JP2003172102A - Turbine blade, its production method, and its thermal barrier coat separation determining method - Google Patents

Turbine blade, its production method, and its thermal barrier coat separation determining method

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Publication number
JP2003172102A
JP2003172102A JP2001374542A JP2001374542A JP2003172102A JP 2003172102 A JP2003172102 A JP 2003172102A JP 2001374542 A JP2001374542 A JP 2001374542A JP 2001374542 A JP2001374542 A JP 2001374542A JP 2003172102 A JP2003172102 A JP 2003172102A
Authority
JP
Japan
Prior art keywords
turbine blade
cooling gas
turbine
thermal barrier
barrier coat
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.)
Pending
Application number
JP2001374542A
Other languages
Japanese (ja)
Inventor
Akitatsu Masaki
彰樹 正木
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.)
IHI Corp
Original Assignee
IHI Corp
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
Family has litigation
First worldwide family litigation filed litigation Critical https://patents.darts-ip.com/?family=19183089&utm_source=google_patent&utm_medium=platform_link&utm_campaign=public_patent_search&patent=JP2003172102(A) "Global patent litigation dataset” by Darts-ip is licensed under a Creative Commons Attribution 4.0 International License.
Application filed by IHI Corp filed Critical IHI Corp
Priority to JP2001374542A priority Critical patent/JP2003172102A/en
Priority to CA002385932A priority patent/CA2385932C/en
Priority to US10/146,964 priority patent/US20030108424A1/en
Priority to EP02012119A priority patent/EP1318273B1/en
Priority to DE60216405T priority patent/DE60216405T2/en
Publication of JP2003172102A publication Critical patent/JP2003172102A/en
Pending legal-status Critical Current

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Classifications

    • 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
    • C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • C23C28/30—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer
    • C23C28/32—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one pure metallic layer
    • C23C28/321—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one pure metallic layer with at least one metal alloy layer
    • C23C28/3215—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one pure metallic layer with at least one metal alloy layer at least one MCrAlX layer
    • 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
    • C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • 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
    • C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • C23C28/30—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer
    • C23C28/32—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one pure metallic layer
    • C23C28/321—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one pure metallic layer with at least one metal alloy layer
    • 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
    • C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • C23C28/30—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer
    • C23C28/34—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one inorganic non-metallic material layer, e.g. metal carbide, nitride, boride, silicide layer and their mixtures, enamels, phosphates and sulphates
    • C23C28/345—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one inorganic non-metallic material layer, e.g. metal carbide, nitride, boride, silicide layer and their mixtures, enamels, phosphates and sulphates with at least one oxide layer
    • 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
    • C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • C23C28/30—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer
    • C23C28/34—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one inorganic non-metallic material layer, e.g. metal carbide, nitride, boride, silicide layer and their mixtures, enamels, phosphates and sulphates
    • C23C28/345—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one inorganic non-metallic material layer, e.g. metal carbide, nitride, boride, silicide layer and their mixtures, enamels, phosphates and sulphates with at least one oxide layer
    • C23C28/3455—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one inorganic non-metallic material layer, e.g. metal carbide, nitride, boride, silicide layer and their mixtures, enamels, phosphates and sulphates with at least one oxide layer with a refractory ceramic layer, e.g. refractory metal oxide, ZrO2, rare earth oxides or a thermal barrier system comprising at least one refractory oxide layer
    • 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
    • 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
    • C23C4/02—Pretreatment of the material to be coated, e.g. for coating on selected surface areas
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/182—Transpiration cooling
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/90—Coating; Surface treatment

Landscapes

  • Chemical & Material Sciences (AREA)
  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Materials Engineering (AREA)
  • Organic Chemistry (AREA)
  • Metallurgy (AREA)
  • Chemical Kinetics & Catalysis (AREA)
  • Inorganic Chemistry (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Plasma & Fusion (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

<P>PROBLEM TO BE SOLVED: To provide a turbine blade operable by raising a temperature of high temperature gas, its manufacturing method, and a method of determining separation of a thermal barrier coat. <P>SOLUTION: Instead of a conventional turbine blade having an outer peripheral surface of a wing shape, this turbine blade has a turbine blade body with a wing-shaped cross section having an inside void arranged inside so as to pass cooling gas, and a plurality of cooling gas ports communicating with the inside void, and arranged on an outer peripheral surface, and the thermal barrier coat being a heat insulating material layer for covering the outer peripheral surface of the turbine blade body. The thermal barrier coat blocks up at least several cooling gas holes. <P>COPYRIGHT: (C)2003,JPO

Description

【発明の詳細な説明】Detailed Description of the Invention

【０００１】[0001]

【発明の属する技術分野】本発明は、タービンエンジン
等で用いられるタービン翼に係る。特にタービン翼の外
周面に施されたサーマルバリアコートの構造、そのター
ビン翼の製造方法及びサーマルバリアコートの剥離の判
定方法に関する。TECHNICAL FIELD The present invention relates to a turbine blade used in a turbine engine or the like. In particular, the present invention relates to a structure of a thermal barrier coat provided on the outer peripheral surface of a turbine blade, a method for manufacturing the turbine blade, and a method for determining peeling of the thermal barrier coat.

【０００２】[0002]

【従来の技術】タービンエンジンに用いられるタービン
翼は高温ガス雰囲気で所定の空力性能と機械的性能を発
揮することを求められる。温度が上がるとタービン翼の
材料は機械的性能が低下するので、タービン翼を所定の
温度以下に維持する必要がある。従来、タービン翼の温
度を所定値以下にするために、タービン翼の外周面に冷
却ガス口を設け、内部空隙を通して、タービン翼の外周
面に冷却ガスを吹きだす。さらに、タービン翼の外周面
にサーマルバリアコートを施し、高温ガスが直接にター
ビン翼に接触することを防止し、タービン翼の材料の温
度が上昇するのを防止する。2. Description of the Related Art Turbine blades used in turbine engines are required to exhibit predetermined aerodynamic performance and mechanical performance in a high temperature gas atmosphere. Since the material of the turbine blade deteriorates in mechanical performance as the temperature rises, it is necessary to keep the turbine blade at a predetermined temperature or lower. Conventionally, in order to keep the temperature of the turbine blade below a predetermined value, a cooling gas port is provided on the outer peripheral surface of the turbine blade, and the cooling gas is blown out to the outer peripheral surface of the turbine blade through an internal gap. Further, a thermal barrier coat is applied to the outer peripheral surface of the turbine blade to prevent the high temperature gas from directly contacting the turbine blade and to prevent the temperature of the material of the turbine blade from rising.

【０００３】従来のタービン翼の構造を、図をもとに、
説明する。図４は、従来のタービン翼の構造図である。
図５は、従来のタービン翼の断面図である。図６は、従
来のタービン翼の断面詳細図である。タービン翼２は、
タービン動翼部材１やタービン静翼部材（図示せず）の
一部である。図はタービン動翼部材１を例に示してい
る。タービン動翼部材１は、タービン翼２とタービン動
翼基部３とタービン動翼取付部４とで構成されている。
タービン翼２は、外周面が翼形状をした柱状構造体であ
り、根元部をタービン動翼基部３に固着し、先端部を自
由にしている。一方、タービン静翼部材の場合は、先端
部もタービン静翼の他方の基部に固着している。冷却ガ
ス供給口５と冷却ガス排出口６が、タービン動翼取付部
４に設けられ、タービン動翼基部３を貫通して後述する
タービン翼の内部空隙に連通している。The structure of a conventional turbine blade is shown in FIG.
explain. FIG. 4 is a structural diagram of a conventional turbine blade.
FIG. 5 is a cross-sectional view of a conventional turbine blade. FIG. 6 is a detailed sectional view of a conventional turbine blade. The turbine blade 2
It is a part of the turbine moving blade member 1 and the turbine stationary blade member (not shown). The figure shows the turbine blade member 1 as an example. The turbine rotor blade member 1 includes a turbine blade 2, a turbine rotor blade base portion 3, and a turbine rotor blade mounting portion 4.
The turbine blade 2 is a columnar structure having an outer peripheral surface in the shape of a blade, and has a root portion fixed to the turbine rotor blade base portion 3 and a free end portion. On the other hand, in the case of the turbine vane member, the tip is also fixed to the other base of the turbine vane. The cooling gas supply port 5 and the cooling gas discharge port 6 are provided in the turbine rotor blade mounting portion 4, penetrate the turbine rotor blade base portion 3 and communicate with the internal air gap of the turbine blade described later.

【０００４】タービン翼２は、タービン翼本体１０とサ
ーマルバリアコート２０とで構成される。タービン翼本
体１０は、内部空隙１２と複数の冷却ガス口１３を有す
る翼形状断面の柱状構造体１１である。内部空隙１２
は、冷却ガスが通過する様に柱状構造体１１の内部に設
けられた空隙であり、柱状構造物を複数の部屋に分け、
その部屋は相互に連通している。タービン翼の先縁側の
内部空隙が冷却ガス供給口５に連通し、タービン翼の後
縁側の内部空隙が冷却ガス排出口６に連通する。冷却ガ
ス口１３は、タービン翼の外側表面の温度が上がりそう
な領域に密に設けられている。冷却ガス口１３は、柱状
構造体１１の外周面に設けられた孔である。冷却ガス口
１３は、内部空隙に連通している。冷却ガス口１３の直
径は、例えば０．４〜０．５ｍｍである。The turbine blade 2 comprises a turbine blade body 10 and a thermal barrier coat 20. The turbine blade body 10 is a columnar structure 11 having an internal cavity 12 and a plurality of cooling gas ports 13 and having a blade-shaped cross section. Internal void 12
Is a void provided inside the columnar structure 11 so that the cooling gas passes therethrough, and the columnar structure is divided into a plurality of chambers,
The rooms are in communication with each other. An inner space on the leading edge side of the turbine blade communicates with the cooling gas supply port 5, and an inner space on the trailing edge side of the turbine blade communicates with the cooling gas discharge port 6. The cooling gas ports 13 are densely provided in the region where the temperature of the outer surface of the turbine blade is likely to rise. The cooling gas port 13 is a hole provided on the outer peripheral surface of the columnar structure 11. The cooling gas port 13 communicates with the internal space. The diameter of the cooling gas port 13 is, for example, 0.4 to 0.5 mm.

【０００５】サーマルバリアコート２０は、前記タービ
ン翼本体の外周面を覆う断熱材層であり、トップコート
２１とインターフェース２２とボンドコート２３とで構
成される。ボンドコート２３は、タービン翼本体の表面
に施される下地処理層であり、例えば、ＮｉＣｏＣｒＡ
ｌＹやＰｔＡｌ等で代表される耐酸化コーティングであ
る。インターフェース２２は、トップコート２１とボン
ドコート２３の間の層であり、例えば、Ａｌ２Ｏ３等で
ある。トップコート２１は、外周面を形成する層であ
り、例えばＺｒＯ２・７％Ｙ2Ｏ3等で代表されるセラミ
ックスコーティングである。例えば、トップコート２１
の厚みは１００〜３００μ、インターフェース２２の厚
みは１μ、ボンドコート２３の厚みは１００μである。The thermal barrier coat 20 is a heat insulating material layer covering the outer peripheral surface of the turbine blade body, and is composed of a top coat 21, an interface 22 and a bond coat 23. The bond coat 23 is a base treatment layer applied to the surface of the turbine blade body, and is, for example, NiCoCrA.
It is an oxidation resistant coating represented by 1Y or PtAl. The interface 22 is a layer between the top coat 21 and the bond coat 23, and is, for example, Al 2 O 3 or the like. The top coat 21 is a layer that forms the outer peripheral surface, and is a ceramic coating represented by, for example, ZrO 2.7% Y 2 O 3 . For example, top coat 21
Has a thickness of 100 to 300 μ, the interface 22 has a thickness of 1 μ, and the bond coat 23 has a thickness of 100 μ.

【０００６】次に従来のタービン翼の作用を説明する。
タービン動翼部材１がタービンディスクに取り付けられ
る。タービンエンジンが定常運転すると、タービン翼の
周囲に高温ガスが流れ、空力を受けてタービンディスク
が回転する。冷却ガスが、タービンディスクを介して、
冷却ガス供給口５に供給される。冷却ガスは、冷却ガス
供給口５から、タービン動翼取付部４とタービン動翼基
部３を通って、内部空隙１２に入る。冷却ガスの一部
は、冷却ガス口１３を通ってタービン翼２の外表面に吹
きだす。残りの冷却ガスは、複数の部屋を流れて、ター
ビン動翼取付部４とタービン動翼基部３を通って、冷却
ガス排出口６から出る。タービン翼の外周に吹きだした
冷却ガスはタービン翼の表面に沿って流れ、周囲の高温
ガスがタービン翼に直接触れるのを防止する。さらに、
サーマルバリアコーティング２０は、タービン翼の外周
を覆い、高温ガスの熱が直接にタービン翼本体に伝わる
のを防止する。（図６（Ｂ）） タービン翼本体は、周囲の高温ガスの熱が入りにくいの
で、材料が高温になるのを防止でき、その機械的強度を
維持できる。Next, the operation of the conventional turbine blade will be described.
The turbine rotor blade member 1 is attached to the turbine disk. When the turbine engine operates steadily, high-temperature gas flows around the turbine blades and receives aerodynamic force to rotate the turbine disk. Cooling gas passes through the turbine disk
It is supplied to the cooling gas supply port 5. The cooling gas enters the internal space 12 from the cooling gas supply port 5 through the turbine rotor blade mounting portion 4 and the turbine rotor blade base portion 3. A part of the cooling gas is blown out to the outer surface of the turbine blade 2 through the cooling gas port 13. The remaining cooling gas flows through the plurality of chambers, passes through the turbine rotor blade mounting portion 4 and the turbine rotor blade base portion 3, and exits from the cooling gas discharge port 6. The cooling gas blown to the outer periphery of the turbine blade flows along the surface of the turbine blade, and prevents the surrounding high temperature gas from directly contacting the turbine blade. further,
The thermal barrier coating 20 covers the outer periphery of the turbine blade and prevents heat of the hot gas from being directly transferred to the turbine blade body. (FIG. 6 (B)) Since the heat of the surrounding high temperature gas is hard to enter in the turbine blade body, it is possible to prevent the material from reaching a high temperature and maintain its mechanical strength.

【０００７】[0007]

【発明が解決しようとする課題】上述のタービン翼を使
用した場合、タービンエンジンの運転を続けると、高温
ガスの酸素分子が、トップコート２１にしみ込み、イン
ターフェース２２で酸化アルミを析出する。インターフ
ェース２２の層厚みが、増加して、約１０μ程度に成長
すると、トップコート２１が剥がれてしまうことが経験
的に知られている。（図６（Ｃ）） 従来は、定期点検において、上記の現象をおこしたター
ビン翼を交換していた。トップコート２１が剥がれる
と、剥がれた部分を通して、高温ガスの熱エネルギーが
タービン翼本体１０に入り、タービン翼本体の温度が所
望の温度以上に上昇する。従って、従来は、高温ガスの
温度を低めにして運用していた。例えば、トップコート
２１の剥がれがなければ１３００度Ｃで運転でできるタ
ービン翼であれば、トップコートの剥がれがあることを
考慮して高温ガス温度を１０００度Ｃにして運転してい
た。従って、今以上にタービンエンジンの効率を上げら
れないという問題があった。また、トップコートの剥が
れは、定期点検による目視検査によらなければ発見でき
ないという問題があった。When the turbine blade described above is used, when the turbine engine continues to be operated, oxygen molecules in the high temperature gas permeate the top coat 21 and deposit aluminum oxide at the interface 22. It is empirically known that the top coat 21 is peeled off when the layer thickness of the interface 22 increases and grows to about 10 μm. (FIG. 6 (C)) Conventionally, in a regular inspection, the turbine blade that caused the above phenomenon was replaced. When the top coat 21 is peeled off, the thermal energy of the high temperature gas enters the turbine blade body 10 through the peeled portion, and the temperature of the turbine blade body rises to a desired temperature or higher. Therefore, in the past, the high temperature gas was operated at a low temperature. For example, in the case of a turbine blade that can be operated at 1300 ° C. without peeling of the top coat 21, the high temperature gas temperature was set to 1000 ° C. in consideration of the peeling of the top coat. Therefore, there has been a problem that the efficiency of the turbine engine cannot be improved further. Further, there is a problem that the peeling of the top coat can be detected only by a visual inspection by a regular inspection.

【０００８】本発明は以上に述べた問題点に鑑み案出さ
れたもので、従来のタービン翼にかわって、高温ガスの
温度を上げて運用できるタービン翼とその製造方法とサ
ーマルバリアコートの剥離を判断する方法を提供しよう
とする。The present invention has been devised in view of the above-mentioned problems, and replaces a conventional turbine blade with a turbine blade that can be operated by raising the temperature of high-temperature gas, a method for manufacturing the turbine blade, and peeling of a thermal barrier coat. Trying to provide a way to judge.

【０００９】[0009]

【課題を解決するための手段】上記目的を達成するた
め、本発明に係る外周面が翼形状をしたタービン翼は、
冷却ガスが通過する様に内部に設けられた内部空隙と該
内部空隙に連通可能にし外周面に設けられた複数の冷却
ガス口とを有する翼形状断面のタービン翼本体と、前記
タービン翼本体の外周面を覆う断熱材層であるサーマル
バリアコートとを備え、前記サーマルバリアコートが前
記冷却ガス口の少なくとも幾つかを塞いでいるものとし
た。In order to achieve the above object, a turbine blade having an outer peripheral surface according to the present invention has a blade shape,
A turbine blade main body having a blade-shaped cross section, which has an internal void provided inside for allowing cooling gas to pass therethrough and a plurality of cooling gas ports provided on the outer peripheral surface and which are capable of communicating with the internal void, and the turbine blade main body A thermal barrier coat that is a heat insulating material layer that covers the outer peripheral surface is provided, and the thermal barrier coat covers at least some of the cooling gas ports.

【００１０】上記本発明の構成により、タービン翼本体
が内部空隙と複数の冷却ガス口とを有する翼形状断面の
柱状構造体であり、断熱材層であるサーマルバリアコー
トが前記タービン翼本体の外周面を覆い、前記サーマル
バリアコートが前記冷却ガス口の少なくとも幾つかを塞
いでいるので、通常運用時には、冷却ガスが内部空隙を
経由して冷却ガス口から吹きだしてタービン翼本体を冷
却し、サーマルバリアコートが剥離すると、剥離した箇
所の冷却ガス口からも冷却ガスが吹きだして、通常運用
時よりも多い冷却ガスでタービン翼を冷却し、タービン
翼の温度が上昇するのを防止できる。With the above-described structure of the present invention, the turbine blade body is a columnar structure having a blade-shaped cross section having an internal void and a plurality of cooling gas ports, and the thermal barrier coat, which is a heat insulating material layer, is the outer periphery of the turbine blade body. Since the surface is covered and the thermal barrier coat covers at least some of the cooling gas ports, during normal operation, the cooling gas is blown out from the cooling gas ports through the internal voids to cool the turbine blade body, and the thermal When the barrier coat is peeled off, the cooling gas is also blown out from the cooling gas port at the peeled portion, and it is possible to prevent the temperature of the turbine blade from rising by cooling the turbine blade with more cooling gas than during normal operation.

【００１１】さらに、本発明に係るタービン翼は、前記
サーマルバリアコートが冷却ガス口の総数のうちの２０
％を越える数の冷却ガス口を塞いでいるものとした。上
記本発明の構成により、前記サーマルバリアコートが冷
却ガス口の総数のうちの２０％を越える数の冷却ガス口
を塞いでいるので、一部のサーマルバリアコートが剥離
すると、剥離した部分の冷却ガス口の数が増え、約２割
を超えて増量した冷却ガスでタービン翼を冷却し、ター
ビン翼の温度が上昇するのを防止できる。Further, in the turbine blade according to the present invention, the thermal barrier coat has 20 of the total number of cooling gas ports.
It is assumed that the cooling gas ports of which the number exceeds 100% are blocked. According to the configuration of the present invention, the thermal barrier coat blocks the cooling gas ports of which the number exceeds 20% of the total number of cooling gas ports. Therefore, when a part of the thermal barrier coat is peeled off, the peeled part is cooled. It is possible to prevent the temperature of the turbine blade from rising by cooling the turbine blade with the cooling gas that has increased the number of gas ports and increased to more than about 20%.

【００１２】また、上記目的を達成するため、本発明に
係る外周面が翼形状をしたタービン翼の製造方法は、冷
却ガスが通過する様に内部に設けられた内部空隙と該内
部空隙に連通可能にし外周面に設けられた複数の冷却ガ
ス口とを有する翼形状断面のタービン翼本体を用意する
タービン翼本体準備工程と、前記冷却ガス口のすくなく
とも幾つかに目止め材を詰める目止め工程と、目止め工
程の後でタービン翼本体の外側表面に断熱材層を設ける
コーティング工程と、を備えたものとした。In order to achieve the above object, in the method for manufacturing a turbine blade having an outer peripheral surface in the shape of a blade according to the present invention, an internal void provided inside to allow cooling gas to pass therethrough and a communication with the internal void are provided. Turbine blade body preparation step of preparing a turbine blade body having a blade-shaped cross section having a plurality of cooling gas ports provided on the outer peripheral surface, and a sealing step of filling a sealing material in at least some of the cooling gas ports And a coating step of providing a heat insulating material layer on the outer surface of the turbine blade body after the sealing step.

【００１３】上記本発明の構成により、タービン翼本体
準備工程で、内部空隙と複数の冷却ガス口とを有する翼
形状断面のタービン翼本体を用意し、目止め工程で前記
冷却ガス口のすくなくとも幾つかに目止め材を詰め、目
止め工程の後のコーティング工程で、タービン翼本体の
外側表面に断熱材層を設けるので、詰まった目止め材に
より断熱材が冷却ガス口を埋めることを防止して、断熱
材層をタービン翼本体の外側表面にコーティングし、内
部空隙と複数の冷却ガス口とを有する翼形状断面の柱状
構造体の外側表面を、断熱材層であるサーマルバリアコ
ートが覆い、前記サーマルバリアコートが前記冷却ガス
口の少なくとも幾つかを塞いでいるタービン翼を製作で
きる。According to the above-described structure of the present invention, in the turbine blade body preparation step, a turbine blade body having an airfoil-shaped cross section having an internal void and a plurality of cooling gas ports is prepared, and at least some of the cooling gas ports are filled in the filling step. The crab filling material is packed, and the heat insulating material layer is provided on the outer surface of the turbine blade body in the coating process after the filling processing, so that the clogging material prevents the heat insulating material from filling the cooling gas port. The outer surface of the turbine blade body is coated with a heat insulating material layer, and the outer surface of the columnar structure having a blade-shaped cross section having an internal void and a plurality of cooling gas ports is covered with a thermal barrier coat that is a heat insulating material layer. Turbine blades can be fabricated with the thermal barrier coat blocking at least some of the cooling gas ports.

【００１４】さらに、本発明に係るタービン翼の製造方
法は、前記目止め工程が、前記冷却ガス口の総数のうち
の２０％を越える数の冷却ガス口に目止めを施すものと
した。上記本発明の構成により、前記目止め工程で、前
記冷却ガス口の総数のうちの２０％を越える数の冷却ガ
ス口に目止めを施すので、内部空隙と複数の冷却ガス口
とを有する翼形状断面の柱状構造体の外側表面を、断熱
材層であるサーマルバリアコートが覆い、前記サーマル
バリアコートが冷却ガス口の総数のうちの２０％を越え
る数の冷却ガス口を塞いでいるタービン翼を製作でき
る。Further, in the turbine blade manufacturing method according to the present invention, in the plugging step, the cooling gas ports of which the number exceeds 20% of the total number of the cooling gas ports are plugged. According to the configuration of the present invention described above, in the plugging step, the cooling gas ports of which the number exceeds 20% of the total number of the cooling gas ports are plugged, so that the blade having the internal void and the plurality of cooling gas ports is provided. A turbine blade in which a thermal barrier coat, which is a heat insulating material layer, covers the outer surface of the columnar structure having a cross section, and the thermal barrier coat blocks cooling gas ports of which the number exceeds 20% of the total number of cooling gas ports. Can be manufactured.

【００１５】さらに、本発明に係るタービン翼の製造方
法は、前記コーティング工程の後に目止め材を除去する
目止め除去工程を備えるものとした。上記本発明の構成
により、前記コーティング工程の後の目止め除去工程
で、目止め材を除去するので、内部空隙と複数の冷却ガ
ス口とを有する翼形状断面の柱状構造体の外側表面を、
断熱材層であるサーマルバリアコートが覆い、前記サー
マルバリアコートにより塞がれた冷却ガス口にはなにも
詰まっていないので、サーマルバリアコートが剥離する
と、剥離した箇所の冷却ガス口からもすぐに冷却ガスが
吹きだして、通常運用時よりも多い冷却ガスでタービン
翼を冷却し、タービン翼の温度が上昇するのを防止でき
る様になったタービン翼を製作できる。Further, the turbine blade manufacturing method according to the present invention comprises a sealing removal step of removing the sealing material after the coating step. According to the configuration of the present invention, in the sealing removal step after the coating step, the sealing material is removed, so that the outer surface of the columnar structure having a blade-shaped cross section having an internal void and a plurality of cooling gas ports,
Since the thermal barrier coat that is the heat insulating material layer covers and the cooling gas port blocked by the thermal barrier coat is not clogged, when the thermal barrier coat is peeled off, it will be immediately from the cooling gas port at the peeled point. It is possible to manufacture a turbine blade that can prevent the temperature of the turbine blade from rising by cooling gas blowing out to cool the turbine blade with more cooling gas than in normal operation.

【００１６】また、上記目的を達成するため、本発明に
係るサーマルバリアコート剥離判断方法は、上述のター
ビン翼を使用したタービンを用意し、タービン翼に供給
する冷却ガスの消費量を検知し、前記消費量の変動によ
って前記サーマルバリアコートの剥離の有ることを判断
するものとした。In order to achieve the above object, the thermal barrier coat peeling determination method according to the present invention prepares a turbine using the above-mentioned turbine blade, detects the consumption amount of the cooling gas supplied to the turbine blade, It was determined that the thermal barrier coat was peeled off based on the fluctuation of the consumption amount.

【００１７】上記本発明の構成により、上述のタービン
翼を使用したタービンを用意し、タービン翼に供給する
冷却ガスの消費量を検知し、前記消費量の変動によって
前記サーマルバリアコートの剥離の有無を判断するの
で、タービンの運転中にサーマルバリアコートの剥離が
発生すると冷却ガスの消費量が増加し、その増加を検知
することでサーマルバリアコートの剥離の有ることを知
ることができ、例えばタービンの異常な温度上昇を防止
する手だてを直ちにとることが出来る。According to the configuration of the present invention, a turbine using the above-mentioned turbine blade is prepared, the consumption amount of the cooling gas supplied to the turbine blade is detected, and the thermal barrier coat is peeled off depending on the variation of the consumption amount. Therefore, if the thermal barrier coat peels off during operation of the turbine, the consumption of the cooling gas increases, and by detecting the increase, it can be known that the thermal barrier coat is peeled off. Immediate measures can be taken to prevent the abnormal temperature rise of the.

【００１８】[0018]

【発明の実施の形態】以下、本発明の好ましい実施形態
を、図面を参照して説明する。なお、各図において、共
通する部分には同一の符号を付し、重複した説明を省略
する。BEST MODE FOR CARRYING OUT THE INVENTION Preferred embodiments of the present invention will be described below with reference to the drawings. In each drawing, common portions are denoted by the same reference numerals, and redundant description will be omitted.

【００１９】本発明の実施形態に係るタービン翼の構造
を説明する。図１は、本発明の実施形態の構造図であ
る。図２は、本発明の実施形態の断面図である。図３
は、本発明の実施形態の断面詳細図である。The structure of the turbine blade according to the embodiment of the present invention will be described. FIG. 1 is a structural diagram of an embodiment of the present invention. FIG. 2 is a sectional view of an embodiment of the present invention. Figure 3
FIG. 3 is a detailed sectional view of an embodiment of the present invention.

【００２０】タービン翼２は、タービン動翼部材１やタ
ービン静翼部材（図示せず）の一部である。図はタービ
ン動翼部材１を例に示している。タービン動翼部材１
は、タービン翼２とタービン動翼基部３とタービン動翼
取付部４で構成されている。タービン翼２は、外周面が
翼形状をした柱状構造であり、根元部をタービン動翼基
部３に固着し、先端部を自由にしている。一方、タービ
ン静翼部材の場合は、先端部もタービン静翼の他方の基
部に固着している。冷却ガス供給口５と冷却ガス排出口
６が、タービン動翼取付部４に設けられ、タービン動翼
基部３を貫通して後述するタービン翼の内部空隙に連通
している。The turbine blade 2 is a part of the turbine rotor blade member 1 and the turbine stationary blade member (not shown). The figure shows the turbine blade member 1 as an example. Turbine blade member 1
Is composed of a turbine blade 2, a turbine rotor blade base 3, and a turbine rotor blade mounting portion 4. The turbine blade 2 has a columnar structure with an outer peripheral surface having a blade shape, and has a root portion fixed to the turbine rotor blade base portion 3 and a free end portion. On the other hand, in the case of the turbine vane member, the tip is also fixed to the other base of the turbine vane. The cooling gas supply port 5 and the cooling gas discharge port 6 are provided in the turbine rotor blade mounting portion 4, penetrate the turbine rotor blade base portion 3 and communicate with the internal air gap of the turbine blade described later.

【００２１】タービン翼２は、タービン翼本体１０とサ
ーマルバリアコート２０とで構成される。タービン翼本
体１０は、内部空隙１２と複数の冷却ガス口１３を有す
る翼形状断面の柱状構造体１１である。内部空隙１２
は、冷却ガスが通過する様に柱状構造体１１の内部に設
けられた空隙であり、柱状構造物を複数の部屋に分け、
その部屋は相互に連通している。タービン翼の先縁側の
内部空隙１２が冷却ガス供給口５に連通し、タービン翼
の後縁側の内部空隙１２が冷却ガス排出口６に連通す
る。The turbine blade 2 is composed of a turbine blade body 10 and a thermal barrier coat 20. The turbine blade body 10 is a columnar structure 11 having an internal cavity 12 and a plurality of cooling gas ports 13 and having a blade-shaped cross section. Internal void 12
Is a void provided inside the columnar structure 11 so that the cooling gas passes therethrough, and the columnar structure is divided into a plurality of chambers,
The rooms are in communication with each other. The inner space 12 on the leading edge side of the turbine blade communicates with the cooling gas supply port 5, and the inner space 12 on the trailing edge side of the turbine blade communicates with the cooling gas discharge port 6.

【００２２】冷却ガス口１３は、柱状構造体１１の外周
面に設けられた孔である。冷却ガス口１３は、内部空隙
に連通している。冷却ガス口１３の直径は、例えば０．
４〜０．５ｍｍである。冷却ガス口１３の幾つかは、後
述するトップコートにより覆われている。ここで、説明
の便宜のために、この冷却ガス口１３を特に消化栓冷却
ガス口１４と呼ぶ。冷却ガス口１３は、冷却が必要な箇
所に均等な間隔で設けられる。消化栓冷却ガス口１４の
数は、冷却ガス口１３の総数の２０％以上であるのが好
ましい。消化栓冷却ガス口１４は、複数の冷却ガス口１
３の内に均等に配されるのが好ましい。消火栓冷却ガス
口１４には、目止め材１５が詰まっている。目止め材の
材料は、後述するトップコートの付着工程での温度によ
り変わるが、例えば、エポキシ樹脂、ＰＥＥＫ樹脂、シ
リコンゴム等の樹脂や、ＣａＯ、ＡｌＦ3、ＮａＡｌ
Ｆ6、ＡｌＰＯ4、ＳｉＯ2等の単体化合物やこれらの混
合物である。例えば、ＡｌＰＯ4とＳｉＯ2との混合物で
もよい。又は、目止め材１５が予め除去され、消火栓冷
却ガス口１４には、なにも詰まっていない。The cooling gas port 13 is a hole provided on the outer peripheral surface of the columnar structure 11. The cooling gas port 13 communicates with the internal space. The diameter of the cooling gas port 13 is, for example, 0.
It is 4 to 0.5 mm. Some of the cooling gas ports 13 are covered with a top coat described later. Here, for convenience of description, this cooling gas port 13 is particularly referred to as a digestion plug cooling gas port 14. The cooling gas ports 13 are provided at equal intervals in places where cooling is required. The number of digestion plug cooling gas ports 14 is preferably 20% or more of the total number of cooling gas ports 13. The digestion plug cooling gas port 14 has a plurality of cooling gas ports 1.
It is preferable that they are evenly distributed among the three. A plug material 15 is clogged in the hydrant cooling gas port 14. The material of the sealing material varies depending on the temperature in the step of attaching the top coat, which will be described later. For example, resin such as epoxy resin, PEEK resin, silicon rubber, CaO, AlF 3 , NaAl, etc.
A single compound such as F 6 , AlPO 4 , and SiO 2 or a mixture thereof. For example, a mixture of AlPO 4 and SiO 2 may be used. Alternatively, the sealing material 15 is removed in advance, and the fire hydrant cooling gas port 14 is not clogged with anything.

【００２３】サーマルバリアコート２０は、前記タービ
ン翼本体の外周面を覆う断熱材層であり、トップコート
２１とインターフェース２２とボンドコート２３とで構
成される。ボンドコート２３は、タービン翼本体の表面
に施される下地処理層であり、例えば、ＮｉＣｏＣｒＡ
ｌＹやＰｔＡｌ等で代表される耐酸化コーティングであ
る。インターフェース２２は、トップコーと２１とボン
ドコート２３の間の層でり、例えば、Ａｌ2Ｏ3等であ
る。トップコート２１は、外周面を形成する層であり、
例えばＺｒＯ２・７％Ｙ2Ｏ3等で代表されるセラミック
スコーティングである。例えば、トップコート２１の厚
みは１００〜３００μ、インターフェース２２の厚みは
１μ、ボンドコート２３の厚みは１００μである。The thermal barrier coat 20 is a heat insulating material layer that covers the outer peripheral surface of the turbine blade body, and is composed of a top coat 21, an interface 22, and a bond coat 23. The bond coat 23 is a base treatment layer applied to the surface of the turbine blade body, and is, for example, NiCoCrA.
It is an oxidation resistant coating represented by 1Y or PtAl. The interface 22 is a layer between the top coat 21 and the bond coat 23, and is, for example, Al 2 O 3 or the like. The top coat 21 is a layer forming the outer peripheral surface,
For example, it is a ceramic coating represented by ZrO2.7% Y 2 O 3 . For example, the top coat 21 has a thickness of 100 to 300 μ, the interface 22 has a thickness of 1 μ, and the bond coat 23 has a thickness of 100 μ.

【００２４】次に本発明の実施形態であるタービン翼の
作用を説明する。タービン動翼部材１がタービンディス
クに取り付けられる。タービンエンジンが定常運転する
と、タービン翼の周囲に高温ガスが流れ、空力を受けて
タービンディスクが回転する。冷却ガスが、タービンデ
ィスクを介して、冷却ガス供給口５に供給される。冷却
ガスは、冷却ガス供給口５から、タービン動翼取付部４
とタービン動翼基部３を通って、内部空隙に入る。冷却
ガスの一部は、消火栓冷却ガス口１４を除く冷却ガス口
１３を通ってタービン翼２の外表面に吹きだす。残りの
冷却ガスは、複数の部屋を流れて、タービン動翼取付部
４とタービン動翼基部３を通って、冷却ガス排出口６か
ら出る。タービン翼の外周に吹きだした冷却ガス、ター
ビン翼の表面に沿って流れ、周囲の高温ガスがタービン
翼に直接触れるのを防止する。さらに、サーマルバリア
コーティング２０は、タービン翼の外周を覆い、高温ガ
スの熱が直接にタービン翼本体に伝わるのを防止する。
（図３（Ｂ））Next, the operation of the turbine blade according to the embodiment of the present invention will be described. The turbine rotor blade member 1 is attached to the turbine disk. When the turbine engine operates steadily, high-temperature gas flows around the turbine blades and receives aerodynamic force to rotate the turbine disk. The cooling gas is supplied to the cooling gas supply port 5 via the turbine disk. The cooling gas is supplied from the cooling gas supply port 5 to the turbine blade mounting portion 4
Through the turbine blade base 3 and into the internal void. A part of the cooling gas is blown to the outer surface of the turbine blade 2 through the cooling gas port 13 except the fire hydrant cooling gas port 14. The remaining cooling gas flows through the plurality of chambers, passes through the turbine rotor blade mounting portion 4 and the turbine rotor blade base portion 3, and exits from the cooling gas discharge port 6. Cooling gas blown to the outer circumference of the turbine blade and flows along the surface of the turbine blade to prevent the hot gas around it from directly contacting the turbine blade. Further, the thermal barrier coating 20 covers the outer periphery of the turbine blade and prevents heat of the hot gas from being directly transferred to the turbine blade body.
(Fig. 3 (B))

【００２５】一部の領域からトップコート２１が剥離す
ると、その領域にある消火栓冷却ガス口１４が露出す
る。目止め材１５が内部空隙の冷却ガスの圧力で押し出
され、冷却ガスが消火栓冷却ガス口１４から吹きだす。
冷却ガスが、トップコート２１の剥離したタービン翼本
体の外表面に沿って流れる。周囲の冷却ガス口１３から
吹き出た冷却ガスと消火栓冷却ガス口から吹き出た冷却
ガスが、トップコート２１の剥離した付近を覆い、高温
ガスがタービン翼本体の表面に触れるのを防止する。
（図３（Ｃ）） タービン翼本体は、周囲の高温ガスの熱が入りにくいの
で、材料が高温になるのを防止でき、その機械的強度を
維持できる。When the top coat 21 is peeled from a part of the area, the fire hydrant cooling gas port 14 in that area is exposed. The sealing material 15 is pushed out by the pressure of the cooling gas in the internal space, and the cooling gas is blown out from the fire hydrant cooling gas port 14.
Cooling gas flows along the outer surface of the stripped turbine blade body of the topcoat 21. The cooling gas blown out from the surrounding cooling gas port 13 and the cooling gas blown out from the fire hydrant cooling gas port cover the peeled vicinity of the top coat 21 and prevent the hot gas from touching the surface of the turbine blade body.
(FIG. 3 (C)) Since the heat of the surrounding high temperature gas is hard to enter the turbine blade body, it is possible to prevent the material from reaching a high temperature and maintain its mechanical strength.

【００２６】次に、本発明の実施形態の製造方法を工程
の順に説明する。 タービン翼本体準備工程：冷却ガスが通過する様に内部
に設けられた内部空隙１２と該内部空隙１２に連通し外
周面に設けられた複数の冷却ガス口とを有するタービン
翼本体を用意する。タービン翼本体の構造は前述したも
のと同じなので、説明を省略する。冷却ガス口１３の一
部は、消火栓冷却ガス口１４となるものである。Next, the manufacturing method of the embodiment of the present invention will be described in the order of steps. Turbine blade main body preparation step: A turbine blade main body having an internal void 12 provided therein so that a cooling gas passes therethrough and a plurality of cooling gas ports provided on the outer peripheral surface and communicating with the internal void 12 is prepared. Since the structure of the turbine blade body is the same as that described above, the description is omitted. A part of the cooling gas port 13 serves as a fire hydrant cooling gas port 14.

【００２７】目止め工程：前記冷却ガス口のすくなくと
も幾つかに目止め材を詰める。好ましくは、前記冷却ガ
ス口の総数のうちの２０％を越える数の冷却ガス口に目
止め材を詰める目止め材の材質は、後工程であるコーテ
ィング工程の処理温度により選択する。目止め材の材料
の耐熱温度は、後工程であるコーティング工程でタービ
ン翼２が置かれる雰囲気温度より高いことが必要であ
る。例えば、コーティングが溶射で行われる場合、ター
ビン翼本体１０の温度は比較的低いので、目止め材の材
料は、エポキシ樹脂、ＰＥＥＫ樹脂、シリコンゴム等の
樹脂が選択される。例えば、コーティングが蒸着でおこ
なわれる場合、タービン翼本体１０の温度が高温になる
ので、 目止め材の材料は、ＣａＯ、ＡｌＦ3、Ｎａ3Ａ
ｌＦ6、ＡｌＰＯ4、ＳｉＯ2等の単体化合物やこれらの
混合物から選ばれる。例えば、ＡｌＰＯ4とＳｉＯ2との
混合物でもよい。Filling process: Fill at least some of the cooling gas ports with a filling material. Preferably, the material of the sealing material that fills the cooling gas openings of more than 20% of the total number of the cooling gas openings with the sealing material is selected according to the processing temperature of the coating step which is the subsequent step. The heat-resistant temperature of the material of the sealing material needs to be higher than the ambient temperature in which the turbine blade 2 is placed in the coating step which is a post-step. For example, when the coating is performed by thermal spraying, the temperature of the turbine blade body 10 is relatively low, and therefore, the material of the sealing material is selected from epoxy resin, PEEK resin, silicone rubber and the like. For example, when the coating is performed by vapor deposition, the temperature of the turbine blade body 10 becomes high, so that the material for the sealing material is CaO, AlF 3 , Na 3 A
It is selected from simple compounds such as IF 6 , AlPO 4 , and SiO 2 and mixtures thereof. For example, a mixture of AlPO 4 and SiO 2 may be used.

【００２８】コーティング工程：目止め工程の後で行わ
れる工程であり、タービン翼本体の外側表面に断熱材層
を設ける。場合によって、消火栓冷却ガス口１４を除く
冷却ガス口１３にマスキングをする。最初に、タービン
翼本体１０の表面にボンドコートをコーティングする。
次に、ボンドコートの表面にインターフェースをコーテ
ィングする。最後に、インターフェースの表面にトップ
コートをコーティングする。Coating process: This process is carried out after the filling process, and a heat insulating material layer is provided on the outer surface of the turbine blade body. In some cases, masking is applied to the cooling gas port 13 excluding the hydrant cooling gas port 14. First, the surface of the turbine blade body 10 is coated with a bond coat.
Next, the interface is coated on the surface of the bond coat. Finally, the surface of the interface is coated with a top coat.

【００２９】目止め除去工程：前記コーティング工程の
後に行われる工程であり、目止め材を除去する。例え
ば、タービン翼を水に浸して、目止め材を溶かし出す。
場合によっては、本工程を実施せずに、目止め材をター
ビン翼本体に残しておいてもよい。Sealing removal step: This is a step performed after the coating step, in which the sealing material is removed. For example, the turbine blade is immersed in water to dissolve the sealing material.
In some cases, the sealing material may be left on the turbine blade body without performing this step.

【００３０】次に、本発明の実施例を使用して行う、サ
ーマルバリアコート剥離判断方法を説明する。 タービン準備工程：タービン翼を使用したタービンを用
意する。 冷却ガス消費量検知工程：タービン翼に供給する冷却ガ
スの消費量を検知する。サーマルバリアコートの剥離の
有無を判断工程：前記消費量の変動によって前記サーマ
ルバリアコートの剥離の有無を判断する。Next, a method for judging peeling of the thermal barrier coat, which is carried out by using the embodiment of the present invention, will be described. Turbine preparation step: Prepare a turbine using turbine blades. Cooling gas consumption detection process: The consumption of cooling gas supplied to the turbine blade is detected. Determining whether the thermal barrier coat is peeled off: Whether the thermal barrier coat is peeled off is determined based on the fluctuation of the consumption amount.

【００３１】上述の実施形態のタービン翼を用いれば、
トップコートが剥離しても、タービン翼の材料の温度が
上昇せず、所定の機械的性能を維持できる。また、トッ
プコートが剥離してもタービン翼の材料の温度が上昇す
る恐れがないので、タービンのガス温度を上げることが
でき、タービン効率が向上する。また、トップコートが
剥がれるまでは、冷却ガス口から吹きだす冷却ガスの消
費量がすくないので、タービンの全体機械効率が向上す
る。上述の実施形態のタービン翼の製造方法を用いれ
ば、従来のタービン翼の製造設備を使用して簡易に本発
明のタービン翼を製作できる。上述の実施形態のトップ
コート剥離検知方法を用いれば、タービンが運転中でも
トップコートの剥離を確認でき、適切なタービン運転を
行うことが出来る。また、トップコートが剥離した際
に、素早くタービン翼を交換することができる。Using the turbine blade of the above embodiment,
Even if the top coat is peeled off, the temperature of the material of the turbine blade does not rise, and the predetermined mechanical performance can be maintained. Further, even if the top coat is peeled off, the temperature of the material of the turbine blade does not rise, so that the gas temperature of the turbine can be raised and the turbine efficiency is improved. Further, until the top coat is peeled off, the consumption of the cooling gas blown out from the cooling gas port is small, so that the overall mechanical efficiency of the turbine is improved. By using the turbine blade manufacturing method of the above-described embodiment, the turbine blade of the present invention can be easily manufactured by using the conventional turbine blade manufacturing equipment. By using the topcoat peeling detection method of the above-described embodiment, the peeling of the topcoat can be confirmed even when the turbine is operating, and appropriate turbine operation can be performed. Further, when the top coat is peeled off, the turbine blade can be quickly replaced.

【００３２】本発明は以上に述べた実施形態に限られる
ものではなく、発明の要旨を逸脱しない範囲で各種の変
更が可能である。The present invention is not limited to the embodiments described above, and various modifications can be made without departing from the gist of the invention.

【００３３】[0033]

【発明の効果】以上説明したように本発明の外周面が翼
形状をしたタービン翼は、その構成により、以下の効果
を有する。通常運用時には、冷却ガスが内部空隙を経由
して冷却ガス口から吹きだしてタービン翼本体を冷却
し、サーマルバリアコートが剥離すると、剥離した箇所
の冷却ガス口からも冷却ガスが吹きだして、通常運用時
よりも多い冷却ガスでタービン翼を冷却し、タービン翼
の温度が上昇するのを防止できる。また、一部のサーマ
ルバリアコートが剥離すると、剥離した部分の冷却ガス
口の数が増加し、約２割を超えて増量した冷却ガスでタ
ービン翼を冷却し、タービン翼の温度が上場するのを防
止できる。また、詰まった目止め材により断熱材が冷却
ガス口を埋めることを防止して、断熱材層をタービン翼
本体の外側表面にコーティングでき、内部空隙と複数の
冷却ガス口とを有する翼形状断面の柱状構造体の外側表
面を、断熱材層であるサーマルバリアコートが覆い、前
記サーマルバリアコートが前記冷却ガス口の少なくとも
幾つかを塞いでいるタービン翼を製作できる。また、目
止め材を冷却ガス口の総数の２０％以上の数の冷却ガス
口に詰めるので、内部空隙と複数の冷却ガス口とを有す
る翼形状断面の柱状構造体の外側表面を、断熱材層であ
るサーマルバリアコートが覆い、前記サーマルバリアコ
ートが冷却ガス口の総数のうちの２０％を越える数の冷
却ガス口を塞いでいるタービン翼を製作できる。また、
前記サーマルバリアコートにより塞がれた冷却ガス口に
はなにも詰まっていないので、サーマルバリアコートが
剥離すると、剥離した箇所の冷却ガス口からもすぐに冷
却ガスが吹きだして、通常運用時よりも多い冷却ガスで
タービン翼を冷却し、タービン翼の温度が上昇するのを
防止できる様になったタービン翼を製作できる。また、
上述のタービン翼を使用したタービンを用意し、タービ
ン翼に供給する冷却ガスの消費量を検知し、前記消費量
の変動によって前記サーマルバリアコートの剥離の有無
を判断するので、タービンの運転中にサーマルバリアコ
ートの剥離を知ることができ、例えばタービンの異常な
温度上昇を防止する手だてを直ちにとることが出来る。
従って、高温ガスの温度を上げて運用できるタービン翼
とその製造方法とサーマルバリアコートの剥離を判断す
る方法を提供できる。As described above, the turbine blade of which the outer peripheral surface has the blade shape according to the present invention has the following effects due to its configuration. During normal operation, cooling gas is blown out from the cooling gas port through the internal gap to cool the turbine blade body, and when the thermal barrier coat is peeled off, cooling gas is blown out from the cooling gas port at the peeled point as well. It is possible to prevent the temperature of the turbine blade from rising by cooling the turbine blade with more cooling gas than is necessary. Further, when a part of the thermal barrier coat is peeled off, the number of cooling gas ports in the peeled part is increased, the turbine blade is cooled by the cooling gas increased by more than about 20%, and the temperature of the turbine blade is listed. Can be prevented. In addition, the plugged sealing material prevents the heat insulating material from filling the cooling gas port, and the heat insulating material layer can be coated on the outer surface of the turbine blade body, and has a blade shape cross section having an internal void and a plurality of cooling gas ports. It is possible to manufacture a turbine blade in which the outer surface of the columnar structure is covered with a thermal barrier coat which is a heat insulating material layer, and the thermal barrier coat covers at least some of the cooling gas ports. Further, since the sealing material is filled in the cooling gas ports of 20% or more of the total number of the cooling gas ports, the outer surface of the columnar structure having the blade-shaped cross section having the internal voids and the plurality of cooling gas ports is provided with the heat insulating material. It is possible to manufacture a turbine blade that is covered with a layer of a thermal barrier coat, and the thermal barrier coat blocks more than 20% of the total number of cooling gas ports. Also,
Since the cooling gas port blocked by the thermal barrier coat is not clogged with anything, when the thermal barrier coat is peeled off, the cooling gas immediately blows out from the cooling gas port at the peeled point, which is more than normal operation. The turbine blade can be manufactured by cooling the turbine blade with a large amount of cooling gas and preventing the temperature of the turbine blade from rising. Also,
Prepare a turbine using the turbine blades described above, detect the consumption of cooling gas to be supplied to the turbine blades, and determine whether or not the thermal barrier coat is peeled off by the fluctuation of the consumption. It is possible to know the peeling of the thermal barrier coat and immediately take measures to prevent an abnormal temperature rise of the turbine, for example.
Therefore, it is possible to provide a turbine blade that can be operated by raising the temperature of the high temperature gas, a method for manufacturing the turbine blade, and a method for determining separation of the thermal barrier coat.

【００３４】[0034]

【図面の簡単な説明】[Brief description of drawings]

【図１】本発明の実施形態の構造図である。FIG. 1 is a structural diagram of an embodiment of the present invention.

【図２】本発明の実施形態の断面図である。FIG. 2 is a cross-sectional view of an embodiment of the present invention.

【図３】本発明の実施形態の断面詳細図である。FIG. 3 is a detailed sectional view of an embodiment of the present invention.

【図４】従来のタービン翼の構造図である。FIG. 4 is a structural diagram of a conventional turbine blade.

【図５】従来のタービン翼の断面図である。FIG. 5 is a cross-sectional view of a conventional turbine blade.

【図６】従来のタービン翼の断面詳細図である。FIG. 6 is a detailed sectional view of a conventional turbine blade.

【符号の説明】[Explanation of symbols]

１ タービン翼部材 ２ タービン翼 ３ タービン翼部材基部 ４ タービン翼取り付け部 ５ 冷却ガス供給口 ６ 冷却ガス排出口 １０ タービン翼本体 １１ 柱状構造体 １２ 内部空隙 １３ 冷却ガス口 １４ 消火栓冷却ガス口（冷却ガス口） ２０ サーマルバリアコート ２１ トップコート ２２ インターフェース ２３ ボンドコート 1 Turbine blade member 2 turbine blades 3 Turbine blade member base 4 Turbine blade mounting part 5 Cooling gas supply port 6 Cooling gas outlet 10 Turbine blade body 11 Columnar structure 12 Internal void 13 Cooling gas port 14 Fire hydrant cooling gas port (cooling gas port) 20 thermal barrier coat 21 top coat 22 Interface 23 bond coat

Claims (6)

【特許請求の範囲】[Claims] 【請求項１】外周面が翼形状をしたタービン翼であっ
て、冷却ガスが通過する様に内部に設けられた内部空隙
と該内部空隙に連通可能にし外周面に設けられた複数の
冷却ガス口とを有する翼形状断面のタービン翼本体と、
前記タービン翼本体の外周面を覆う断熱材層であるサー
マルバリアコートとを備え、前記サーマルバリアコート
が前記冷却ガス口の少なくとも幾つかを塞いでいること
を特徴とするタービン翼。1. A turbine blade having an outer peripheral surface in the shape of a blade, wherein a plurality of cooling gases are provided on the outer peripheral surface so as to allow the cooling gas to pass therethrough, and an internal space provided inside. A turbine blade body having a blade-shaped cross section having a mouth,
A turbine blade, comprising: a thermal barrier coat, which is a heat insulating material layer that covers an outer peripheral surface of the turbine blade body, and the thermal barrier coat blocks at least some of the cooling gas ports. 【請求項２】前記サーマルバリアコートが冷却ガス口の
総数のうちの２０％を越える数の冷却ガス口を塞いでい
ることを特徴とする請求項１に記載のタービン翼。2. The turbine blade according to claim 1, wherein the thermal barrier coat closes more than 20% of the total number of cooling gas ports. 【請求項３】外周面が翼形状をしたタービン翼の製造方
法であって、冷却ガスが通過する様に内部に設けられた
内部空隙と該内部空隙に連通可能にし外周面に設けられ
た複数の冷却ガス口とを有する翼形状断面のタービン翼
本体を用意するタービン翼本体準備工程と、前記冷却ガ
ス口のすくなくとも幾つかに目止め材を詰める目止め工
程と、目止め工程の後でタービン翼本体の外側表面に断
熱材層を設けるコーティング工程と、を備えたことを特
徴とするタービン翼の製造方法3. A method of manufacturing a turbine blade, the outer peripheral surface of which is blade-shaped, wherein a plurality of internal voids are provided inside of the inner wall to allow cooling gas to pass through and a plurality of inner voids are provided so as to communicate with the inner voids. A turbine blade body preparation step of preparing a turbine blade body having a blade-shaped cross section having a cooling gas port, a sealing step of filling at least some of the cooling gas ports with a sealing material, and a turbine after the sealing step And a coating step of providing a heat insulating material layer on the outer surface of the blade body. 【請求項４】前記目止め工程が、前記冷却ガス口の総数
のうちの２０％を越える数の冷却ガス口に目止め材を詰
めることを特徴とする請求項３に記載のタービン翼の製
造方法4. The turbine blade manufacturing method according to claim 3, wherein in the plugging step, the plugging material is filled in the cooling gas ports of which the number exceeds 20% of the total number of the cooling gas ports. Method 【請求項５】前記コーティング工程の後に目止め材を除
去する目止め除去工程を備えることを特徴とする請求項
３または請求項４のうちの一つに記載のタービン翼の製
造方法。5. The method for manufacturing a turbine blade according to claim 3, further comprising a sealing removal step of removing a sealing material after the coating step. 【請求項６】請求項１乃至請求項２のタービン翼を使用
したタービンを用意し、タービン翼に供給する冷却ガス
の消費量を検知し、前記消費量の変動によって前記サー
マルバリアコートの剥離の有ることを判断することを特
徴とするサーマルバリアコート剥離判断方法。6. A turbine using the turbine blade according to any one of claims 1 and 2 is prepared, the consumption of cooling gas supplied to the turbine blade is detected, and the thermal barrier coat is peeled off by the fluctuation of the consumption. A thermal barrier coat peeling judgment method, characterized by judging that there is.
JP2001374542A 2001-12-07 2001-12-07 Turbine blade, its production method, and its thermal barrier coat separation determining method Pending JP2003172102A (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
JP2001374542A JP2003172102A (en) 2001-12-07 2001-12-07 Turbine blade, its production method, and its thermal barrier coat separation determining method
CA002385932A CA2385932C (en) 2001-12-07 2002-05-08 Turbine blade, manufacturing method of turbine blade, and strip judging method of thermal barrier coat
US10/146,964 US20030108424A1 (en) 2001-12-07 2002-05-17 Turbine blade, manufacturing method of turbine blade, and strip judging method of thermal barrier coat
EP02012119A EP1318273B1 (en) 2001-12-07 2002-05-31 Turbine blade, manufacturing method of turbine blade, and strip judging method of a thermal barrier coat
DE60216405T DE60216405T2 (en) 2001-12-07 2002-05-31 Turbine blade, turbine blade manufacturing method, and evaluation method for peeling a thermal barrier coating

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
JP2001374542A JP2003172102A (en) 2001-12-07 2001-12-07 Turbine blade, its production method, and its thermal barrier coat separation determining method

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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
KR101828543B1 (en) * 2013-07-23 2018-02-12 한화테크윈 주식회사 Turbine blade, turbine comprising the same and method for manufacturing the turbine blade
WO2019164683A3 (en) * 2018-02-08 2019-11-28 General Electric Company Method of masking apertures in a component and processing the component

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US20090074576A1 (en) * 2006-04-20 2009-03-19 Florida Turbine Technologies, Inc. Turbine blade with cooling breakout passages
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JP2006138624A (en) * 2004-11-09 2006-06-01 General Electric Co <Ge> Gas turbine engine component
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
KR101828543B1 (en) * 2013-07-23 2018-02-12 한화테크윈 주식회사 Turbine blade, turbine comprising the same and method for manufacturing the turbine blade
WO2019164683A3 (en) * 2018-02-08 2019-11-28 General Electric Company Method of masking apertures in a component and processing the component
US11814742B2 (en) 2018-02-08 2023-11-14 General Electric Company Method of masking apertures in a component and processing the component

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CA2385932A1 (en) 2003-06-07

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