JP6054137B2 - High temperature member for gas turbine having thermal barrier coating - Google Patents

High temperature member for gas turbine having thermal barrier coating Download PDF

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JP6054137B2
JP6054137B2 JP2012234278A JP2012234278A JP6054137B2 JP 6054137 B2 JP6054137 B2 JP 6054137B2 JP 2012234278 A JP2012234278 A JP 2012234278A JP 2012234278 A JP2012234278 A JP 2012234278A JP 6054137 B2 JP6054137 B2 JP 6054137B2
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alloy
gas turbine
temperature member
temperature
thermal barrier
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JP2014084791A (en
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秀行 有川
秀行 有川
児島　慶享
慶享 児島
忠 粕谷
忠 粕谷
輝 目幡
輝 目幡
市川　国弘
国弘 市川
宏之 遠藤
宏之 遠藤
遠藤　孝夫
孝夫 遠藤
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Mitsubishi Power Ltd
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Mitsubishi Hitachi Power Systems Ltd
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Priority to JP2012234278A priority Critical patent/JP6054137B2/en
Priority to EP13189647.4A priority patent/EP2725120B1/en
Priority to US14/061,306 priority patent/US20140112758A1/en
Priority to CN201310508420.6A priority patent/CN103774134A/en
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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
    • 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
    • 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
    • C23C24/00—Coating starting from inorganic powder
    • C23C24/02—Coating starting from inorganic powder by application of pressure only
    • C23C24/04—Impact or kinetic deposition of particles
    • 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
    • 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
    • C23C4/04—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the coating material
    • C23C4/06—Metallic material
    • C23C4/073—Metallic material containing MCrAl or MCrAlY alloys, where M is nickel, cobalt or iron, with or without non-metal elements
    • 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
    • C23C4/04—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the coating material
    • C23C4/06—Metallic material
    • C23C4/08—Metallic material containing only metal elements
    • 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
    • F01D5/183—Blade walls being porous
    • 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
    • F01D5/186—Film cooling

Description

本発明は、耐熱性に優れた遮熱コーティングを有する、ガスタービンの動静翼，燃焼器，シュラウド等のガスタービン用高温部材に関する。 The present invention relates to a high-temperature member for a gas turbine such as a moving and stationary blade, a combustor, and a shroud of a gas turbine, which has a thermal barrier coating excellent in heat resistance.

ガスタービンは効率向上を目的として運転温度が年々高くなってきている。運転温度の高温化に対処するために、ガスタービン高温部品には、耐熱性に優れた材料が用いられ、加えて、燃焼ガスに晒される面の反対面を、空気や蒸気等の流体冷媒で冷却する構造が採用されている。さらに、温度環境を和らげる目的で、表面に低熱伝導性のセラミックスよりなる遮熱コーティング（Thermal Barrier Coating：以下ＴＢＣと称す）を施すことが行われている。使用条件にもよるが、一般的にＴＢＣの適用により、基材温度が５０〜１００℃低減できる。例えば、特開昭６２−２１１３８７号公報（特許文献１）などには、基材に対して、ＭＣｒＡｌＹ合金層を介して、低熱伝導性で耐熱性に優れた部分安定化ジルコニアよりなる遮熱層を有するＴＢＣが開示されている。ここで、Ｍは鉄（Ｆｅ），Ｎｉ及びＣｏからなるグループから選ばれた少なくとも１種を表し、Ｃｒはクロム、Ａｌはアルミニウム、Ｙはイットリウムを表す。   The operation temperature of gas turbines is increasing year by year for the purpose of improving efficiency. In order to cope with higher operating temperatures, high-temperature materials are used for high-temperature parts of the gas turbine. In addition, the surface exposed to the combustion gas is covered with a fluid refrigerant such as air or steam. A cooling structure is adopted. Furthermore, for the purpose of relaxing the temperature environment, a thermal barrier coating (Thermal Barrier Coating: hereinafter referred to as TBC) made of ceramic with low thermal conductivity is performed on the surface. Although it depends on the use conditions, the substrate temperature can generally be reduced by 50 to 100 ° C. by application of TBC. For example, JP-A-62-211387 (Patent Document 1) and the like disclose a thermal barrier layer made of partially stabilized zirconia having low thermal conductivity and excellent heat resistance with respect to a base material via an MCrAlY alloy layer. A TBC is disclosed. Here, M represents at least one selected from the group consisting of iron (Fe), Ni, and Co, Cr represents chromium, Al represents aluminum, and Y represents yttrium.

このような、ＴＢＣと冷却構造を有するガスタービン高温部品は、優れた耐熱性を示すが、さらなるガスタービンの性能向上に向け、より冷却効率の高い浸み出し冷却方式の採用が望まれている。浸み出し冷却は、部材の表面全体から、微細流路（一般的には多孔体）を通じて、微量の冷却媒体を均一に高温部材の表面から浸み出させることで、効率良く冷却を行う方法である。例えば、特開平１０−２３１７０４号公報（特許文献２）、特開２０１０−６５６３４号公報（特許文献３）には多孔質金属上に多孔質セラミック層による浸み出し冷却を採用したガスタービン高温部材が開示されている。また、特開２００５−３５０３４１号公報（特許文献４）には、多孔質セラミックと耐熱合金基材を鋳造時に一体化した構造で、浸み出し冷却構造を採用したガスタービン高温部材が開示されている。 Such a gas turbine high-temperature component having a TBC and a cooling structure exhibits excellent heat resistance, but in order to further improve the performance of the gas turbine, it is desired to employ a leaching cooling method with higher cooling efficiency. . The leaching cooling is a method for efficiently cooling a small amount of cooling medium from the entire surface of the member through the fine flow path (generally a porous body) uniformly from the surface of the high temperature member. It is. For example, JP-A 10 -231704 (Patent Document 2), JP 2010-65634 (Patent Document 3) to the gas turbine hot member employing the out soak by the porous ceramic layer on a porous metal cooling Is disclosed. Japanese Patent Application Laid-Open No. 2005-350341 (Patent Document 4) discloses a gas turbine high-temperature member that employs a leaching cooling structure in which a porous ceramic and a heat-resistant alloy base material are integrated during casting. Yes.

特開昭６２−２１１３８７号公報JP-A-62-211387 特開平１０−２３１７０４号公報Japanese Patent Laid-Open No. 10-231704 特開２０１０−６５６３４号公報JP 2010-65634 A 特開２００５−３５０３４１号公報JP 2005-350341 A

上記従来技術では、一部でＴＢＣの遮熱セラミック層が採用されているものの、いずれの公知例においても、ＴＢＣの合金下地層に相当する層が、皮膜ではなく多孔質金属で代替されているか、あるいは、合金下地層に相当する層が省略されている。これは、冷却媒体の流路となる微細通路を従来の合金下地層の成膜方法を用いて形成することが困難なためである。ＴＢＣでは、遮熱セラミック層は、燃焼ガスからの熱を遮蔽する役割を担っており、みかけの熱伝導を低く抑え、さらに熱応力を緩和する効果を期待できることから、多孔質セラミック層が採用されている。一方、合金下地層は、セラミック層と基材の密着を確保すると共に、燃焼ガスによる酸化や腐食から基材を保護する役割を担っており、より緻密な組織が採用されている。このため、ＴＢＣと浸み出し冷却を組み合わせた高温部材を実現するためには、従来とは異なる、冷却媒体の流路を有した合金下地層が必要である。
そこで、本発明の目的は、浸み出し冷却に適した微細な冷却媒体の流路を有した合金下地層を実現し、これを用いた浸み出し冷却機能と遮熱コーティングを備え、耐熱性に優れるガスタービン用高温部材を提供することにある。 In the above prior art, a TBC thermal barrier ceramic layer is used in part, but in any known example, the layer corresponding to the TBC alloy underlayer is replaced with a porous metal instead of a film. Alternatively, a layer corresponding to the alloy underlayer is omitted. This is because it is difficult to form a fine passage serving as a cooling medium flow path by using a conventional alloy underlayer film forming method. In TBC, the thermal barrier ceramic layer plays the role of shielding the heat from the combustion gas, and it can be expected to have the effect of suppressing the apparent thermal conduction and further reducing the thermal stress. Therefore, the porous ceramic layer is adopted. ing. On the other hand, the alloy underlayer ensures the adhesion between the ceramic layer and the base material and plays a role of protecting the base material from oxidation and corrosion due to combustion gas, and has a finer structure. Therefore, in order to realize a high-temperature member combining TBC and leaching cooling, an alloy underlayer having a cooling medium flow path different from the conventional one is required.
Accordingly, an object of the present invention is to realize an alloy underlayer having a flow path of a fine cooling medium suitable for leaching cooling, and is provided with a leaching cooling function and a thermal barrier coating using the alloy underlayer. It is in providing the high temperature member for gas turbines which is excellent in.

本発明は、上記課題を鑑み、高温の燃焼ガスに曝される基材表面に、合金下地層を設け、さらに、その表面上に遮熱セラミック層を設けてなる遮熱コーティングを有し、かつ、合金下地層は、合金粉末粒子が概略球形を維持した状態で積層されたものであり、流体冷媒による冷却構造を有するガスタービン用高温部材において、合金下地層、及び、遮熱セラミック層に、基材側から表面側に連通した微細通路を設け、微細通路を合金下地層の概略球形が維持された合金粉末粒子間の間隙により形成されたものとし、部材を冷却する冷媒の一部を、これら微細通路を通じて、部材外部に流出させることを最も主要な特徴とする。 In view of the above problems, the present invention has a thermal barrier coating in which an alloy base layer is provided on the surface of a substrate exposed to a high-temperature combustion gas, and further, a thermal barrier ceramic layer is provided on the surface, and The alloy underlayer is formed by laminating the alloy powder particles in a state of maintaining a substantially spherical shape.In the high temperature member for a gas turbine having a cooling structure with a fluid refrigerant, the alloy underlayer and the heat-shielding ceramic layer are A fine passage communicating from the substrate side to the surface side is provided, the fine passage is formed by a gap between alloy powder particles in which the approximately spherical shape of the alloy underlayer is maintained, and a part of the refrigerant for cooling the member is The most important feature is that it flows out of the member through these fine passages.

本発明は、ＴＢＣの合金下地層と遮熱セラミック層内に、基材側から表面側に連通した微細通路を設け、部材を冷却する冷媒の一部を、これら微細通路を通じて部材外部に流出させることで、ＴＢＣ、特に合金下地層が効率良く冷却される。また、高温部材の表面全面から均一に冷媒が浸み出すことによって、均一で効率的なフィルム冷却効果が期待できる。これらの効果により、燃焼ガス温度の高温化に伴う部材温度上昇によって、従来技術の適用が困難となるような過酷な条件下でも、使用が可能となるという利点がある。また、本発明の遮熱コーティングと冷却構造を有するガスタービン用高温部材を用いたガスタービンは、より高温で運転が可能で、効率を高めることができるという利点がある。   In the present invention, a fine passage communicating from the base material side to the surface side is provided in the TBC alloy underlayer and the heat-shielding ceramic layer, and a part of the refrigerant for cooling the member flows out of the member through the fine passage. As a result, the TBC, particularly the alloy underlayer, is efficiently cooled. Moreover, a uniform and efficient film cooling effect can be expected by the refrigerant leaching uniformly from the entire surface of the high temperature member. Due to these effects, there is an advantage that it can be used even under severe conditions in which application of the prior art becomes difficult due to an increase in the temperature of the member accompanying an increase in the combustion gas temperature. Moreover, the gas turbine using the high-temperature member for gas turbines having the thermal barrier coating and the cooling structure of the present invention has an advantage that it can be operated at a higher temperature and efficiency can be increased.

本発明の遮熱コーティングと冷却構造を有するガスタービン用高温部材の構造を示す断面模式図である。It is a cross-sectional schematic diagram which shows the structure of the high temperature member for gas turbines which has the thermal barrier coating of this invention, and a cooling structure. ガスタービンの構造を示す断面模式図である。It is a cross-sectional schematic diagram which shows the structure of a gas turbine.

以下、図面を用いて本発明を詳細に説明する。   Hereinafter, the present invention will be described in detail with reference to the drawings.

本発明は、図１に示すように、基材１上に合金下地層２を設け、さらに、その上に遮熱セラミック層３を設けた構成である。基材１には、基材１の冷却媒体通路から、合金下地層２を設けた表面に向けて、基材１を貫通した冷却孔４が複数設けられている。合金下地層２は、多数の概略球形の合金粉末粒子５が積層され、基材１側からコーティング表面まで連通した粒子間の間隙６が存在する構造を有することを特徴とする。さらに、合金下地層２の上には、遮熱セラミック層３が設けられ、遮熱セラミック層３は、多数の縦方向クラック７を有する。基材１の冷却媒体通路から、冷却孔４を通じて、合金下地層２に達した流体冷媒８は、合金下地層内の粒子間の間隙６を通じて、合金下地層２内を拡散しながら表面側に流れ、遮熱セラミック層３に達し、遮熱セラミック層内の縦方向クラック７を通じて、遮熱セラミック層３の表面から流出する。
基材１は、ニッケル基，コバルト基、または、鉄基の耐熱合金を用いることができる。合金下地層２は、ニッケル基，コバルト基、または、鉄基の耐熱合金を用いることができるが、好ましくは、ＭＣｒＡｌＹ（Ｍは、Ｆｅ，Ｎｉ，Ｃｏのうちの何れか、または複数）合金を用いることが望ましい。ＭＣｒＡｌＹ合金は、耐酸化性に優れるため、好適である。
また、合金下地層２は、多数の概略球形の合金粒子５が積層され、基材１側から合金下地層２の表面まで、連通した粒子間の間隙６が存在する構造を有する。このような構造の皮膜を形成するためには、例えば、ガスアトマイズ法で製造した概略球形の合金粉末を原料として用い、合金粉末を基材表面に高速で衝突させて積層する方法を用いることが好ましい。具体的には、例えば、プラズマ溶射法、高速ガス溶射（ＨＶＯＦ）法、コールドスプレー法等の方法を用いることができる。中でも、コールドスプレー法が最も好適に用いられる。
本発明の特徴である、基材１側からコーティング表面まで連通した粒子間の間隙６が存在する構造を有する合金下地層２を形成するには、アーク溶射や火炎（フレーム）溶射のように、合金粉末粒子を高温で溶融して基材に衝突させて積層する方法では、溶融した粉末粒子が基材に衝突した際に大きく扁平して積層するため、連通しない気孔（いわゆる閉気孔）が形成されやすくなる。また、大気中で溶融する温度まで加熱された合金粉末では、表面に酸化物が生じ、この酸化物が皮膜内に混入することで皮膜の耐酸化性を低下させる。また、粒子同士の結合が酸化物によって妨げられ、皮膜の強度が低下するという問題も生じる。
従って、本発明の遮熱コーティングの合金下地層２を形成する際には、原料として用いる概略球形の合金粉末を溶融、酸化させず、そのままの球形に近い形状を維持したまま積層することが望ましい。これには、より低温で成膜を行うことできるコールドスプレー法が好適である。しかし、低温でも粒子速度が高速になり過ぎると、基材に衝突した際に粉末粒子の扁平が生じ、皮膜が緻密化して連通気孔が減少するため、本発明の合金下地層２を形成できなくなるため、成膜条件を適当に調整する必要がある。なお、同様に成膜条件を、適宜、調整することで、プラズマ溶射法、高速ガス溶射（ＨＶＯＦ）法等を用いることもできる。
前記の成膜法を用いて形成した、本発明の基材側から表面側に連通した粒子間の間隙６を有する合金下地層２の連通間隙は、皮膜内体積分率が３０〜７０％の範囲が好ましい。間隙の体積分率が３０％未満では、流通する冷却媒体量が少なくなり、浸み出し冷却の効果が十分に得られない。一方、間隙の体積分率が増加すると冷却効果は高まるが、皮膜強度が低下し、間隙の体積分率が７０％を超えると、使用中にコーティングの損傷が生じ易くなってしまう。より好ましくは、間隙の体積分率が４０〜６０％の範囲が好適である。
また、本発明の遮熱コーティングは、合金下地層２、遮熱セラミック層３のいずれも、成膜後に熱処理を施すことが好ましい。合金下地層２では、熱処理による固相拡散によって粒子間の結合を強化することで皮膜強度を向上することができる。また、遮熱セラミック層３では、縦方向クラックの開口を促し、冷却媒体の流通を円滑にすることが期待できる。熱処理方法は、合金下地層２の酸化を防止するため、真空中で行うことが望ましい。熱処理条件は、コーティング、基材材料にも依るが、概ね、１０００℃以上で２ｈ以上保持することが好ましい。
以下、実施例を説明する。
（実施例１）
基体として、ニッケル基耐熱合金ＩＮ７３８（１６％Cr−８．５％Ｃｏ−３．４％Ｔｉ−３．４％Ａｌ−２．６％Ｗ−１．７％Ｍｏ−１．７％Ｔａ−０．９％Ｎｂ−０．１％Ｃ−０．０５％Ｚｒ−０．０１％Ｂ−残部Ｎｉ、重量％）製で、内部に冷却空気通路を有する、ガスタービン１段動翼を準備した。動翼には、放電加工によって、基体表面から内部冷却通路まで貫通した冷却孔を複数加工した。また、原料粉末として、ガスアトマイズ法で製造された、概略球状で平均粒径約４０μｍのＣｏＮｉＣｒＡｌＹ合金粉末（Ｃｏ−３２％Ｎｉ−２１％Ｃｒ−８％Ａｌ−０．５％Ｙ、重量％）を準備した。コールドスプレー装置を用い、原料粉末を動翼の燃焼ガス通路面に対し、成膜した。成膜条件は、作動ガスに窒素ガスを用い、ガス圧力３ＭＰａ、ガス温度８００℃、粉末供給量２０ｇ／ｍｉｎ、成膜距離１５ｍｍの条件を用い、合金下地層２の厚さが約０．３ｍｍまで成膜を実施した。
その後、合金下地層２を設けた基材１上に、イットリア部分安定化ジルコニア（ＺｒＯ2−８ｗｔ％Ｙ2Ｏ3）粉末を用い、大気中プラズマ溶射（プラズマ出力約１００ｋＷ）にて約０.６ｍｍの厚さ、気孔率が約８％の縦クラックを有する遮熱セラミック層３を設けた。この際の成膜条件としては、予熱温度が約８００℃、溶射ガンの移動速度は３０ｍ／ｍｉｎ、溶射距離は９０ｍｍとし、熱流速約０．４ＭＷ／ｍ2とした。さらに、遮熱セラミック層３を成膜後の動翼に対し、真空中で１１２０℃×２ｈ、８４０℃×２４ｈの熱処理を実施した。
このようにして製作した動翼を切断して断面組織を確認したところ、図１に示したように、合金下地層２は、多数の概略球形の合金粒子５が積層され、基材１側から合金下地層２の表面まで連通した粒子間の間隙６が存在する組織を呈していた。相対密度から気孔の体積分率を測定したところ、約５０％であった。
前記手順で作製した別試験翼をガスタービンに組込み、1年間の試験運転を行った。この際、翼の冷却空気入口にオリフィスを設け、従来設計よりも冷却空気量を３０％減じた。
試験運転後の、本発明のＴＢＣを用いた動翼は、外観、切断調査のいずれにおいても、損傷はほとんど認められなかった。一方、比較のため、同時に冷却空気量を減じて運転に供した従来技術のＴＢＣを設けた動翼では、外観上、ＴＢＣの剥離が部分的に認められ、さらに断面調査では、剥離部以外に合金下地層の酸化損傷が認められた。これらの結果から、本発明のＴＢＣを設けたガスタービン高温部品が優れた耐熱性を有することが確認された。

（実施例２）
図２は発電用ガスタービン主要部の断面模式図である。ガスタービンは、タービンケーシング４８の内部に、中心に回転軸（ロータ）４９と、回転軸４９の周囲に設置される動翼４６とケーシング４８側に支持される静翼４５、タービンシュラウド４７を有するタービン部４４を備える。このタービン部４４に連結され、大気を吸込み、燃焼用及び冷却媒体用の圧縮空気を得る圧縮機５０と、燃焼器４０を有する。燃焼器４０は、圧縮機５０から供給される圧縮空気と、供給される燃料（図示せず）を混合して噴射する燃焼器ノズル４１を有し、この混合気を燃焼器ライナ４２内で燃焼させて高温高圧の燃焼ガスを発生し、トランジションピース（尾筒）４３を介して、この燃焼ガスがタービン４４に供給されることで、ロータ４９が高速で回転する。圧縮機５０より吐出された圧縮空気の一部は、燃焼器４０のライナ４２，トランジションピース４３やタービン静翼４５，タービン動翼４６の内部冷却空気として用いられる。燃焼器４０で発生した高温高圧の燃焼ガスは、トランジションピース４３を経てタービン静翼４５で整流され、動翼４６に噴射されてタービン部４４を回転駆動する。そして図示はしていないが、一般的には回転軸４９の端部に結合されている発電機により発電するように構成されている。 As shown in FIG. 1, the present invention has a configuration in which an alloy base layer 2 is provided on a base material 1, and a thermal barrier ceramic layer 3 is further provided thereon. The base material 1 is provided with a plurality of cooling holes 4 penetrating the base material 1 from the cooling medium passage of the base material 1 toward the surface on which the alloy base layer 2 is provided. The alloy underlayer 2 has a structure in which a large number of roughly spherical alloy powder particles 5 are laminated, and there is a structure in which there are gaps 6 between particles communicating from the substrate 1 side to the coating surface. Furthermore, a heat insulating ceramic layer 3 is provided on the alloy base layer 2, and the heat insulating ceramic layer 3 has a number of longitudinal cracks 7. The fluid refrigerant 8 that has reached the alloy underlayer 2 from the cooling medium passage of the base material 1 through the cooling hole 4 is diffused in the alloy underlayer 2 through the gaps 6 between the particles in the alloy underlayer to the surface side. It flows, reaches the thermal barrier ceramic layer 3, and flows out from the surface of the thermal barrier ceramic layer 3 through the longitudinal crack 7 in the thermal barrier ceramic layer.
The substrate 1 can be made of a nickel-based, cobalt-based, or iron-based heat-resistant alloy. The alloy base layer 2 may be a nickel-based, cobalt-based, or iron-based heat-resistant alloy, but preferably an MCrAlY (M is any one of Fe, Ni, Co, or a plurality) alloy. It is desirable to use it. The MCrAlY alloy is preferable because it is excellent in oxidation resistance.
The alloy underlayer 2 has a structure in which a large number of approximately spherical alloy particles 5 are laminated, and there are gaps 6 between the communicating particles from the substrate 1 side to the surface of the alloy underlayer 2. In order to form a film having such a structure, for example, it is preferable to use a substantially spherical alloy powder produced by a gas atomization method as a raw material, and use a method in which the alloy powder collides with the substrate surface at high speed and is laminated. . Specifically, for example, a plasma spraying method, a high-speed gas spraying (HVOF) method, a cold spray method, or the like can be used. Of these, the cold spray method is most preferably used.
In order to form the alloy underlayer 2 having a structure in which the gap 6 between the particles communicating from the substrate 1 side to the coating surface, which is a feature of the present invention, exists, like arc spraying or flame (frame) spraying, In the method in which alloy powder particles are melted at a high temperature and collide with the base material, the molten powder particles are flattened and laminated when they collide with the base material, so that pores that do not communicate (so-called closed pores) are formed. It becomes easy to be done. Moreover, in the alloy powder heated to the temperature which melt | dissolves in air | atmosphere, an oxide arises on the surface and this oxide mixes in a membrane | film | coat, and reduces the oxidation resistance of a membrane | film | coat. In addition, the bonding between the particles is hindered by the oxide, resulting in a problem that the strength of the film is lowered.
Therefore, when forming the alloy underlayer 2 of the thermal barrier coating of the present invention, it is desirable to laminate the substantially spherical alloy powder used as a raw material while maintaining the shape close to the spherical shape without melting and oxidizing it. . For this, a cold spray method that can form a film at a lower temperature is suitable. However, if the particle velocity becomes too high even at a low temperature, the powder particles become flat when they collide with the substrate, and the coating becomes dense and the continuous air holes are reduced, so that the alloy underlayer 2 of the present invention cannot be formed. Therefore, it is necessary to appropriately adjust the film forming conditions. Similarly, a plasma spraying method, a high-speed gas spraying (HVOF) method, or the like can be used by appropriately adjusting the film forming conditions.
The communicating gap of the alloy underlayer 2 having the gap 6 between the particles communicating from the base material side to the surface side of the present invention formed using the film forming method described above has an in-film volume fraction of 30 to 70%. A range is preferred. When the volume fraction of the gap is less than 30%, the amount of the circulating cooling medium decreases, and the leaching cooling effect cannot be sufficiently obtained. On the other hand, when the volume fraction of the gap increases, the cooling effect increases, but the film strength decreases, and when the volume fraction of the gap exceeds 70%, the coating is liable to be damaged during use. More preferably, the volume fraction of the gap is in the range of 40 to 60%.
In the thermal barrier coating of the present invention, it is preferable that both the alloy underlayer 2 and the thermal barrier ceramic layer 3 are subjected to heat treatment after film formation. In the alloy underlayer 2, the strength of the film can be improved by strengthening the bond between particles by solid phase diffusion by heat treatment. Moreover, in the thermal-insulation ceramic layer 3, it can be expected that the opening of the longitudinal crack is promoted and the circulation of the cooling medium is made smooth. The heat treatment method is desirably performed in a vacuum in order to prevent oxidation of the alloy underlayer 2. Although the heat treatment conditions depend on the coating and the base material, it is generally preferable to hold at 1000 ° C. or higher for 2 hours or longer.
Examples will be described below.
Example 1
Nickel-based heat-resistant alloy IN738 (16% Cr-8.5% Co-3.4% Ti-3.4% Al-2.6% W-1.7% Mo-1.7% Ta-0 9% Nb-0.1% C-0.05% Zr-0.01% B-balance Ni, wt%), and a gas turbine first stage blade having a cooling air passage therein was prepared. In the rotor blade, a plurality of cooling holes penetrating from the surface of the base to the internal cooling passage were machined by electric discharge machining. Further, as a raw material powder, a CoNiCrAlY alloy powder (Co-32% Ni-21% Cr-8% Al-0.5% Y, weight%) produced by a gas atomization method and having an approximately spherical average particle diameter of about 40 μm is used. Got ready. Using a cold spray apparatus, the raw material powder was deposited on the combustion gas passage surface of the rotor blade. The film formation conditions are as follows. Nitrogen gas is used as the working gas, the gas pressure is 3 MPa, the gas temperature is 800 ° C., the powder supply amount is 20 g / min, and the film formation distance is 15 mm. Film formation was carried out.
After that, yttria partially stabilized zirconia (ZrO 2 -8 wt% Y 2 O 3 ) powder is used on the base material 1 provided with the alloy underlayer 2 and is about 0.00 by atmospheric plasma spraying (plasma output about 100 kW). A thermal barrier ceramic layer 3 having a longitudinal crack with a thickness of 6 mm and a porosity of about 8% was provided. Film formation conditions at this time were a preheating temperature of about 800 ° C., a spray gun moving speed of 30 m / min, a spraying distance of 90 mm, and a heat flow rate of about 0.4 MW / m 2 . Further, the rotor blades after the formation of the thermal barrier ceramic layer 3 were heat-treated in vacuum at 1120 ° C. × 2 h and 840 ° C. × 24 h.
When the rotor blade manufactured in this way was cut and the cross-sectional structure was confirmed, as shown in FIG. A structure in which gaps 6 between the particles communicating to the surface of the alloy underlayer 2 existed was exhibited. When the volume fraction of the pores was measured from the relative density, it was about 50%.
Another test blade produced in the above procedure was incorporated into a gas turbine and a one-year test operation was performed. At this time, an orifice was provided at the cooling air inlet of the blade, and the amount of cooling air was reduced by 30% compared to the conventional design.
After the test operation, the rotor blades using the TBC of the present invention showed almost no damage in both appearance and cutting investigation. On the other hand, for comparison, in the rotor blade provided with the TBC of the prior art that was used for operation while reducing the cooling air amount at the same time, TBC peeling was partially recognized in appearance, and in the cross-sectional investigation, other than the peeling portion Oxidative damage of the alloy underlayer was observed. From these results, it was confirmed that the gas turbine high-temperature component provided with the TBC of the present invention has excellent heat resistance.

(Example 2)
FIG. 2 is a schematic cross-sectional view of the main part of the power generation gas turbine. The gas turbine has a rotating shaft (rotor) 49 at the center, a moving blade 46 installed around the rotating shaft 49, a stationary blade 45 supported on the casing 48 side, and a turbine shroud 47 inside the turbine casing 48. A turbine unit 44 is provided. A combustor 40 and a compressor 50 are connected to the turbine section 44 and suck in the atmosphere to obtain compressed air for combustion and a cooling medium. The combustor 40 includes a combustor nozzle 41 that mixes and injects compressed air supplied from the compressor 50 and supplied fuel (not shown), and combusts the mixture in the combustor liner 42. Thus, high-temperature and high-pressure combustion gas is generated, and this combustion gas is supplied to the turbine 44 via the transition piece (tail tube) 43, whereby the rotor 49 rotates at a high speed. A part of the compressed air discharged from the compressor 50 is used as internal cooling air for the liner 42, the transition piece 43, the turbine stationary blade 45, and the turbine rotor blade 46 of the combustor 40. The high-temperature and high-pressure combustion gas generated in the combustor 40 is rectified by the turbine stationary blade 45 through the transition piece 43 and injected to the rotor blade 46 to rotate and drive the turbine unit 44. Although not shown, it is generally configured to generate power with a generator coupled to the end of the rotating shaft 49.

本実施例は、前述の実施例１に記載の本発明のＴＢＣを、動翼４５に加え、さらに、静翼４６、燃焼器ライナ４２，トランジションピース４３の燃焼ガス通路にあたる内周面、および、初段のタービンシュラウド４７の燃焼ガス通路面に、実施例１に記載の方法に準じた方法によってＴＢＣを設けた構成とした。具体的には各部品の燃焼ガス通路表面に、放電加工によって、基体表面から内部冷却通路まで貫通した冷却孔を複数加工した。また、原料粉末として、ガスアトマイズ法で製造された、概略球状で平均粒径約５０μｍのＮｉＣｏＣｒＡｌＹ合金粉末（Ｎｉ−２３％Ｃｏ−１７％Ｃｒ−１２．５％Ａｌ−０．５％Ｙ、重量％）を準備した。コールドスプレー装置を用い、原料粉末を各部品の燃焼ガス通路面に対し、成膜した。成膜条件は、作動ガスに窒素ガスを用い、ガス圧力３ＭＰａ、ガス温度９００℃、粉末供給量１５ｇ／ｍｉｎ、成膜距離２０ｍｍの条件を用い、合金下地層２の厚さが約０．３ｍｍまで成膜を実施した。その後、合金下地層２を設けた基材１上に、イットリア部分安定化ジルコニア（ＺｒＯ2−８ｗｔ％Ｙ2Ｏ3）粉末を用い、大気中プラズマ溶射（プラズマ出力約５０ｋＷ）にて約０.３ｍｍの厚さ、気孔率が約２５％の連通気孔を有する多孔質遮熱セラミック層を設けた。この際の成膜条件としては、予熱温度が約１５０℃、溶射ガンの移動速度は４５ｍ／ｍｉｎ、溶射距離は１００ｍｍとした。さらに、遮熱セラミック層を成膜後の各部品に対し、それぞれの部品の基材として用いられている合金の熱処理条件に準じた真空中熱処理を実施した。
なお、本実施例では、３段で構成されるタービン部４４の動翼４５，静翼４６，シュラウド４７の各初段のみに、本発明のＴＢＣを設けた構成を採用したが、さらに後段の２段，３段に適用することも可能である。さらには、他の段数で構成されるタービン、例えば、２段，４段で構成されるタービンの全段落、乃至は選択された段落に適用することも可能である。 In this embodiment, the TBC of the present invention described in the first embodiment is added to the moving blade 45, and further, the inner peripheral surface corresponding to the combustion gas passage of the stationary blade 46, the combustor liner 42, and the transition piece 43, and The TBC was provided on the combustion gas passage surface of the turbine shroud 47 in the first stage by a method according to the method described in the first embodiment. Specifically, a plurality of cooling holes penetrating from the substrate surface to the internal cooling passage were machined on the surface of the combustion gas passage of each component by electric discharge machining. Further, as a raw material powder, a NiCoCrAlY alloy powder (Ni-23% Co-17% Cr-12.5% Al-0.5% Y, weight%) manufactured by a gas atomizing method and having an approximately spherical average particle diameter of about 50 μm ) Was prepared. Using a cold spray apparatus, the raw material powder was deposited on the combustion gas passage surface of each part . The film formation conditions are as follows. Nitrogen gas is used as the working gas, the gas pressure is 3 MPa, the gas temperature is 900 ° C., the powder supply rate is 15 g / min, the film formation distance is 20 mm, and the thickness of the alloy underlayer 2 is about 0.3 mm. Film formation was carried out. After that, yttria partially stabilized zirconia (ZrO 2 -8 wt% Y 2 O 3 ) powder is used on the base material 1 provided with the alloy underlayer 2 and is about 0.00 by atmospheric plasma spraying (plasma output about 50 kW). A porous heat-insulating ceramic layer having a continuous ventilation hole having a thickness of 3 mm and a porosity of about 25% was provided. The film forming conditions at this time were a preheating temperature of about 150 ° C., a spray gun moving speed of 45 m / min, and a spray distance of 100 mm. Further, heat treatment in a vacuum was performed on each component after the thermal barrier ceramic layer was formed in accordance with the heat treatment conditions of the alloy used as the base material of each component.
In the present embodiment, a configuration in which the TBC of the present invention is provided only in the first stage of each of the moving blade 45, the stationary blade 46, and the shroud 47 of the turbine section 44 configured in three stages is employed. It is also possible to apply to three or three stages. Further, the present invention can be applied to all or selected paragraphs of a turbine configured with other stages, for example, a turbine configured with two stages or four stages.

以上の構成による本実施例のガスタービンにおいて、本発明のＴＢＣを設けた部品については冷却空気を約３０％減じて運転した。２年間の運転後、各部品を観察したところ、本実施例のＴＢＣを設けたガスタービン部品では、ＴＢＣにほとんど損傷は認められず健全であった。一方、冷却空気を減じたことにより、タービンンの効率は向上した。   In the gas turbine according to the present embodiment having the above-described configuration, the parts provided with the TBC of the present invention were operated with cooling air reduced by about 30%. When each part was observed after two years of operation, the gas turbine part provided with the TBC of this example was healthy with almost no damage observed in the TBC. On the other hand, the efficiency of the turbine was improved by reducing the cooling air.

以上の結果から、本実施例のガスタービンは、その優れた高温部品の耐熱性により、高温で運転することが可能となり、経済性，安定運用性に優れる。   From the above results, the gas turbine according to the present embodiment can be operated at a high temperature due to the excellent heat resistance of the high-temperature parts, and is excellent in economic efficiency and stable operation.

１ 基材
２ 合金下地層
３ 遮熱セラミック層
４ 冷却孔
５ 合金粉末粒子
６ 粒子間の間隙
７ クラック
８ 流体冷媒
４０ 燃焼器
４１ 燃焼器ノズル
４２ 燃焼器ライナ
４３ トランジションピース
４４ タービン
４５ タービン動翼
４６ タービン静翼
４７ タービンシュラウド
４８ タービンケーシング
４９ タービンロータ
５０ 圧縮機 DESCRIPTION OF SYMBOLS 1 Base material 2 Alloy base layer 3 Thermal insulation ceramic layer 4 Cooling hole 5 Alloy powder particle 6 Intergranular space 7 Crack 8 Fluid refrigerant 40 Combustor 41 Combustor nozzle 42 Combustor liner 43 Transition piece 44 Turbine
45 Turbine blade 46 Turbine stationary blade 47 Turbine shroud 48 Turbine casing
49 Turbine rotor 50 Compressor

Claims (10)

高温の燃焼ガスに曝される基材表面に、合金下地層を設け、さらに、その表面上に遮熱セラミック層を設けてなる遮熱コーティングを有するガスタービン用高温部材において、前記合金下地層は、合金粉末粒子が概略球形を維持した状態で積層されたものであり、前記合金下地層と遮熱セラミック層に、基材側から表面側に連通した微細通路を設け、前記微細通路を前記合金下地層の概略球形が維持された合金粉末粒子間の間隙により形成されたものとし、部材を冷却する冷媒の一部を、これら微細通路を通じて、部材外部に流出させることを特徴とするガスタービン用高温部材。 In a high-temperature member for a gas turbine having a thermal barrier coating in which an alloy base layer is provided on the surface of a base material exposed to a high-temperature combustion gas and a thermal barrier ceramic layer is provided on the surface, the alloy base layer is The alloy powder particles are laminated while maintaining a substantially spherical shape, and the alloy base layer and the heat-shielding ceramic layer are provided with a fine passage communicating from the substrate side to the surface side, and the fine passage is provided in the alloy. It is formed by a gap between alloy powder particles in which the substantially spherical shape of the underlayer is maintained, and a part of the refrigerant that cools the member flows out of the member through these fine passages. High temperature member. 前記基材が、Ｎｉ基、Ｃｏ基、またはＦｅ基の耐熱合金からなることを特徴とする請求項１記載のガスタービン用高温部材。   The high temperature member for a gas turbine according to claim 1, wherein the base material is made of a heat resistant alloy of Ni base, Co base, or Fe base. 前記合金下地層が、ＭＣｒＡｌＹ（Ｍは、Ｆｅ，Ｎｉ，Ｃｏから選ばれる少なくとも１種）合金からなることを特徴とする請求項１記載のガスタービン用高温部材。   The high-temperature member for a gas turbine according to claim 1, wherein the alloy underlayer is made of an MCrAlY (M is at least one selected from Fe, Ni, and Co) alloy. 前記合金下地層は、粒径の範囲が５〜１００μｍの前記合金粉末粒子の積層組織を有し、かつ、積層粒子間の隙間によって形成される連通した微細通路の皮膜内体積分率が３０〜７０％であることを特徴とする、請求項１記載のガスタービン用高温部材。 The alloy undercoating layer is in the range of particle size has a stacked structure of the alloy powder particles of 5 to 100 [mu] m, and coating the volume fraction of fine passages in communication formed by gaps between the stacked particles is 30 The high-temperature member for a gas turbine according to claim 1, wherein the high-temperature member is 70%. 前記合金下地層が、前記合金粉末粒子を、合金の融点以下の温度の作動ガスによって加速して溶融を伴わずに、高速で基材表面に衝突させる方法で形成されることを特徴とする、請求項１記載のガスタービン用高温部材。 The alloy undercoating layer, the alloy powder particles, without melting and accelerated by working gas temperature below the melting point of the alloy, and wherein the high speed by being formed by a method of colliding to the substrate surface, The high temperature member for gas turbines of Claim 1. 前記遮熱セラミック層が、部分安定化ジルコニアであることを特徴とする、請求項１記載のガスタービン用高温部材。   The high-temperature member for a gas turbine according to claim 1, wherein the thermal barrier ceramic layer is partially stabilized zirconia. 前記遮熱セラミック層の微細通路が、クラックによって形成されていることを特徴とする、請求項１記載のガスタービン用高温部材。   The high-temperature member for a gas turbine according to claim 1, wherein the fine passages of the thermal barrier ceramic layer are formed by cracks. 前記遮熱セラミック層の微細通路が、気孔によって形成されていることを特徴とする、請求項１記載のガスタービン用高温部材。   The high-temperature member for a gas turbine according to claim 1, wherein the fine passage of the heat-insulating ceramic layer is formed by pores. 請求項１乃至８のいずれかに記載のガスタービン用高温部材を備えたことを特徴とするガスタービン。   A gas turbine comprising the high-temperature member for a gas turbine according to claim 1. 請求項９に記載のガスタービンを備えたことを特徴とするガスタービン複合発電プラント。   A gas turbine combined power plant comprising the gas turbine according to claim 9.
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