JPH04236757A - Method for masking turbine blade - Google Patents
Method for masking turbine bladeInfo
- Publication number
- JPH04236757A JPH04236757A JP1573191A JP1573191A JPH04236757A JP H04236757 A JPH04236757 A JP H04236757A JP 1573191 A JP1573191 A JP 1573191A JP 1573191 A JP1573191 A JP 1573191A JP H04236757 A JPH04236757 A JP H04236757A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- cooling hole
- pin
- cooling
- masking
- 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.)
- Withdrawn
Links
- 230000000873 masking effect Effects 0.000 title abstract description 16
- 238000000034 method Methods 0.000 title description 11
- 238000001816 cooling Methods 0.000 claims abstract description 39
- 239000011248 coating agent Substances 0.000 claims abstract description 17
- 238000000576 coating method Methods 0.000 claims abstract description 17
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 claims abstract description 15
- 229910002804 graphite Inorganic materials 0.000 claims abstract description 15
- 239000010439 graphite Substances 0.000 claims abstract description 15
- 230000003647 oxidation Effects 0.000 claims abstract description 7
- 238000007254 oxidation reaction Methods 0.000 claims abstract description 7
- 238000007750 plasma spraying Methods 0.000 claims description 9
- 238000010438 heat treatment Methods 0.000 claims description 4
- 239000000463 material Substances 0.000 abstract description 9
- 238000010290 vacuum plasma spraying Methods 0.000 abstract 2
- 238000005336 cracking Methods 0.000 abstract 1
- 230000002542 deteriorative effect Effects 0.000 abstract 1
- 239000011247 coating layer Substances 0.000 description 3
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000003116 impacting effect Effects 0.000 description 1
- 239000002184 metal Substances 0.000 description 1
- 230000000116 mitigating effect Effects 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
- Coating By Spraying Or Casting (AREA)
Abstract
Description
【０００１】0001
【産業上の利用分野】本発明は、ガスタービンのタービ
ン動静翼の減圧プラズマ溶射施行時に適用される冷却孔
のマスキング方法に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a cooling hole masking method applied when applying low pressure plasma spraying to turbine rotor and stationary blades of a gas turbine.
【０００２】0002
【従来の技術】ガスタービンのタービン動静翼には、図
４，図５に示すようにタービン入口ガス温度の上昇に伴
うタービン動静翼１，２の母材冷却強化の観点から、動
静翼のプロファイル部及びシュラウド部にはφ０．６〜
φ１．１ｍｍ程度の冷却孔が穿設されている。このター
ビン動静翼には、近年図７に示すように減圧プラズマ溶
射による耐高温酸化によるコーティング４が施行されて
いるが、減圧プラズマ溶射の施行時には減圧雰囲気中の
プラズマ炎の温度が高く、冷却孔に対する適当なマスキ
ング法がなかった。なお図７における３は母材を示す。[Prior Art] As shown in FIGS. 4 and 5, the turbine rotor and stationary blades of a gas turbine have a profile that is designed to strengthen the cooling of the base material of the turbine rotor and stationary blades 1 and 2 as the turbine inlet gas temperature rises. φ0.6~ for the part and shroud part
Cooling holes with a diameter of about 1.1 mm are bored. In recent years, these turbine rotor and stationary blades have been coated with high-temperature oxidation-resistant coating 4 by low-pressure plasma spraying, as shown in Figure 7. However, when low-pressure plasma spraying is applied, the temperature of the plasma flame in the reduced-pressure atmosphere is high, and the cooling holes There was no suitable masking method for this. Note that 3 in FIG. 7 indicates the base material.
【０００３】0003
【発明が解決しようとする課題】ところで前述のように
従来技術では、冷却孔に対する適当なマスキング方法が
なかったことから、冷却孔内径部に図７に示すようにタ
ービン翼母材３表面にコーティングしたコーティング層
４のコーティング巻込みを生じ、冷却孔径の縮小が生じ
、これにより、ガスタービン動静翼の冷却性能を低下さ
せる不具合があった。[Problems to be Solved by the Invention] However, as mentioned above, in the prior art, there was no appropriate masking method for the cooling holes, so it was necessary to coat the inner diameter of the cooling holes on the surface of the turbine blade base material 3 as shown in FIG. This causes the coating layer 4 to become engulfed, resulting in a reduction in the diameter of the cooling hole, which causes a problem of lowering the cooling performance of the gas turbine rotor and stationary blades.
【０００４】本発明は上記冷却孔詰りの防止のため黒鉛
を用いたマスキング方法により前記不具合点を解消する
新たなタービン翼のマスキング方法を提供しようとする
ものである。[0004] The present invention aims to provide a new method for masking a turbine blade, which eliminates the above-mentioned disadvantages by using a masking method using graphite to prevent clogging of the cooling holes.
【０００５】[0005]
【課題を解決するための手段】このため本発明のタービ
ン翼のマスキング方法は、冷却孔を有するタービン翼に
おいて、前記冷却孔をテーパ形状に構成した黒鉛ピンを
用いて止栓した後、タービン翼表面に減圧プラズマ溶射
により耐酸化コーティングを施し、しかる後前記黒鉛ピ
ンを加熱燃焼して除去することを特徴としている。[Means for Solving the Problems] Therefore, in the turbine blade masking method of the present invention, in a turbine blade having a cooling hole, after the cooling hole is plugged using a graphite pin configured in a tapered shape, the turbine blade is It is characterized in that an oxidation-resistant coating is applied to the surface by low-pressure plasma spraying, and then the graphite pins are removed by heating and burning.
【０００６】[0006]
【作用】上述の本発明のマスキング方法は、タービン動
静翼表面への減圧プラズマ溶射施行時に、タービン翼に
穿設した冷却孔にテーパ状の黒鉛ピンを使用して止栓す
ることにより、冷却孔へのコーティングの巻込みを防止
することができる。この冷却孔へのマスキングに対し、
金属等のピンを使用すると、このピンに対しても、コー
ティング材が付着し、コーティング後のピンの除去作業
が困難であるばかりか、ピンの除去時の衝撃により、コ
ーティング皮膜中にクラックを発生する恐れがあり、コ
ーティング剥離の原因にもなり得る。しかし、本発明の
ように黒鉛ピンを使用した場合、コーティング後の加熱
（５００℃程度）により、黒鉛ピンは燃焼し、コーティ
ング皮膜に衝撃を与えることなく、ピンを除去できるこ
とから、コーティング皮膜内のクラック発生の心配がな
い。[Operation] The above-mentioned masking method of the present invention uses tapered graphite pins to stop the cooling holes drilled in the turbine blades during low-pressure plasma spraying on the surfaces of the turbine rotor and stationary blades. It is possible to prevent the coating from getting caught up in the coating. For masking this cooling hole,
If a metal pin is used, the coating material will adhere to the pin, making it difficult to remove the pin after coating, and the impact of removing the pin may cause cracks in the coating film. This may cause the coating to peel off. However, when graphite pins are used as in the present invention, the graphite pins are burned by heating (approximately 500°C) after coating, and the pins can be removed without impacting the coating film. There is no need to worry about cracks occurring.
【０００７】[0007]
【実施例】以下図面により本発明の１実施例について説
明すると、図１〜図３は本発明マスキング方法の作業工
程を示す説明図、図４は冷却孔を有するガスタービン動
翼の外観図例を示し、図５は同じくガスタービン静翼の
外観図例を示す。図４，図５に示すようにガスタービン
の動翼及び静翼にはタービン入口ガス温度の上昇に伴う
動静翼の母材冷却のための冷却孔が翼面に穿設されてお
り、ガスタービンの仕様によっては、静翼のシュラウド
面に穿設されているものもある。この冷却孔６の孔径は
０．６〜１．１ｍｍ程度の大きさである。断面でみると
図６のように外面より内面に向けてストレートの孔が穿
設されている。[Embodiment] An embodiment of the present invention will be described below with reference to the drawings. Figs. 1 to 3 are explanatory diagrams showing the working steps of the masking method of the present invention, and Fig. 4 is an example of an external view of a gas turbine rotor blade having cooling holes. Similarly, FIG. 5 shows an example of an external view of a gas turbine stationary blade. As shown in Figures 4 and 5, cooling holes are drilled in the blade surfaces of the moving and stationary blades of the gas turbine to cool the base material of the moving and stationary blades as the turbine inlet gas temperature rises. Depending on the specifications, holes may be drilled into the shroud surface of the stator vane. The hole diameter of this cooling hole 6 is approximately 0.6 to 1.1 mm. When viewed in cross section, as shown in Figure 6, a straight hole is drilled from the outer surface toward the inner surface.
【０００８】本発明マスキング方法は、図１〜図３の順
序で施行されるもので、図１に示すように１０°程度の
テーパをつけたマスキング用の黒鉛製のピン５を先づ母
材３に穿設した各冷却孔６にそれぞれ差し込む。この時
、母材表面からのピン５の突出量は０．３ｍｍ程度にす
る。〔コーティング時のプラズマジェットのシェイドエ
リア（陰影領域）を極力少なくする為。〕この後、減圧
プラズマ溶射により、タービン翼表面に耐酸化コーティ
ングを施行し、母材表面にコーティング層４を形成する
と、ピン５を挿入した冷却孔６入口箇所の断面は図２の
ようになり、ピン５は残った状態となる。次ぎにこの黒
鉛製ピン５を大気中でバーナ等で加熱（５００℃程度）
することにより、酸化・消失させる。この時の断面を示
したのが図３である。The masking method of the present invention is carried out in the order shown in FIGS. 1 to 3. As shown in FIG. 3 into each of the cooling holes 6 drilled therein. At this time, the amount of protrusion of the pin 5 from the surface of the base material is set to about 0.3 mm. [To minimize the plasma jet shade area during coating. [After this, an oxidation-resistant coating is applied to the turbine blade surface by low-pressure plasma spraying and a coating layer 4 is formed on the base material surface, and the cross section of the inlet of the cooling hole 6 into which the pin 5 is inserted becomes as shown in Figure 2. , pin 5 remains. Next, this graphite pin 5 is heated in the atmosphere with a burner, etc. (about 500℃)
By doing so, it is oxidized and disappears. FIG. 3 shows a cross section at this time.
【０００９】[0009]
【発明の効果】以上述べたように本発明のタービン翼の
マスキング方法を採用することにより、冷却孔を有する
ガスタービン動静翼への減圧プラズマ溶射による耐酸化
コーティング時の冷却孔詰りが防止でき、ガスタービン
動静翼の冷却性能を低下させることがない。また、冷却
孔へ挿入するのが黒鉛製造のピンである為、その後大気
中で加熱することにより、コーティング皮膜にクラック
等を発生させることなく、容易に孔中より除去（消失）
できる。更には、テーパ型の黒鉛ピンを使用することに
より冷却孔近傍断面は図３に示すように、コーティング
皮膜にテーパが付くことから端部における熱衝撃緩和効
果があり、冷却孔近傍の剥離防止効果がある。As described above, by employing the turbine blade masking method of the present invention, cooling hole clogging can be prevented when oxidation-resistant coating is applied by low-pressure plasma spraying to gas turbine rotor and stationary blades having cooling holes. The cooling performance of gas turbine rotor and stationary blades is not reduced. In addition, since the graphite pin is inserted into the cooling hole, it can be easily removed (disappeared) from the hole by heating it in the atmosphere without causing any cracks in the coating film.
can. Furthermore, by using a tapered graphite pin, the cross-section near the cooling hole is tapered as shown in Figure 3, which has a thermal shock mitigation effect at the end and prevents peeling near the cooling hole. There is.
【図１】本発明の１実施例に係る黒鉛ピンによる冷却孔
マスキング状態を示す断面図である。FIG. 1 is a sectional view showing a cooling hole masking state by a graphite pin according to an embodiment of the present invention.
【図２】本発明の１実施例に係る減圧プラズマ溶射後の
冷却孔の断面図である。FIG. 2 is a cross-sectional view of a cooling hole after reduced pressure plasma spraying according to an embodiment of the present invention.
【図３】本発明の１実施例に係る黒鉛ピンの加熱除去後
の冷却孔の断面図である。FIG. 3 is a cross-sectional view of a cooling hole after heat removal of a graphite pin according to an embodiment of the present invention.
【図４】冷却孔を有するガスタービン動翼の外観図であ
る。FIG. 4 is an external view of a gas turbine rotor blade having cooling holes.
【図５】冷却孔を有するガスタービン静翼の外観図であ
る。FIG. 5 is an external view of a gas turbine stationary blade having cooling holes.
【図６】冷却孔の断面図である。FIG. 6 is a cross-sectional view of a cooling hole.
【図７】従来のマスキングのない場合の冷却孔へのコー
ティング巻込み状態を示す断面図である。FIG. 7 is a cross-sectional view showing a state in which a coating is wrapped around a cooling hole without conventional masking.
１ ガスタービン動翼 ２ ガスタービン静翼 ３ タービン翼母材 ４ コーティング層 ５ マスキング用黒鉛ピン ６ 冷却孔 1 Gas turbine rotor blades 2 Gas turbine stator blade 3 Turbine blade base material 4 Coating layer 5 Graphite pin for masking 6 Cooling holes
Claims (1)
前記冷却孔をテーパ形状に構成した黒鉛ピンを用いて止
栓した後タービン翼表面に減圧プラズマ溶射により耐酸
化コーティングを施し、しかる後前記黒鉛ピンを加熱燃
焼して除去することを特徴とするタービン翼のマスキン
グ方法。Claim 1: A turbine blade having cooling holes,
A turbine characterized in that after the cooling hole is plugged using a graphite pin having a tapered shape, an oxidation-resistant coating is applied to the surface of the turbine blade by low pressure plasma spraying, and then the graphite pin is removed by heating and burning. How to mask the wings.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP1573191A JPH04236757A (en) | 1991-01-17 | 1991-01-17 | Method for masking turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP1573191A JPH04236757A (en) | 1991-01-17 | 1991-01-17 | Method for masking turbine blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH04236757A true JPH04236757A (en) | 1992-08-25 |
Family
ID=11896914
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP1573191A Withdrawn JPH04236757A (en) | 1991-01-17 | 1991-01-17 | Method for masking turbine blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH04236757A (en) |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0861918A1 (en) * | 1994-01-26 | 1998-09-02 | United Technologies Corporation | Improved pack coating process for particles containing small passageways |
| US6247895B1 (en) * | 1998-06-17 | 2001-06-19 | United Technologies Corporation | Locking member for processing a flow directing assembly |
| WO2004013368A1 (en) * | 2002-08-02 | 2004-02-12 | Mitsubishi Heavy Industries, Ltd. | Method for forming heat shielding film, masking pin and tail pipe of combustor |
| WO2007134916A1 (en) * | 2006-05-19 | 2007-11-29 | Siemens Aktiengesellschaft | Method for preparing a component consisting of an electroconductive base material in order to carry out an erosive process |
| JP2011106448A (en) * | 2009-11-16 | 2011-06-02 | Siemens Ag | Coating method for component with partially closed holes and method for opening the coated hole |
| JP2014109274A (en) * | 2012-12-04 | 2014-06-12 | General Electric Co <Ge> | Coated article |
| CN109653805A (en) * | 2018-12-07 | 2019-04-19 | 中国航发沈阳发动机研究所 | The air film hole and thermal barrier coating matching process of guide vane of high pressure turbine |
| US10272461B2 (en) | 2017-04-04 | 2019-04-30 | General Electric Company | Method for masking cooling passages |
| US10815783B2 (en) | 2018-05-24 | 2020-10-27 | General Electric Company | In situ engine component repair |
-
1991
- 1991-01-17 JP JP1573191A patent/JPH04236757A/en not_active Withdrawn
Cited By (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0861918A1 (en) * | 1994-01-26 | 1998-09-02 | United Technologies Corporation | Improved pack coating process for particles containing small passageways |
| US6247895B1 (en) * | 1998-06-17 | 2001-06-19 | United Technologies Corporation | Locking member for processing a flow directing assembly |
| US9051879B2 (en) | 2002-08-02 | 2015-06-09 | Mitsubishi Heavy Industries, Ltd. | Thermal barrier coating method, masking pin and combustor transition piece |
| US8722144B2 (en) | 2002-08-02 | 2014-05-13 | Mitsubishi Heavy Industries, Ltd. | Thermal barrier coating method, masking pin and combustor transition piece |
| WO2004013368A1 (en) * | 2002-08-02 | 2004-02-12 | Mitsubishi Heavy Industries, Ltd. | Method for forming heat shielding film, masking pin and tail pipe of combustor |
| DE10392994B4 (en) * | 2002-08-02 | 2006-12-14 | Mitsubishi Heavy Industries, Ltd. | Thermal barrier coating method, its use and cover pins |
| CN100368588C (en) * | 2002-08-02 | 2008-02-13 | 三菱重工业株式会社 | Method for forming heat shielding film, masking pin and tail pipe of combustor |
| DE10392994C5 (en) * | 2002-08-02 | 2013-08-14 | Mitsubishi Heavy Industries, Ltd. | Thermal barrier coating method and its use |
| WO2007134620A1 (en) * | 2006-05-19 | 2007-11-29 | Siemens Aktiengesellschaft | Method for preparing a component consisting of an electroconductive base material in order to carry out an erosive process |
| WO2007134916A1 (en) * | 2006-05-19 | 2007-11-29 | Siemens Aktiengesellschaft | Method for preparing a component consisting of an electroconductive base material in order to carry out an erosive process |
| JP2011106448A (en) * | 2009-11-16 | 2011-06-02 | Siemens Ag | Coating method for component with partially closed holes and method for opening the coated hole |
| US8980372B2 (en) | 2009-11-16 | 2015-03-17 | Siemens Aktiengesellschaft | Process for coating a component having partially closed holes and process for opening the holes |
| JP2014109274A (en) * | 2012-12-04 | 2014-06-12 | General Electric Co <Ge> | Coated article |
| US10272461B2 (en) | 2017-04-04 | 2019-04-30 | General Electric Company | Method for masking cooling passages |
| US10815783B2 (en) | 2018-05-24 | 2020-10-27 | General Electric Company | In situ engine component repair |
| CN109653805A (en) * | 2018-12-07 | 2019-04-19 | 中国航发沈阳发动机研究所 | The air film hole and thermal barrier coating matching process of guide vane of high pressure turbine |
| CN109653805B (en) * | 2018-12-07 | 2021-08-17 | 中国航发沈阳发动机研究所 | Method for matching air film hole of high-pressure turbine guide vane with thermal barrier coating |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| A300 | Application deemed to be withdrawn because no request for examination was validly filed |
Free format text: JAPANESE INTERMEDIATE CODE: A300 Effective date: 19980514 |