, and in the range of 0.50mm-0.60mm, the cooling hole is not blocked, and meets the fault inspection standard, and can be installed for use.
② after vacuum pretreatment and aging treatment, the durability of 870 ℃/370MPa of the turbine blade base material reaches 105% of the original value, and the specific results are shown in Table 3 below.
TABLE 3 high temperature durability of the turbine blade base material before and after vacuum pretreatment and aging treatment (test conditions: 870 deg.C, 370MPa)
3.3 comparative experiment:
after the turbine blade is installed on site and used for a maintenance period (16000EOH), the thermal barrier coating of the turbine blade is good, and faults such as layer rising, block falling and the like do not exist. The results show that the thermal barrier coating prepared in example 3 of the present invention performs better than the coating performance of the original manufacturer's new product.
The above description is only for the specific embodiments of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art can easily conceive of the changes or substitutions within the technical scope of the present invention, and all the changes or substitutions should be covered within the scope of the present invention.
Claims (10)
1. A method for repairing a long-life thermal barrier coating of a F-grade ground heavy gas turbine blade is characterized by comprising the following steps: the method comprises the following steps:
(1) stripping an old thermal barrier coating surface layer on the turbine blade;
(2) removing the old thermal barrier coating bonding layer on the turbine blade;
(3) preparing a new thermal barrier coating bonding layer by adopting supersonic flame spraying;
(4) preparing a new thermal barrier coating surface layer by adopting plasma spraying;
(5) vacuum pretreatment:
putting the turbine blade sprayed with the new thermal barrier coating surface layer in the step (4) into a vacuum furnace,
under the vacuum degree of less than or equal to 1.0 multiplied by 10-3Pa, keeping the temperature at 1070-1090 ℃ for 2-2.5 h, cooling the temperature to 900 ℃ along with the furnace after the heat preservation is finished, and then cooling the temperature to room temperature;
(6) aging treatment:
putting the turbine blade subjected to vacuum pretreatment in the step (5) into an air furnace, preserving the heat for 16-18 h at the temperature of 840-860 ℃, and air-cooling to room temperature after the heat preservation is finished;
(7) processing the surface of the thermal barrier coating and the cooling hole:
and (4) carrying out surface roughness reduction treatment on the turbine blade subjected to aging treatment in the step (6), and dredging cooling holes in the turbine blade.
2. The method of claim 1 for repairing a long life thermal barrier coating on a class F ground heavy gas turbine blade, comprising: and (3) after the step (7), the step (8) of detecting the thickness of the coating is further included, and the method specifically comprises the following steps: and detecting whether the thickness of the thermal barrier coating on the turbine blade is 0.5mm-0.6mm by using an eddy current thickness gauge.
3. The method of claim 2 for repairing a long life thermal barrier coating on a class F ground heavy gas turbine blade, wherein: and (3) water circulation inspection after the step (8), specifically comprising the following steps: the cooling holes were checked for 100% open flow.
4. The method of claim 3 for repairing a long life thermal barrier coating on a class F earth heavy gas turbine blade, wherein: in the step (1), stripping an old thermal barrier coating surface layer on the turbine blade specifically comprises the following steps: stripping the surface layer of the old thermal barrier coating by adopting a sand blowing method, wherein the white corundum particles are 60 meshes, the sand blowing pressure is 0.3MPa-0.5MPa, and the distance is 140mm-160 mm.
5. The method of claim 4 for repairing a long life thermal barrier coating on a class F earth heavy gas turbine blade, wherein: in the step (2), removing the old thermal barrier coating bonding layer on the turbine blade specifically comprises the following steps: removing the old thermal barrier coating bonding layer by adopting a hydrochloric acid solution, and immersing the turbine blade stripped of the old thermal barrier coating surface layer in the hydrochloric acid solution with the temperature of 18-22 ℃ for 42-48 h; the turbine blades are then rinsed clean.
6. The method of claim 5 for repairing a long life thermal barrier coating on a class F earth heavy gas turbine blade, wherein: in the step (3), a new thermal barrier coating bonding layer is prepared by supersonic flame spraying, which specifically comprises the following steps: the bonding layer material is selected from CoNiCrAlY multi-element alloy powder, and the thickness of the bonding layer is 0.10mm-0.15 mm.
7. The method of claim 6 for repairing a long life thermal barrier coating on a class F ground heavy gas turbine blade, wherein: in the step (4), a new thermal barrier coating surface layer is prepared by adopting plasma spraying, and specifically, 8% Y is adopted as a surface layer material2O3Stabilized ZrO2Powder, the thickness of the surface layer is 0.40mm-0.45 mm.
8. The method of claim 7 for repairing a long life thermal barrier coating on a class F ground heavy gas turbine blade, wherein: in the step (7), the thermal barrier coating surface and the cooling hole are processed, specifically: the surface of the surface layer of the thermal barrier coating is polished by adopting flexible diamond polishing cloth; the cooling hole is dredged by a diamond grinding head with the phi of 1.0 mm.
9. The method for repairing a long-life thermal barrier coating of a class F earth-based heavy gas turbine blade according to any one of claims 1-8, wherein: the vacuum pretreatment in the step (5) is to place the turbine blade sprayed with the new thermal barrier coating surface layer in the step (4) into a vacuum furnace, and the vacuum degree is 8.0 multiplied by 10-4Pa, keeping the temperature at 1080 ℃ for 2h, cooling to 900 ℃ along with the furnace after the heat preservation is finished, and then cooling to room temperature by air.
10. The method of claim 9 for repairing a long life thermal barrier coating on a class F ground heavy gas turbine blade, comprising: and (4) the aging treatment in the step (6) is to place the turbine blade subjected to the vacuum pretreatment in the step (5) into an air furnace, keep the temperature for 16 hours at 850 ℃, and cool the turbine blade to room temperature after the heat preservation is finished.
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Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113106532A (en) * | 2021-04-07 | 2021-07-13 | 江苏源清动力技术有限公司 | Process for removing thermal barrier coating of thermal component of aero-engine and gas turbine |
| CN113957378A (en) * | 2021-10-22 | 2022-01-21 | 中国航发贵州黎阳航空动力有限公司 | Coating repair method for turbine blade |
| CN113969405A (en) * | 2021-10-27 | 2022-01-25 | 中国航发贵州黎阳航空动力有限公司 | Method for removing NiCrAlY coating of aeroengine blade |
| CN113981457A (en) * | 2021-10-27 | 2022-01-28 | 中国航发贵州黎阳航空动力有限公司 | Method for removing failure coating of single crystal high-pressure turbine blade |
| CN114990468A (en) * | 2022-06-08 | 2022-09-02 | 哈尔滨汽轮机厂有限责任公司 | Preparation method of gas turbine combustor transition section thermal barrier coating for protecting gas film divergence hole |
| CN115961248A (en) * | 2022-12-07 | 2023-04-14 | 中国航发动力股份有限公司 | Single crystal blade thermal barrier coating repairing method |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2008240563A (en) * | 2007-03-26 | 2008-10-09 | Toshiba Corp | Gas turbine high-temperature component repairing method, and gas turbine high-temperature component |
| CN106637031A (en) * | 2015-10-28 | 2017-05-10 | 三菱日立电力系统株式会社 | Thermal barrier coating applying method, thermal barrier coating repairing method, gas turbine member manufacture method and shielding pin |
| CN106637030A (en) * | 2015-10-28 | 2017-05-10 | 三菱日立电力系统株式会社 | Thermal barrier coating stripping method and thermal barrier coating repair method |
| US20180128166A1 (en) * | 2015-11-20 | 2018-05-10 | Federal-Mogul Llc | Steel piston crown and/or combustion engine components with dynamic thermal insulation coating and method of making and using such a coating |
| CN108118278A (en) * | 2016-11-29 | 2018-06-05 | 沈阳黎明航空发动机(集团)有限责任公司 | One kind is used for the low guide vane method for preparing heat barrier coating of IC10 alloys |
| CN110629152A (en) * | 2019-10-09 | 2019-12-31 | 中国人民解放军第五七一九工厂 | Spraying method of thermal barrier coating with longitudinal cracks for combustion chamber |
-
2020
- 2020-01-22 CN CN202010075337.4A patent/CN111254377A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2008240563A (en) * | 2007-03-26 | 2008-10-09 | Toshiba Corp | Gas turbine high-temperature component repairing method, and gas turbine high-temperature component |
| CN106637031A (en) * | 2015-10-28 | 2017-05-10 | 三菱日立电力系统株式会社 | Thermal barrier coating applying method, thermal barrier coating repairing method, gas turbine member manufacture method and shielding pin |
| CN106637030A (en) * | 2015-10-28 | 2017-05-10 | 三菱日立电力系统株式会社 | Thermal barrier coating stripping method and thermal barrier coating repair method |
| US20180128166A1 (en) * | 2015-11-20 | 2018-05-10 | Federal-Mogul Llc | Steel piston crown and/or combustion engine components with dynamic thermal insulation coating and method of making and using such a coating |
| CN108118278A (en) * | 2016-11-29 | 2018-06-05 | 沈阳黎明航空发动机(集团)有限责任公司 | One kind is used for the low guide vane method for preparing heat barrier coating of IC10 alloys |
| CN110629152A (en) * | 2019-10-09 | 2019-12-31 | 中国人民解放军第五七一九工厂 | Spraying method of thermal barrier coating with longitudinal cracks for combustion chamber |
Non-Patent Citations (1)
| Title |
|---|
| 王铁军 等: "《先进燃气轮机设计制造基础专著系列 热障涂层强度理论与检测技术》", 31 December 2016, 西安交通大学出版社 * |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113106532A (en) * | 2021-04-07 | 2021-07-13 | 江苏源清动力技术有限公司 | Process for removing thermal barrier coating of thermal component of aero-engine and gas turbine |
| CN113957378A (en) * | 2021-10-22 | 2022-01-21 | 中国航发贵州黎阳航空动力有限公司 | Coating repair method for turbine blade |
| CN113969405A (en) * | 2021-10-27 | 2022-01-25 | 中国航发贵州黎阳航空动力有限公司 | Method for removing NiCrAlY coating of aeroengine blade |
| CN113981457A (en) * | 2021-10-27 | 2022-01-28 | 中国航发贵州黎阳航空动力有限公司 | Method for removing failure coating of single crystal high-pressure turbine blade |
| CN114990468A (en) * | 2022-06-08 | 2022-09-02 | 哈尔滨汽轮机厂有限责任公司 | Preparation method of gas turbine combustor transition section thermal barrier coating for protecting gas film divergence hole |
| CN115961248A (en) * | 2022-12-07 | 2023-04-14 | 中国航发动力股份有限公司 | Single crystal blade thermal barrier coating repairing method |
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