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 PDF

Info

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
Application number
CN201811497744.3A
Other languages
Chinese (zh)
Other versions
CN109653805A (en
Inventor
曾令玉
王富强
张志强
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
AECC Shenyang Engine Research Institute
Original Assignee
AECC Shenyang Engine Research Institute
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by AECC Shenyang Engine Research Institute filed Critical AECC Shenyang Engine Research Institute
Priority to CN201811497744.3A priority Critical patent/CN109653805B/en
Publication of CN109653805A publication Critical patent/CN109653805A/en
Application granted granted Critical
Publication of CN109653805B publication Critical patent/CN109653805B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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

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

Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating
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:
Figure BDA0001897325470000021
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:
Figure BDA0001897325470000022
if it is
Figure BDA0001897325470000023
Then is unified with
Figure BDA0001897325470000024
As a shrinkage factor;
if it is
Figure BDA0001897325470000025
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:
Figure BDA0001897325470000031
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:
Figure BDA0001897325470000041
if it is
Figure BDA0001897325470000042
Then is unified with
Figure BDA0001897325470000043
As a shrinkage factor;
if it is
Figure BDA0001897325470000044
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:
Figure FDA0003025182590000011
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:
Figure FDA0003025182590000021
if it is
Figure FDA0003025182590000022
Then is unified with
Figure FDA0003025182590000023
As a shrinkage factor;
if it is
Figure FDA0003025182590000024
The shrinkage cavity value of the ith row of film holes is processed separately.
CN201811497744.3A 2018-12-07 2018-12-07 Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating Active CN109653805B (en)

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)

* Cited by examiner, † Cited by third party
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)

* Cited by examiner, † Cited by third party
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

Patent Citations (6)

* Cited by examiner, † Cited by third party
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)

* Cited by examiner, † Cited by third party
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
US8197184B2 (en) Vane with enhanced heat transfer
EP2935792B1 (en) Vane device for a gas turbine and corresponding method of manufacturing
JP2009517576A (en) Repair method for shroud segment of gas turbine
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
EP3133252B1 (en) Rotor tip clearance
WO2014046901A1 (en) System and method for machining aircraft components
JP4981828B2 (en) Method of forming HVOF sprayed coating layer and turbine member holding device
JP2011256865A (en) Method, system, and computer program product for life management of gas turbine
JP2017096935A (en) Methods for monitoring components
KR20170054571A (en) Template for forming cooling passages in a turbine engine component
US20180347378A1 (en) Adaptively opening cooling pathway
US20150354406A1 (en) Blade outer air seal and method of manufacture
US20140193664A1 (en) Recoating process and recoated turbine blade
US9919391B2 (en) Method for manufacturing a turbine assembly
EP3591172B1 (en) Aircraft component qualification system and process
US20180347368A1 (en) Airfoil and method of fabricating same
US20200317370A1 (en) Aircraft component repair system and process
US20160290645A1 (en) Axisymmetric offset of three-dimensional contoured endwalls
US10533433B2 (en) Turbine blade with hot-corrosion-resistant coating
US8869739B2 (en) Wheel coating method and apparatus for a turbine
US20150004308A1 (en) Method for creating a textured bond coat surface
US10995619B2 (en) Airfoil and method of fabricating same

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