CN109653805A - The air film hole and thermal barrier coating matching process of guide vane of high pressure turbine - Google Patents
The air film hole and thermal barrier coating matching process of guide vane of high pressure turbine Download PDFInfo
- Publication number
- CN109653805A CN109653805A CN201811497744.3A CN201811497744A CN109653805A CN 109653805 A CN109653805 A CN 109653805A CN 201811497744 A CN201811497744 A CN 201811497744A CN 109653805 A CN109653805 A CN 109653805A
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- China
- Prior art keywords
- diameter
- air film
- film hole
- high pressure
- pressure turbine
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- 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.)
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Classifications
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- 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
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- 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
This application provides a kind of air film holes of guide vane of high pressure turbine and thermal barrier coating matching process, comprising: designs the diameter design value of air film hole；Guide vane of high pressure turbine is sprayed with predetermined process parameter, to form thermal barrier coating；The second diameter after first diameter and spraying before measuring the air film hole spraying respectively；According to the first diameter and the second diameter, shrinkage cavity rule is determined；The air film hole in the guide vane of high pressure turbine is processed, until the processing diameter of air film hole is equal to the diameter design value；The guide vane of high pressure turbine is processed according to the diameter of the air film hole after test, and carries out thermal boundary spraying according to the predetermined process parameter.
Description
Technical field
This application involves aero-engine technology fields, specifically provide the air film hole and heat of a kind of guide vane of high pressure turbine
Barrier coating matching process.
Background technique
Gaseous film control and thermal barrier coating are two kinds of important technologies of guide vane of high pressure turbine thermal protection, air film hole aperture and
Thermal barrier coating thickness is an important factor for influencing cooling effect and heat-insulating capability, and current technology is first to process air film hole,
After spray thermal barrier coating, when carrying out thermal barrier coating spraying, powdered coating can be deposited in air film hole, reduce aperture, shadow
Ring Film Cooling.
It not fully achieved design requirement, air film according to the guide vane of high pressure turbine that current technology scheme is processed
Bore dia and thermal barrier coating thickness are both less than design value, its due cooling cannot be played by leading to gaseous film control and thermal barrier coating all
Effect and heat-insulating capability, cooling deficiency cause the actual temperature of turbo blade matrix to be higher than expected design, reduce blade use
Reliability, while leaf longevity being brought to shorten, change the high problem of part rate, improve the cost of engine.
Summary of the invention
At least one in order to solve the above-mentioned technical problem, this application provides a kind of air film holes of guide vane of high pressure turbine
With thermal barrier coating matching process, comprising:
Step 1, the diameter design value of air film hole is designed；
Step 2, guide vane of high pressure turbine is sprayed with predetermined process parameter, to form thermal barrier coating；
Step 3, the second diameter after first diameter and spraying before measuring the air film hole spraying respectively；
Step 4, according to the first diameter and the second diameter, shrinkage cavity rule is determined；
Step 5, step 2 to step 4 is tested as a wheel, according to
dq=d0+Δd
The air film hole in the guide vane of high pressure turbine is processed, repeats step 2 to step 4, Zhi Daoqi
The processing diameter of fenestra is equal to the diameter design value；
Wherein, q is test round, dqIt is the diameter of air film hole after testing, Δ d is shrinkage cavity value；
The guide vane of high pressure turbine is processed according to the diameter of the air film hole after test, and according to described default
Technological parameter carries out thermal boundary spraying.
According at least one embodiment of the application, the predetermined process parameter includes: default spray angle, default coating
Thickness, default torch parameters, default spray gun walking path.
According at least one embodiment of the application, the air film hole, will be tangential throughout the guide vane of high pressure turbine
The air film hole that position is identical, radial position is different is classified as a row,
According to the first diameter and the second diameter, shrinkage cavity rule is determined, comprising:
Shrinkage cavity value is obtained according to the following formula:
Δ d=dBefore-dAfterwards,
Wherein, dBeforeIt is the diameter before spraying, dAfterwardsIt is the diameter after spraying；
Then in the guide vane of high pressure turbine m be vented fenestra average shrinkage cavity value:
Wherein, n is the number of m exhaust fenestra, and k is testpieces group number, Δ dmijIt is jth group testpieces m exhaust fenestra
The shrinkage cavity value in i-th of hole；
According to the average shrinkage cavity value, shrinkage cavity rule is obtained:
IfThen unify withAs shrinkage cavity value；
IfThen the m shrinkage cavity value for arranging i-th of air film hole is individually handled.
In the air film hole and thermal barrier coating matching process of guide vane of high pressure turbine provided by the embodiments of the present application, Neng Goubao
Card gaseous film control and thermal barrier coating meet design requirement simultaneously, improve cooling effect and heat-insulating capability, reduce turbine leaf chip base
Temperature improves blade service life, and part bring cost is changed in reduction, improves designed reliability.
Detailed description of the invention
Fig. 1 is the air film hole of guide vane of high pressure turbine provided by the embodiments of the present application and the stream of thermal barrier coating matching process
Journey schematic diagram.
Specific embodiment
The application is described in further detail with reference to the accompanying drawings and examples.It is understood that this place is retouched
The specific embodiment stated is used only for explaining related application, rather than the restriction to this application.It also should be noted that in order to
Convenient for description, part relevant to the application is illustrated only in attached drawing.
It should be noted that in the absence of conflict, the features in the embodiments and the embodiments of the present application can phase
Mutually combination.The application is described in detail below with reference to the accompanying drawings and in conjunction with the embodiments.
Fig. 1 is the air film hole of guide vane of high pressure turbine provided by the embodiments of the present application and the stream of thermal barrier coating matching process
Journey schematic diagram.
As shown in Figure 1, the matching process the following steps are included:
Step 1, the diameter design value of air film hole is designed.
In the present embodiment, it can be determined by experiment air film hole shrinkage cavity rule caused by thermal barrier coating spraying in advance, so
The processing diameter for amplifying air film hole in air film hole machined afterwards, the satisfaction design of air film bore dia is wanted after guaranteeing spraying thermal barrier coating
It asks, note air film hole diameter design value is d0。
Step 2, guide vane of high pressure turbine is sprayed with predetermined process parameter, to form thermal barrier coating.
Wherein, predetermined process parameter includes: default spray angle, default coating layer thickness, default torch parameters, default spray gun
Walking path.
As an alternative embodiment, spray painting control program can be write, using manipulator according to program command come
The spraying for completing thermal barrier coating, solidifies the spraying process of thermal barrier coating.
Step 3, the second diameter after first diameter and spraying before measurement air film hole sprays respectively.
For example, carrying out thermal barrier coating spraying using the cured program in step 2, the air film hole before measurement spraying is straight respectively
Air film bore dia after diameter, spraying.
Step 4, according to the first diameter and the second diameter, shrinkage cavity rule is determined.
In the present embodiment, air film hole is throughout guide vane of high pressure turbine, chordwise location is identical, radial position is different
Air film hole is classified as a row, obtains shrinkage cavity value according to the following formula:
Δ d=dBefore-dAfterwards,
Wherein, dBeforeIt is the diameter before spraying, dAfterwardsIt is the diameter after spraying；
Then in guide vane of high pressure turbine m be vented fenestra average shrinkage cavity value:
Wherein, n is the number of m exhaust fenestra, and k is testpieces group number, Δ dmijIt is jth group testpieces m exhaust fenestra
The shrinkage cavity value in i-th of hole；
Then shrinkage cavity rule are as follows:
IfThen unify withAs shrinkage cavity value；
IfThen the m shrinkage cavity value for arranging i-th of air film hole is individually handled.
Step 5, step 2 to step 4 is tested as a wheel, according to
dq=d0+Δd
Air film hole in guide vane of high pressure turbine is processed, repeats step 2 to step 4, until adding for air film hole
Work diameter is equal to the diameter design value.
Wherein, q is test round, dqIt is the diameter of air film hole after testing, Δ d is shrinkage cavity value；
Step 6, the guide vane of high pressure turbine is processed according to the diameter of the air film hole after test, and according to institute
It states predetermined process parameter and carries out thermal boundary spraying.
According to diameter dqThe air film hole machined of guide vane of high pressure turbine is carried out, and according to spray procedure cured in test
Carry out thermal barrier coating spraying, can guarantee finished product air film bore dia and thermal barrier coating thickness simultaneously meet design requirement, realization
Air film hole and thermal barrier coating matched design.
So far, it has been combined preferred embodiment shown in the drawings and describes the technical solution of the application, still, this field
Technical staff is it is easily understood that the protection scope of the application is expressly not limited to these specific embodiments.Without departing from this
Under the premise of the principle of application, those skilled in the art can make equivalent change or replacement to the relevant technologies feature, these
Technical solution after change or replacement is fallen within the protection scope of the application.
Claims (3)
1. a kind of air film hole of guide vane of high pressure turbine and thermal barrier coating matching process characterized by comprising
Step 1, the diameter design value of air film hole is designed；
Step 2, guide vane of high pressure turbine is sprayed with predetermined process parameter, to form thermal barrier coating；
Step 3, the second diameter after first diameter and spraying before measuring the air film hole spraying respectively；
Step 4, according to the first diameter and the second diameter, shrinkage cavity rule is determined；
Step 5, step 2 to step 4 is tested as a wheel, according to
dq=d0+Δd
The air film hole in the guide vane of high pressure turbine is processed, repeats step 2 to step 4, until air film hole
Processing diameter be equal to the diameter design value；
Wherein, q is test round, dqIt is the diameter of air film hole after testing, Δ d is shrinkage cavity value；
The guide vane of high pressure turbine is processed according to the diameter of the air film hole after test, and according to the predetermined process
Parameter carries out thermal boundary spraying.
2. the air film hole of guide vane of high pressure turbine according to claim 1 and thermal barrier coating matching process, feature exist
In the predetermined process parameter includes: default spray angle, default coating layer thickness, default torch parameters, default spray gun walking road
Diameter.
3. the air film hole of guide vane of high pressure turbine according to claim 1 and thermal barrier coating matching process, feature exist
In the air film hole is throughout the guide vane of high pressure turbine, chordwise location is identical, radial position the is different air film hole
It is classified as a row,
According to the first diameter and the second diameter, shrinkage cavity rule is determined, comprising:
Shrinkage cavity value is obtained according to the following formula:
Δ d=dBefore-dAfterwards,
Wherein, dBeforeIt is the diameter before spraying, dAfterwardsIt is the diameter after spraying；
Then in the guide vane of high pressure turbine m be vented fenestra average shrinkage cavity value:
Wherein, n is the number of m exhaust fenestra, and k is testpieces group number, Δ dmijIt is jth group testpieces m exhaust fenestra i-th
The shrinkage cavity value in a hole；
According to the average shrinkage cavity value, shrinkage cavity rule is obtained:
IfThen unify withAs shrinkage cavity value；
IfThen the m shrinkage cavity value for arranging i-th of air film hole is individually handled.
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 true CN109653805A (en) | 2019-04-19 |
| CN109653805B CN109653805B (en) | 2021-08-17 |
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| Application Number | Title | Priority Date | Filing Date |
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| 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 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112380696A (en) * | 2020-11-13 | 2021-02-19 | 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 |
|---|
| 康军卫等: "航空发动机涡轮叶片缩孔问题及控制研究", 《沈阳航空航天大学学报》 * |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112380696A (en) * | 2020-11-13 | 2021-02-19 | 中国航发沈阳发动机研究所 | Turbine air cooling blade design method based on additive manufacturing process |
| 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 |
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| CN109653805B (en) | 2021-08-17 |
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