JPH09144504A - Turbine cooling blade and its working method - Google Patents
Turbine cooling blade and its working methodInfo
- Publication number
- JPH09144504A JPH09144504A JP7304592A JP30459295A JPH09144504A JP H09144504 A JPH09144504 A JP H09144504A JP 7304592 A JP7304592 A JP 7304592A JP 30459295 A JP30459295 A JP 30459295A JP H09144504 A JPH09144504 A JP H09144504A
- Authority
- JP
- Japan
- Prior art keywords
- cooling
- blade
- turbine
- film
- film cooling
- 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
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- 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
-
- 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
- F05D2250/00—Geometry
- F05D2250/10—Two-dimensional
- F05D2250/11—Two-dimensional triangular
-
- 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
- F05D2250/00—Geometry
- F05D2250/10—Two-dimensional
- F05D2250/14—Two-dimensional elliptical
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Laser Beam Processing (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【０００１】[0001]
【発明の属する技術分野】本発明は、タービン翼の表面
に冷却ガスの薄いフィルムを形成するためのフィルム冷
却孔を備えたタービン冷却翼とその加工方法に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a turbine cooling blade provided with film cooling holes for forming a thin film of cooling gas on the surface of a turbine blade and a method for processing the same.
【０００２】[0002]
【従来の技術】ガスタービン及びガスタービンを用いた
複合サイクルの熱効率は、タービン入口温度を高めるこ
とによって著しく向上させることができる。このため、
図８に示すように、これまで耐熱材料の開発や冷却技術
の改良によって、タービン入口温度の高温化が図られて
きている。The thermal efficiency of gas turbines and combined cycles using gas turbines can be significantly improved by increasing the turbine inlet temperature. For this reason,
As shown in FIG. 8, the turbine inlet temperature has been increased by developing heat-resistant materials and improving cooling technology.
【０００３】かかるタービン翼に用いる耐熱材料には、
従来、主としてＣｏ基及びＮｉ基の耐熱金属が用いられ
てきたが、現在、耐熱合金より優れた耐熱材料として、
ＳｉＣやＳｉ3 Ｎ4 などのエンジニアリングセラミック
が注目され、研究開発が行われている（図９）。かかる
セラミック翼の適用により、耐用温度を約１４００℃前
後にまで高めることができる。Heat resistant materials used for such turbine blades include
Conventionally, mainly Co-based and Ni-based refractory metals have been used, but nowadays, as a heat-resistant material superior to heat-resistant alloys,
Engineering ceramics such as SiC and Si 3 N 4 are drawing attention and are being researched and developed (Fig. 9). By applying such a ceramic blade, the service temperature can be increased to about 1400 ° C.
【０００４】また、タービン翼の耐熱材料を許容温度以
下に保つために、圧縮空気の一部を用いて翼を冷却する
対流冷却、インピンジ冷却、フィルム冷却、トランスピ
レーション冷却、等の冷却技術が、従来、単独に或いは
組合わせて用いられている。このうち、フィルム冷却
は、図１０に例示するように、タービン翼１の表面に空
気吹出孔２（フィルム冷却孔）を設け、この孔から冷却
空気３を吹き出してタービン翼の表面に冷却空気の薄い
フィルムを形成するものである。かかるフィルム冷却孔
２には、丸孔又は矩形孔が主として用いられている（例
えば、特開昭６２−１６５５０５号公報、特願平６−１
６１７３４号、未公開、等）。かかるフィルム冷却は、
翼への流入熱量を減少させる効果が高く、タービン入口
温度の高温化のために特に適していることが知られてい
る。Further, in order to keep the heat-resistant material of the turbine blade below the allowable temperature, there are cooling techniques such as convection cooling, impingement cooling, film cooling, and transpiration cooling in which a part of compressed air is used to cool the blade. Conventionally, they are used alone or in combination. Among them, in the film cooling, as illustrated in FIG. 10, an air blowout hole 2 (film cooling hole) is provided on the surface of the turbine blade 1, and cooling air 3 is blown from this hole to cool the cooling air on the surface of the turbine blade. It forms a thin film. A round hole or a rectangular hole is mainly used as the film cooling hole 2 (for example, JP-A-62-165505, Japanese Patent Application No. 6-1).
61734, unpublished, etc.). Such film cooling is
It is known that the effect of reducing the amount of heat flowing into the blade is high, and that it is particularly suitable for increasing the turbine inlet temperature.
【０００５】[0005]
【発明が解決しようとする課題】航空エンジンの性能向
上と低公害化を図るために、燃料に水素を用いた水素燃
焼タービンの開発が進められている。この水素燃焼ター
ビンは、従来の航空エンジンと異なり、主流ガスと冷却
ガスが主として水蒸気となる。空気と比べて水蒸気は熱
伝導が良いことから、ガスタービンの動翼，静翼に流入
する熱量が増大し、熱負荷が大きくなる。これに対処
し、ガスタービンの性能向上を図るためには、翼への流
入熱量を減少させる効果があるフィルム冷却の性能を更
に高める必要がある。In order to improve the performance of an aviation engine and reduce pollution, a hydrogen combustion turbine using hydrogen as a fuel is under development. Unlike the conventional aviation engine, this hydrogen combustion turbine mainly uses steam as the mainstream gas and the cooling gas. Since water vapor has better heat conduction than air, the amount of heat flowing into the moving blades and stationary blades of the gas turbine increases, and the heat load increases. In order to cope with this and improve the performance of the gas turbine, it is necessary to further enhance the film cooling performance, which has the effect of reducing the amount of heat flowing into the blade.
【０００６】一方、航空エンジンの性能を飛躍的に高め
るためには、タービン入口温度を従来の最高１３００℃
前後から約１７００℃前後にまで高める必要があり、こ
れを実現するためには、タービン翼を従来の耐熱材料か
ら高温性能の優れたエンジニアリングセラミック（Ｓｉ
Ｃ，Ｓｉ3 Ｎ4 ，等）に変更する必要がある。On the other hand, in order to dramatically improve the performance of an aero engine, the turbine inlet temperature is set to a maximum of 1300 ° C, which is the conventional maximum.
It is necessary to raise the temperature from around 1700 ° C to around 1700 ° C. In order to realize this, turbine blades can be made from conventional heat-resistant materials to engineering ceramics (Si
C, Si 3 N 4 , etc.).
【０００７】しかし、従来のフィルム冷却孔は、丸孔又
は矩形孔であり、丸孔のフィルム冷却孔は、フィルム効
率ηが低く、矩形孔のフィルム冷却孔は、加工が困難で
ある問題点があった。すなわち、従来の矩形孔は、矩形
孔に相当する電極を用いて放電加工により形成していた
が、この放電加工は電極によって孔形状が決まるため、
形状の微調整が困難であり、各フィルム冷却孔の流量を
均一化するには、一部に圧力損失の大きい絞りを設ける
必要があった。また、かかる放電加工は導電性のある金
属材料にしか適用できず、高温性能の優れたエンジニア
リングセラミックの加工ができない問題点があった。一
方、丸孔は、例えばレーザ加工により加工できるが、上
述したフィルム効率ηが低い問題点があった。However, the conventional film cooling holes are round holes or rectangular holes. The film cooling holes with round holes have a low film efficiency η, and the film cooling holes with rectangular holes are difficult to process. there were. That is, the conventional rectangular hole was formed by electric discharge machining using an electrode corresponding to the rectangular hole, but since the hole shape is determined by the electrode in this electric discharge machining,
It is difficult to finely adjust the shape, and in order to make the flow rate of each film cooling hole uniform, it was necessary to provide a throttle with a large pressure loss in part. Further, such electric discharge machining can be applied only to a conductive metal material, and there is a problem that an engineering ceramic excellent in high temperature performance cannot be machined. On the other hand, round holes can be processed by, for example, laser processing, but there is a problem that the above-mentioned film efficiency η is low.
【０００８】なお、フィルム効率ηは、フィルム冷却の
有効性を示すものであり、図１０において、主流の温度
をＴｇ、冷却流体膜の温度をＴｆ、空気吹出孔２におけ
る冷却空気の温度をＴｃ、とする場合に、η＝（Ｔｇ−
Ｔｆ）／（Ｔｇ−Ｔｃ）で定義される。The film efficiency η indicates the effectiveness of film cooling. In FIG. 10, the temperature of the main stream is Tg, the temperature of the cooling fluid film is Tf, and the temperature of the cooling air in the air outlet 2 is Tc. , And η = (Tg−
It is defined by Tf) / (Tg-Tc).
【０００９】本発明は、上述した問題点を解決するため
に創案されたものである。すなわち、本発明の目的は、
個々のフィルム冷却孔を自由に調整でき、圧力損失を過
大にすることなく流量の均一化を図ることができ、高温
性能の優れたエンジニアリングセラミック製のタービン
翼にも適用でき、かつフィルム効率ηを高めることがで
きるタービン冷却翼とその加工方法を提供することにあ
る。The present invention has been made to solve the above problems. That is, the object of the present invention is:
Individual film cooling holes can be adjusted freely, the flow rate can be made uniform without excessive pressure loss, it can be applied to engineering ceramic turbine blades with excellent high-temperature performance, and the film efficiency η It is an object of the present invention to provide a turbine cooling blade that can be enhanced and a processing method thereof.
【００１０】[0010]
【課題を解決するための手段】本発明によれば、中空タ
ービン翼の表面に冷却ガスを吹き出すための複数のフィ
ルム冷却孔を備え、該フィルム冷却孔は、タービン翼の
内部から主流の下流側に拡がるデュフューザー状に形成
されている、ことを特徴とするタービン冷却翼が提供さ
れる。本発明の好ましい実施形態によれば、前記フィル
ム冷却孔は、タービン翼の内部に頂点を有する円錐形状
を有する。また、前記フィルム冷却孔は、タービン翼の
内部に頂点を有する円錐形状が翼の長さ方向に連なった
横長デュフューザー形状を有する、ことが好ましい。According to the present invention, a plurality of film cooling holes for blowing a cooling gas are provided on the surface of a hollow turbine blade, and the film cooling holes are provided from the inside of the turbine blade on the downstream side of the main flow. A turbine cooling blade is provided, which is formed in the shape of a diffuser that spreads over. According to a preferred embodiment of the present invention, the film cooling hole has a conical shape having an apex inside the turbine blade. Further, it is preferable that the film cooling hole has a laterally long diffuser shape in which a conical shape having an apex inside the turbine blade is continuous in the blade length direction.
【００１１】更に、本発明によれば、中空タービン翼の
表面にタービン翼の内部に焦点を有するレーザビームを
照射し、該レーザビームにより中空タービン翼にフィル
ム冷却孔を形成する、ことを特徴とするタービン冷却翼
の加工方法が提供される。Further, according to the present invention, the surface of the hollow turbine blade is irradiated with a laser beam having a focus inside the turbine blade, and the laser beam forms a film cooling hole in the hollow turbine blade. A method for processing a turbine cooling blade is provided.
【００１２】上述した本発明のタービン冷却翼とその加
工方法によれば、タービン翼の内部から主流の下流側に
拡がるデュフューザー状のフィルム冷却孔を、タービン
翼の内部に焦点を有するレーザビームにより加工するの
で、従来の金属材料のみならず、従来放電加工では加工
できなかったエンジニアリングセラミック製のタービン
翼にも適用できる。また、デュフューザー状のフィルム
冷却孔をレーザビームにより加工することにより、レー
ザビームの調整により、従来の放電加工に比較して個々
のフィルム冷却孔を自由に調整でき、圧力損失を過大に
することなく流量の均一化を図ることができる。更に、
種々の実験を行った結果、後述するように、かかるデュ
フューザー状のフィルム冷却孔のフィルム効率ηを、従
来の丸孔のみならず矩形孔よりも高めることができるこ
とがわかった。According to the above-described turbine cooling blade of the present invention and the processing method thereof, the diffuser-shaped film cooling hole extending from the inside of the turbine blade to the downstream side of the main flow is processed by the laser beam having the focus inside the turbine blade. Therefore, it can be applied not only to conventional metal materials but also to turbine blades made of engineering ceramics, which could not be processed by conventional electric discharge machining. Also, by processing the film cooling holes in the shape of a diffuser with a laser beam, by adjusting the laser beam, individual film cooling holes can be freely adjusted as compared with conventional electrical discharge machining, without increasing pressure loss excessively. The flow rate can be made uniform. Furthermore,
As a result of various experiments, it was found that the film efficiency η of such a diffuser-like film cooling hole can be increased more than that of a conventional round hole as well as a rectangular hole, as described later.
【００１３】[0013]
【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て使用する。図１は、本発
明によるタービン冷却翼とその加工方法を示す全体構成
図である。この図に示すように、本発明のタービン冷却
翼１０は、中空タービン翼の表面に冷却ガスを吹き出す
ための複数のフィルム冷却孔１２を備えており、このフ
ィルム冷却孔１２は、タービン翼の内部から主流の下流
側に拡がるデュフューザー状に形成されている。DESCRIPTION OF THE PREFERRED EMBODIMENTS Preferred embodiments of the present invention will be described below with reference to the drawings. In the drawings, common parts are denoted by the same reference numerals. FIG. 1 is an overall configuration diagram showing a turbine cooling blade and a processing method thereof according to the present invention. As shown in this figure, a turbine cooling blade 10 of the present invention is provided with a plurality of film cooling holes 12 for blowing out a cooling gas on the surface of a hollow turbine blade. To a downstream side of the mainstream.
【００１４】図１に示すように、このフィルム冷却孔１
２は、レーザビーム５をレンズ１４ａで絞って焦点１４
ｂに集光するようになったレーザ加工装置１４により加
工される。すなわち、このレーザ加工装置１４を用い
て、中空タービン翼の表面にタービン翼の内部に焦点１
４ｂを有するレーザビーム５を照射し、このレーザビー
ム５により中空タービン翼にフィルム冷却孔１２を容易
に形成することができる。As shown in FIG. 1, this film cooling hole 1
2 is focused on the laser beam 5 by the lens 14a.
It is processed by the laser processing device 14 adapted to focus on b. That is, by using this laser processing device 14, the focal point 1 inside the turbine blade is formed on the surface of the hollow turbine blade.
A laser beam 5 having 4b is irradiated, and the laser beam 5 can easily form the film cooling hole 12 in the hollow turbine blade.
【００１５】この加工方法により、従来の金属材料のみ
ならず、従来放電加工では加工できなかったエンジニア
リングセラミック製のタービン翼にも適用できる。ま
た、デュフューザー状のフィルム冷却孔をレーザビーム
により加工することにより、レーザビームの調整によ
り、従来の放電加工に比較して個々のフィルム冷却孔を
自由に調整でき、圧力損失を過大にすることなく流量の
均一化を図ることができる。This processing method can be applied not only to conventional metal materials, but also to turbine blades made of engineering ceramics, which could not be processed by conventional electric discharge machining. Also, by processing the film cooling holes in the shape of a diffuser with a laser beam, by adjusting the laser beam, individual film cooling holes can be freely adjusted as compared with conventional electrical discharge machining, without increasing pressure loss excessively. The flow rate can be made uniform.
【００１６】[0016]
【実施例】図１に示した加工方法により、図２及び図３
に示す種々の形状のフィルム冷却孔を加工し、その性能
を試験した。図２及び図３において、ＴＰ１とＴＰ２
は、従来の丸孔と矩形孔であり、ＴＰ３，ＴＰ４，ＴＰ
５，ＴＰ６は、上述した本発明の方法により加工したフ
ィルム冷却孔１２である。これらの図から明らかなよう
に、本発明によるフィルム冷却孔１２は、ＴＰ３，ＴＰ
４，ＴＰ５，ＴＰ６に示すように、タービン翼の内部か
ら主流の下流側に拡がるデュフューザー状に形成されて
いる。EXAMPLE FIG. 2 and FIG. 3 are carried out by the processing method shown in FIG.
The film cooling holes of various shapes shown in Figure 3 were processed and their performance was tested. 2 and 3, TP1 and TP2
Are conventional round holes and rectangular holes, and are TP3, TP4, TP
5, TP6 are the film cooling holes 12 processed by the above-described method of the present invention. As is clear from these figures, the film cooling holes 12 according to the present invention have TP3, TP
As shown by 4, TP5 and TP6, it is formed in a diffuser shape that extends from the inside of the turbine blade to the downstream side of the main flow.
【００１７】ＴＰ３のフィルム冷却孔１２は、本発明に
よるフィルム冷却孔のうち最も基本的なものであり、タ
ービン翼の内部に頂点を有する円錐形状を有している。
この孔形状（以下、「単純円錐孔」と呼ぶ）は、図１の
加工方向で、加工中にレーザビームの軸線を一定位置に
保持することにより、容易に加工することができる。Ｔ
Ｐ４とＴＰ５のフィルム冷却孔１２は、タービン翼の内
部に頂点を有する円錐形状が翼の長さ方向に連なった横
長デュフューザー形状を有している。この孔形状（以
下、「横長円錐孔」と呼ぶ）は、図１の加工方向で、単
純円錐孔を加工中にレーザビームの軸線を翼の長さ方向
に振ることにより、容易に加工することができる。The film cooling hole 12 of TP3 is the most basic film cooling hole according to the present invention and has a conical shape having an apex inside the turbine blade.
This hole shape (hereinafter referred to as “simple conical hole”) can be easily processed in the processing direction of FIG. 1 by holding the axis of the laser beam at a constant position during processing. T
The film cooling holes 12 of P4 and TP5 have a laterally long diffuser shape in which conical shapes having vertices inside the turbine blades are continuous in the blade length direction. This hole shape (hereinafter referred to as "horizontal elongated conical hole") can be easily machined in the machining direction of Fig. 1 by swaying the axis of the laser beam in the blade length direction during machining of the simple conical hole. You can
【００１８】ＴＰ６のフィルム冷却孔１２は、図１の加
工方向で、単純円錐孔を加工中にレーザビームの軸線を
正三角形に振ることにより、加工でき、この孔形状を以
下、「三角錐孔」と呼ぶ。なお、図３に示した各寸法
は、試験において実際の使用状態との相似則を満たす寸
法であり、実機においては、ほぼこの１／１０程度の寸
法となる。また、本発明はこれらの寸法に限定されない
のは勿論である。The film cooling hole 12 of the TP6 can be processed by swinging the axis line of the laser beam into an equilateral triangle during the processing of the simple conical hole in the processing direction of FIG. ". Each dimension shown in FIG. 3 is a dimension that satisfies the rule of similarity with the actual usage in the test, and is about 1/10 of this in an actual machine. Of course, the present invention is not limited to these dimensions.
【００１９】図４は、各冷却孔のフィルム効率の比較図
であり、（Ａ）は密度比１．１の場合、（Ｂ）は密度比
２．０の場合を示している。図４において、性能に影響
を与える種々のパラメータについて、曲率は翼背側の典
型的な曲率半径と冷却孔形の比を模擬し、Blowing Rati
o （主流と冷却ガスの質量流速比：以下、質量流速比と
呼ぶ）は、０．５から２．５までについて、また密度比
は、１〜２の実機の想定範囲をカバーする条件としてい
る。この図から、ＴＰ３のフィルム冷却孔（単純円錐
孔）とＴＰ４とＴＰ５のフィルム冷却孔（横長円錐孔）
は、従来の丸孔（ＴＰ１）と矩形孔（ＴＰ２）よりも、
ほとんどの条件で高いフィルム効率を示しているのがわ
かる。FIG. 4 is a comparison diagram of the film efficiency of each cooling hole. FIG. 4A shows the case where the density ratio is 1.1, and FIG. 4B shows the case where the density ratio is 2.0. In Fig. 4, for various parameters that affect performance, the curvature simulates the ratio of the typical radius of curvature on the blade back side to the cooling hole shape, and
o (Mass flow velocity ratio between main flow and cooling gas: hereinafter referred to as mass flow velocity ratio) is from 0.5 to 2.5, and density ratio is a condition that covers the assumed range of 1 to 2 actual machines. . From this figure, the film cooling holes of TP3 (simple conical holes) and the film cooling holes of TP4 and TP5 (horizontal conical holes)
Is better than the conventional round hole (TP1) and rectangular hole (TP2).
It can be seen that the film efficiency is high under most conditions.
【００２０】図５は、従来の丸孔（ＴＰ１）と矩形孔
（ＴＰ２）に対する各冷却孔の平均フィルム効率の比率
を示す図であり、（Ａ）は丸孔に対する比率、（Ｂ）は
矩形孔に対する比率を示している。この図から、本発明
によるＴＰ３（単純円錐孔）とＴＰ４及びＴＰ５（横長
円錐孔）は、従来の丸孔（ＴＰ１）はもとより従来の放
電加工を前提としたデュフューザー孔、すなわち矩形孔
（ＴＰ２）よりも高いフィルム効率を有することが確認
できた。FIG. 5 is a diagram showing the ratio of the average film efficiency of each cooling hole to the conventional round hole (TP1) and rectangular hole (TP2), where (A) is the ratio to the round hole and (B) is the rectangular shape. The ratio to the holes is shown. From this figure, TP3 (simple conical hole) and TP4 and TP5 (horizontally conical hole) according to the present invention are not only conventional round holes (TP1) but also conventional diffuser holes, that is, rectangular holes (TP2). It was confirmed that the film had a higher film efficiency than the above.
【００２１】また、図５において、ＴＰ４とＴＰ５のフ
ィルム冷却孔（横長円錐孔）が特に高いフィルム効率を
示しており、デュフューザー孔形状については、冷却孔
の並んでいるピッチ方向に孔断面積を拡大し、孔出口に
おいて隣の孔との間隔を実質的に狭めることがフィルム
効率の向上に寄与することが確認できた。Further, in FIG. 5, the film cooling holes (horizontally elongated conical holes) of TP4 and TP5 show particularly high film efficiency. Regarding the shape of the diffuser hole, the hole cross-sectional area is arranged in the pitch direction in which the cooling holes are lined up. It was confirmed that the enlargement and the substantial reduction of the space between the adjacent holes at the hole exit contributed to the improvement of the film efficiency.
【００２２】図６及び図７は、動翼及び静翼の予測冷却
性能をそれぞれ示す図である。図４及び図５の試験結果
で得られたフィルム効率のデータを基に、１７００℃級
の水素燃焼タービンに適用した場合、１５％程度のフィ
ルム効率の向上が期待できることから、図６及び図７に
示すように、冷却効率で０．０３の効率向上が見込ま
れ、冷却ガス流量にすると、タービン入口流量に対する
割合として１．２％〜１．７％程度の流量削減効果があ
ることがわかった。FIGS. 6 and 7 are views showing the predicted cooling performances of the moving blade and the stationary blade, respectively. Based on the film efficiency data obtained in the test results of FIGS. 4 and 5, when applied to a 1700 ° C. class hydrogen combustion turbine, an improvement in film efficiency of about 15% can be expected. As shown in Fig. 7, the cooling efficiency is expected to improve by 0.03, and it has been found that when the cooling gas flow rate is used, there is a flow rate reduction effect of about 1.2% to 1.7% as a ratio to the turbine inlet flow rate. .
【００２３】なお、本発明は上述した実施例に限定され
ず、本発明の要旨を逸脱しない範囲で自由に変更できる
ことは勿論である。The present invention is not limited to the above-described embodiments, and it goes without saying that the present invention can be freely modified without departing from the gist of the present invention.
【００２４】[0024]
【発明の効果】上述したように、本発明によるタービン
冷却翼の冷却孔形状は、レーザー加工によりセラミック
に対して適用可能であり、しかも従来のデュフューザー
孔より性能を向上させることができ、高温の水素燃焼タ
ービンを小流量の冷却ガスで効率的に冷却することが期
待できる。As described above, the cooling hole shape of the turbine cooling blade according to the present invention can be applied to the ceramic by the laser processing, and further, the performance can be improved as compared with the conventional diffuser hole, and high temperature It can be expected that the hydrogen combustion turbine can be efficiently cooled with a small flow rate of cooling gas.
【００２５】すなわち、本発明のタービン冷却翼とその
加工方法は、個々のフィルム冷却孔を自由に調整でき、
圧力損失を過大にすることなく流量の均一化を図ること
ができ、高温性能の優れたエンジニアリングセラミック
製のタービン翼にも適用でき、かつフィルム効率ηを高
めることができる、等の優れた効果を有する。That is, according to the turbine cooling blade of the present invention and the processing method thereof, individual film cooling holes can be freely adjusted,
The uniform flow rate can be achieved without excessive pressure loss, it can be applied to engineering ceramic turbine blades with excellent high-temperature performance, and the film efficiency η can be increased. Have.
【図１】本発明によるタービン冷却翼とその加工方法を
示す全体構成図である。FIG. 1 is an overall configuration diagram showing a turbine cooling blade and a processing method thereof according to the present invention.
【図２】試験した種々のフィルム冷却孔の斜視図であ
る。FIG. 2 is a perspective view of various film cooling holes tested.
【図３】試験した種々のフィルム冷却孔の部分断面図で
ある。FIG. 3 is a partial cross-sectional view of various film cooling holes tested.
【図４】各冷却孔のフィルム効率の比較図である。FIG. 4 is a comparative diagram of film efficiency of each cooling hole.
【図５】従来の丸孔と矩形孔に対する各冷却孔の平均フ
ィルム効率の比率を示す図である。FIG. 5 is a diagram showing a ratio of an average film efficiency of each cooling hole to a conventional round hole and a rectangular hole.
【図６】動翼の予測冷却性能図である。FIG. 6 is a predicted cooling performance diagram of a moving blade.
【図７】静翼の予測冷却性能図である。FIG. 7 is a predicted cooling performance diagram of a stationary blade.
【図８】ガスタービンの高温化の推移を示す図である。FIG. 8 is a diagram showing changes in the temperature of the gas turbine.
【図９】耐熱合金の進歩を示す図である。FIG. 9 is a diagram showing progress in heat resistant alloys.
【図１０】従来のフィルム冷却を示す模式図である。FIG. 10 is a schematic view showing conventional film cooling.
１ タービン翼 ２ フィルム冷却孔（空気吹出孔） ３ 冷却空気 ５ レーザビーム １０ タービン冷却翼 １２ フィルム冷却孔 １４ レーザ加工装置 １４ａ レンズ １４ｂ 焦点 1 Turbine Blade 2 Film Cooling Hole (Air Blowing Hole) 3 Cooling Air 5 Laser Beam 10 Turbine Cooling Blade 12 Film Cooling Hole 14 Laser Processing Device 14a Lens 14b Focus
Claims (4)
出すための複数のフィルム冷却孔を備え、該フィルム冷
却孔は、タービン翼の内部から主流の下流側に拡がるデ
ュフューザー状に形成されている、ことを特徴とするタ
ービン冷却翼。1. A plurality of film cooling holes for blowing a cooling gas are provided on the surface of a hollow turbine blade, and the film cooling holes are formed in a diffuser shape extending from the inside of the turbine blade to the downstream side of the mainstream. A turbine cooling blade characterized by the above.
部に頂点を有する円錐形状を有する、ことを特徴とする
請求項１に記載のタービン冷却翼。2. The turbine cooling blade according to claim 1, wherein the film cooling hole has a conical shape having an apex inside the turbine blade.
部に頂点を有する円錐形状が翼の長さ方向に連なった横
長デュフューザー形状を有する、ことを特徴とする請求
項１に記載のタービン冷却翼。3. The turbine cooling blade according to claim 1, wherein the film cooling hole has a laterally long diffuser shape in which a conical shape having an apex inside the turbine blade is continuous in a length direction of the blade. .
部に焦点を有するレーザビームを照射し、該レーザビー
ムにより中空タービン翼にフィルム冷却孔を形成する、
ことを特徴とするタービン冷却翼の加工方法。4. The surface of the hollow turbine blade is irradiated with a laser beam having a focus inside the turbine blade, and the laser beam forms a film cooling hole in the hollow turbine blade.
A method for processing a turbine cooling blade, characterized in that
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP7304592A JPH09144504A (en) | 1995-11-22 | 1995-11-22 | Turbine cooling blade and its working method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP7304592A JPH09144504A (en) | 1995-11-22 | 1995-11-22 | Turbine cooling blade and its working method |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH09144504A true JPH09144504A (en) | 1997-06-03 |
Family
ID=17934863
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP7304592A Pending JPH09144504A (en) | 1995-11-22 | 1995-11-22 | Turbine cooling blade and its working method |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH09144504A (en) |
Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2001254633A (en) * | 2000-01-19 | 2001-09-21 | General Electric Co <Ge> | Method to form cooling hole |
| JP2006138624A (en) * | 2004-11-09 | 2006-06-01 | General Electric Co <Ge> | Gas turbine engine component |
| US7273351B2 (en) | 2004-11-06 | 2007-09-25 | Rolls-Royce, Plc | Component having a film cooling arrangement |
| JP2009162224A (en) * | 2007-12-28 | 2009-07-23 | General Electric Co <Ge> | Method of forming cooling hole and turbine airfoil with hybrid-formed cooling holes |
| JP2010014113A (en) * | 2008-06-30 | 2010-01-21 | General Electric Co <Ge> | Rear frame having elliptic cooling slot and associated method |
| WO2012005324A1 (en) * | 2010-07-09 | 2012-01-12 | 株式会社Ihi | Turbine blade and engine component |
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-
1995
- 1995-11-22 JP JP7304592A patent/JPH09144504A/en active Pending
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|---|---|---|---|---|
| JP2001254633A (en) * | 2000-01-19 | 2001-09-21 | General Electric Co <Ge> | Method to form cooling hole |
| JP4693250B2 (en) * | 2000-01-19 | 2011-06-01 | ゼネラル・エレクトリック・カンパニイ | Method for forming cooling holes |
| US7273351B2 (en) | 2004-11-06 | 2007-09-25 | Rolls-Royce, Plc | Component having a film cooling arrangement |
| JP2006138624A (en) * | 2004-11-09 | 2006-06-01 | General Electric Co <Ge> | Gas turbine engine component |
| JP2009162224A (en) * | 2007-12-28 | 2009-07-23 | General Electric Co <Ge> | Method of forming cooling hole and turbine airfoil with hybrid-formed cooling holes |
| JP2010014113A (en) * | 2008-06-30 | 2010-01-21 | General Electric Co <Ge> | Rear frame having elliptic cooling slot and associated method |
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| WO2012005324A1 (en) * | 2010-07-09 | 2012-01-12 | 株式会社Ihi | Turbine blade and engine component |
| US9376919B2 (en) | 2010-07-09 | 2016-06-28 | Ihi Corporation | Turbine blade and engine component |
| JP2012017721A (en) * | 2010-07-09 | 2012-01-26 | Ihi Corp | Turbine blade and engine component |
| CN103406669A (en) * | 2012-04-19 | 2013-11-27 | 中国航空工业集团公司北京航空制造工程研究所 | Method for directly machining turbine blade air film abnormal-shaped hole through laser |
| CN106583949B (en) * | 2016-11-29 | 2018-09-07 | 沈阳黎明航空发动机(集团)有限责任公司 | The processing method of the low damage of the high whirlpool hollow blade air film hole of aero-engine monocrystalline |
| CN106583949A (en) * | 2016-11-29 | 2017-04-26 | 沈阳黎明航空发动机(集团)有限责任公司 | Low-damage processing method of film holes in single crystal high-pressure-turbine hollow blade of aircraft engine |
| CN107999974A (en) * | 2017-12-08 | 2018-05-08 | 中国航发动力股份有限公司 | A kind of annular spread group hole laser spot quickly measures and its group's hole forming method |
| CN107999974B (en) * | 2017-12-08 | 2019-06-04 | 中国航发动力股份有限公司 | A kind of annular spread group hole laser spot rapid survey and its group's hole forming method |
| CN110026677A (en) * | 2019-04-24 | 2019-07-19 | 西安中科微精光子制造科技有限公司 | The laser processing of special-shaped air film hole |
| CN110202277A (en) * | 2019-04-25 | 2019-09-06 | 青岛理工大学 | A kind of blade of aviation engine air film hole processing device and working method |
| WO2020215716A1 (en) * | 2019-04-25 | 2020-10-29 | 青岛理工大学 | Aero-engine blade film hole processing apparatus and working method |
| CN112443361A (en) * | 2020-11-04 | 2021-03-05 | 西北工业大学 | A reverse air film pore structure of pit for turbine blade |
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