CN109653805B - Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating - Google Patents
Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating Download PDFInfo
- Publication number
- CN109653805B CN109653805B CN201811497744.3A CN201811497744A CN109653805B CN 109653805 B CN109653805 B CN 109653805B CN 201811497744 A CN201811497744 A CN 201811497744A CN 109653805 B CN109653805 B CN 109653805B
- Authority
- CN
- China
- Prior art keywords
- diameter
- film hole
- pressure turbine
- spraying
- turbine guide
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
-
- 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
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Mechanical Engineering (AREA)
- Physics & Mathematics (AREA)
- Plasma & Fusion (AREA)
- Chemical Kinetics & Catalysis (AREA)
- Materials Engineering (AREA)
- Metallurgy (AREA)
- Organic Chemistry (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application provides a method for matching a film hole and a thermal barrier coating of a high-pressure turbine guide blade, which comprises the following steps: designing a diameter design value of the air film hole; spraying the high-pressure turbine guide blade according to preset process parameters to form a thermal barrier coating; respectively measuring a first diameter of the air film hole before spraying and a second diameter of the air film hole after spraying; determining a shrinkage cavity rule according to the first diameter and the second diameter; machining the film hole on the high-pressure turbine guide vane until the machining diameter of the film hole is equal to the designed diameter value; and machining the high-pressure turbine guide blade according to the diameter of the tested air film hole, and performing thermal barrier spraying according to the preset process parameters.
Description
Technical Field
The application relates to the technical field of aero-engines, and particularly provides a method for matching a film hole and a thermal barrier coating of a high-pressure turbine guide blade.
Background
The air film cooling and the thermal barrier coating are two important technologies for the thermal protection of the guide blade of the high-pressure turbine, the aperture of an air film hole and the thickness of the thermal barrier coating are important factors influencing the cooling effect and the heat insulation capability, the existing technology is to process the air film hole firstly and then spray the thermal barrier coating, and when the thermal barrier coating is sprayed, the powdery coating is accumulated in the air film hole, so that the aperture is reduced, and the air film cooling effect is influenced.
The high-pressure turbine guide blade processed according to the current technical scheme does not completely meet the design requirement, the diameter of an air film hole and the thickness of a thermal barrier coating are smaller than the design values, so that the air film cooling and the thermal barrier coating cannot exert the due cooling effect and the thermal insulation capability, the actual temperature of a turbine blade substrate is higher than the design expectation due to insufficient cooling, the use reliability of the blade is reduced, the problems of shortened service life and high replacement rate of the blade are caused, and the cost of an engine is improved.
Disclosure of Invention
In order to solve at least one of the above technical problems, the present application provides a method for matching a film hole of a high-pressure turbine guide blade with a thermal barrier coating, comprising:
step 1, designing a diameter design value of a gas film hole;
step 2, spraying the high-pressure turbine guide blade according to preset process parameters to form a thermal barrier coating;
step 3, respectively measuring a first diameter of the air film hole before spraying and a second diameter of the air film hole after spraying;
step 4, determining a shrinkage cavity rule according to the first diameter and the second diameter;
step 5, taking the steps 2 to 4 as a round of test according to
dq＝d0+Δd
Processing the film hole on the high-pressure turbine guide vane, and repeating the steps 2 to 4 until the processing diameter of the film hole is equal to the designed diameter value;
wherein q is the test run, dqIs the diameter of the gas film hole after the test, and Δ d is the shrinkage cavity value;
and machining the high-pressure turbine guide blade according to the diameter of the tested air film hole, and performing thermal barrier spraying according to the preset process parameters.
According to at least one embodiment of the present application, the preset process parameters include: presetting a spraying angle, presetting a coating thickness, presetting spray gun parameters and presetting a spray gun walking path.
According to at least one embodiment of the present application, the film holes are distributed over the high pressure turbine guide vane, the film holes having the same chord-wise position and different radial positions are grouped into a row,
determining a shrinkage cavity law according to the first diameter and the second diameter, comprising:
the shrinkage cavity value was obtained as follows:
Δd＝dfront side-dRear end，
Wherein d isFront sideIs the diameter before spraying, dRear endIs the diameter after spraying;
the average shrinkage cavity value of the m-th exhaust film hole on the high-pressure turbine guide blade is as follows:
wherein n is the number of the m-th exhaust film holes, k is the number of test piece groups, and Δ dmijIs the shrinkage cavity value of the ith hole of the mth exhaust film hole of the jt group of test pieces;
and obtaining a shrinkage cavity rule according to the average shrinkage cavity value:
if it is
Then is unified with
As a shrinkage factor;
if it is
The shrinkage cavity value of the ith row of film holes is processed separately.
According to the matching method of the air film hole and the thermal barrier coating of the high-pressure turbine guide blade, the air film cooling and the thermal barrier coating can meet the design requirements at the same time, the cooling effect and the heat insulation capacity are improved, the temperature of the base body of the turbine blade is reduced, the service life of the blade is prolonged, the cost caused by replacing the blade is reduced, and the design reliability is improved.
Drawings
FIG. 1 is a schematic flow chart of a method for matching a film hole and a thermal barrier coating of a high-pressure turbine guide vane provided by an embodiment of the application.
Detailed Description
The present application will be described in further detail with reference to the following drawings and examples. It is to be understood that the specific embodiments described herein are merely illustrative of the relevant application and are not limiting of the application. It should be noted that, for convenience of description, only the portions related to the present application are shown in the drawings.
It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The present application will be described in detail below with reference to the embodiments with reference to the attached drawings.
FIG. 1 is a schematic flow chart of a method for matching a film hole and a thermal barrier coating of a high-pressure turbine guide vane provided by an embodiment of the application.
As shown in fig. 1, the matching method includes the steps of:
step 1, designing a diameter design value of a gas film hole.
In this embodiment, the shrinkage rule of the air film hole caused by thermal barrier coating spraying can be determined in advance through experiments, then the processing diameter of the air film hole is enlarged during air film hole processing, it is ensured that the diameter of the air film hole after thermal barrier coating spraying meets the design requirement, and the design value of the diameter of the air film hole is recorded as d0。
And 2, spraying the high-pressure turbine guide blade according to preset process parameters to form a thermal barrier coating.
Wherein, the preset process parameters comprise: presetting a spraying angle, presetting a coating thickness, presetting spray gun parameters and presetting a spray gun walking path.
As an alternative embodiment, a spraying control program can be written, and a manipulator is adopted to complete the spraying of the thermal barrier coating according to a program command, and the spraying process of the thermal barrier coating is cured.
And 3, respectively measuring the first diameter of the air film hole before spraying and the second diameter of the air film hole after spraying.
For example, the thermal barrier coating is sprayed by the curing procedure in step 2, and the diameter of the film hole before spraying and the diameter of the film hole after spraying are measured respectively.
And 4, determining a shrinkage cavity rule according to the first diameter and the second diameter.
In this embodiment, the film holes are distributed over the high-pressure turbine guide vane, the film holes with the same chord-wise position and different radial positions are grouped into a row, and the shrinkage cavity value is obtained according to the following formula:
Δd＝dfront side-dRear end，
Wherein d isFront sideIs the diameter before spraying, dRear endIs the diameter after spraying;
the average shrinkage cavity value of the m-th exhaust film hole on the high-pressure turbine guide blade is as follows:
wherein n is the number of the m-th exhaust film holes, k is the number of test piece groups, and Δ dmijIs the shrinkage cavity value of the ith hole of the mth exhaust film hole of the jt group of test pieces;
the shrinkage cavity rule is as follows:
if it is
Then is unified with
As a shrinkage factor;
if it is
The shrinkage cavity value of the ith row of film holes is processed separately.
Step 5, taking the steps 2 to 4 as a round of test according to
dq＝d0+Δd
And (4) processing the film hole on the guide vane of the high-pressure turbine, and repeating the steps from 2 to 4 until the processing diameter of the film hole is equal to the designed diameter value.
Wherein q is the test run, dqIs the diameter of the gas film hole after the test, and Δ d is the shrinkage cavity value;
and 6, machining the high-pressure turbine guide blade according to the diameter of the tested air film hole, and performing thermal barrier spraying according to the preset process parameters.
According to diameter dqAnd processing an air film hole of the high-pressure turbine guide blade, and spraying a thermal barrier coating according to a spraying program solidified in a test, so that the diameter of the air film hole of a finished product and the thickness of the thermal barrier coating can meet the design requirement at the same time, and the matching design of the air film hole and the thermal barrier coating is realized.
So far, the technical solutions of the present application have been described in connection with the preferred embodiments shown in the drawings, but it is easily understood by those skilled in the art that the scope of the present application is obviously not limited to these specific embodiments. Equivalent changes or substitutions of related technical features can be made by those skilled in the art without departing from the principle of the present application, and the technical scheme after the changes or substitutions will fall into the protection scope of the present application.
Claims (1)
1. A method for matching a film hole and a thermal barrier coating of a high-pressure turbine guide vane is characterized by comprising the following steps:
step 1, designing a diameter design value of a gas film hole;
step 2, spraying the high-pressure turbine guide blade according to preset process parameters to form a thermal barrier coating;
step 3, respectively measuring a first diameter of the air film hole before spraying and a second diameter of the air film hole after spraying;
step 4, determining a shrinkage cavity rule according to the first diameter and the second diameter;
step 5, taking the steps 2 to 4 as a round of test according to
dq＝d0+△d
Processing the film hole on the high-pressure turbine guide vane, and repeating the steps 2 to 4 until the processing diameter of the film hole is equal to the designed diameter value;
wherein q is the test run, dqIs the diameter of the gas film hole after the test, and Δ d is the shrinkage cavity value;
processing the high-pressure turbine guide blade according to the diameter of the tested air film hole, and performing thermal barrier spraying according to the preset process parameters;
the preset process parameters comprise: presetting a spraying angle, a coating thickness, a spray gun parameter and a spray gun walking path;
the air film holes are distributed over the high-pressure turbine guide vane, the air film holes with the same chord direction position and different radial positions are grouped into a row,
determining a shrinkage cavity law according to the first diameter and the second diameter, comprising:
the shrinkage cavity value was obtained as follows:
△d＝dfront side-dRear end，
Wherein d isFront sideIs the diameter before spraying, dRear endIs the diameter after spraying;
the average shrinkage cavity value of the m-th exhaust film hole on the high-pressure turbine guide blade is as follows:
wherein n is the number of the mth exhaust film hole, k is the number of test piece groups, and Δ dmijIs the shrinkage cavity value of the ith hole of the mth exhaust film hole of the jt group of test pieces;
and obtaining a shrinkage cavity rule according to the average shrinkage cavity value:
if it is
Then is unified with
As a shrinkage factor;
if it is
The shrinkage cavity value of the ith row of film holes is processed separately.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811497744.3A CN109653805B (en) | 2018-12-07 | 2018-12-07 | Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811497744.3A CN109653805B (en) | 2018-12-07 | 2018-12-07 | Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN109653805A CN109653805A (en) | 2019-04-19 |
| CN109653805B true CN109653805B (en) | 2021-08-17 |
Family
ID=66113829
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201811497744.3A Active CN109653805B (en) | 2018-12-07 | 2018-12-07 | Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN109653805B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112380696B (en) * | 2020-11-13 | 2022-08-19 | 中国航发沈阳发动机研究所 | Turbine air cooling blade design method based on additive manufacturing process |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH04236757A (en) * | 1991-01-17 | 1992-08-25 | Mitsubishi Heavy Ind Ltd | Method for masking turbine blade |
| US5702288A (en) * | 1995-08-30 | 1997-12-30 | United Technologies Corporation | Method of removing excess overlay coating from within cooling holes of aluminide coated gas turbine engine components |
| US5771577A (en) * | 1996-05-17 | 1998-06-30 | General Electric Company | Method for making a fluid cooled article with protective coating |
| DE102005015153A1 (en) * | 2005-03-31 | 2006-10-05 | Alstom Technology Ltd. | Method of renewing cooling aperture e.g. of gas turbine involves applying new lamination on component in aperture zone in length-wise section |
| CN101120156A (en) * | 2005-04-12 | 2008-02-06 | 西门子公司 | Component with film cooling holes |
| CN107762565A (en) * | 2016-08-16 | 2018-03-06 | 通用电气公司 | Has the porose component for turbogenerator |
-
2018
- 2018-12-07 CN CN201811497744.3A patent/CN109653805B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH04236757A (en) * | 1991-01-17 | 1992-08-25 | Mitsubishi Heavy Ind Ltd | Method for masking turbine blade |
| US5702288A (en) * | 1995-08-30 | 1997-12-30 | United Technologies Corporation | Method of removing excess overlay coating from within cooling holes of aluminide coated gas turbine engine components |
| US5771577A (en) * | 1996-05-17 | 1998-06-30 | General Electric Company | Method for making a fluid cooled article with protective coating |
| DE102005015153A1 (en) * | 2005-03-31 | 2006-10-05 | Alstom Technology Ltd. | Method of renewing cooling aperture e.g. of gas turbine involves applying new lamination on component in aperture zone in length-wise section |
| CN101120156A (en) * | 2005-04-12 | 2008-02-06 | 西门子公司 | Component with film cooling holes |
| CN107762565A (en) * | 2016-08-16 | 2018-03-06 | 通用电气公司 | Has the porose component for turbogenerator |
Non-Patent Citations (1)
| Title |
|---|
| 航空发动机涡轮叶片缩孔问题及控制研究;康军卫等;《沈阳航空航天大学学报》;20180831;第35卷(第4期);第60-66页 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN109653805A (en) | 2019-04-19 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| EP2950942B1 (en) | Coating process for gas turbine engine component with cooling holes | |
| US10907502B2 (en) | System and method of fabricating and repairing a gas turbine component | |
| US20080226462A1 (en) | Cast features for a turbine engine airfoil | |
| EP2935792B1 (en) | Vane device for a gas turbine and corresponding method of manufacturing | |
| US20160083829A1 (en) | Coating process | |
| US20080095635A1 (en) | Vane with enhanced heat transfer | |
| JP2009517576A (en) | Repair method for shroud segment of gas turbine | |
| US11407067B2 (en) | Method for repairing a part | |
| CN109653805B (en) | Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating | |
| JP2016020689A (en) | Methods for producing strain sensors on turbine components | |
| JP4981828B2 (en) | Method of forming HVOF sprayed coating layer and turbine member holding device | |
| WO2014046901A1 (en) | System and method for machining aircraft components | |
| KR20170054571A (en) | Template for forming cooling passages in a turbine engine component | |
| US11755791B2 (en) | Aircraft component qualification system and process | |
| US20180347378A1 (en) | Adaptively opening cooling pathway | |
| CN110647052A (en) | Variable cycle engine mode switching self-adaptive identity card model construction method | |
| CN109261466A (en) | A kind of modified functionally gradient coating design method of consideration deposition | |
| US20160290645A1 (en) | Axisymmetric offset of three-dimensional contoured endwalls | |
| US9919391B2 (en) | Method for manufacturing a turbine assembly | |
| US20180347368A1 (en) | Airfoil and method of fabricating same | |
| JP2009518571A (en) | How to make a familiar layer | |
| KR102375979B1 (en) | Spray apparatus | |
| US8869739B2 (en) | Wheel coating method and apparatus for a turbine | |
| US20090107211A1 (en) | Airfoil shot peening test strips | |
| US20150004308A1 (en) | Method for creating a textured bond coat surface |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |