JP3422611B2 - Gas turbine blade - Google Patents
Gas turbine bladeInfo
- Publication number
- JP3422611B2 JP3422611B2 JP386096A JP386096A JP3422611B2 JP 3422611 B2 JP3422611 B2 JP 3422611B2 JP 386096 A JP386096 A JP 386096A JP 386096 A JP386096 A JP 386096A JP 3422611 B2 JP3422611 B2 JP 3422611B2
- Authority
- JP
- Japan
- Prior art keywords
- blade
- thinning
- gas turbine
- tip
- present
- 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.)
- Expired - Fee Related
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
- F01D5/20—Specially-shaped blade tips to seal space between tips and stator
Description
【発明の詳細な説明】
【０００１】
【発明の属する技術分野】本発明は動翼先端部にシンニ
ング部を設けたガスタービン動翼に関する。
【０００２】
【従来の技術】従来のタービン動翼の翼先端部構造を図
６，図７に示す。図６は動翼先端部におけるシンニング
部の外観斜視図、図７は図６のＢ−Ｂ線に沿う断面図で
ある。図６，図７に示すように、ガスタービン動翼の翼
先端は、静止部２と萬が一接触を生じても動翼本体３に
損傷を生じないように、先端は翼型をしているが、内側
をくりぬき部４によって薄くしてシンニング７を形成し
ている。
【０００３】一方、動翼腹側５から動翼背側６に向って
図のように高温の主流ガス１がリークする。動翼本体３
は冷却空気によって内面から冷却されその冷却空気を冷
却孔１０から吹き出し、シンニング部をも冷却しようと
しているが、この翼先端のシンニング７は、実質上、動
翼本体３と接続している部分の熱伝導で冷却されるのみ
で、高温の主流ガス１に３方を囲れている事から高温に
なり易い。この温度が翼材料の酸化限界を越えると、シ
ンニング７は酸化し、流体力によって摩耗し、高さを減
じ、高温の主流ガス１のリークをシールする役割を減じ
てしまう。
【０００４】また、従来のガスタービン動翼におけるシ
ンニングでは、シンニング部が高温となり、一方、翼部
は冷却されていてメタル温度は低く、この大きな温度差
のためシンニング部には引張の熱応力が働き、クラック
を生じる危険があった。
【０００５】
【発明が解決しようとする課題】この為、前記したよう
に動翼本体３を冷却する空気を、冷却孔１０から吹き出
し、シンニング部７を冷い冷却空気でおおう事により酸
化を防止する方法がとられるが、冷却空気は、先端のへ
こみ部で主流高温ガス１と混合し、冷却空気を多く流す
割りには、翼先端を効果的に冷却する事が困難であっ
た。
【０００６】本発明は、従来のガスタービン動翼に見ら
れたように、翼先端が高温になって酸化し摩耗し、翼腹
側から翼背側に主流ガスがリークするのを防止するよう
翼先端を効果的に冷却可能なガスタービン動翼を提供す
ることを課題としている。
【０００７】
【課題を解決するための手段】本発明は、ガスタービン
動翼における前記課題を解決するため、動翼先端部にお
互いの間に小隙をもって形成されたフィンを有するシン
ニング部を設けた構成のガスタービン動翼を提供する。
【０００８】本発明のガスタービン動翼に設けるフィン
は、図６及び図７で説明したような従来のシンニングか
ら放電加工で作ってもよいし、また、精密鋳造で翼部と
同時に製作してもよい。或いはまた、別途フィンを製作
したものをロー付け等で動翼先端に取り付けても良い。
【０００９】本発明によってガスタービン動翼先端部に
設けるフィンはシンニング部の厚さに依存するが２〜３
列設けた構造とするのが好ましい。また、フィンはお互
いの間に小隙を有するように列をなして形成する。な
お、フィン単体の形状は限定されない。
【００１０】本発明によるガスタービン動翼では、動翼
先端部に２〜３列のフィンを有するシンニング部を設け
ているので、主流ガスが翼腹側から翼背側にリークする
場合、流体抵抗が大きくなりリークしにくくなる。そし
て、翼の腹側から吹き出した冷却空気は、フィン形状の
シンニング部を流れフィンの側面を効果的に冷却する。
【００１１】このように、本発明のガスタービン動翼で
は、翼先端が高温になって酸化し摩耗を生じて翼腹側か
ら翼背側に主流ガスがリークするような事態が生ずるの
を防止することができる。
【００１２】
【発明の実施の形態】以下、本発明によるガスタービン
動翼について図１〜図３に示した実施の形態に基づいて
具体的に説明する。なお、以下の実施の形態において、
図６及び図７に示した従来の装置と同じ構成の部分には
説明を簡単にするため同じ符号を付してある。図１に本
発明の実施の一形態によるガスタービン動翼のシンニン
グ部の外観斜視図を、図２には、図１のＡ−Ａ線に沿う
断面図を、図３には同じく図１のＣ矢視の概略図を示
す。
【００１３】本実施の形態によるガスタービン動翼で
は、動翼先端に、２列フィン構造からなるシンニング９
を構成してあり、フィルム冷却孔１０より吹き出した冷
却空気８によってこのフィン構造のシンニング９を冷却
する。このように、フィン形状のシンニング９はフィン
の側面をフィルム冷却空気で冷却する事が出来る。
【００１４】このシンニング９部分が萬が一、静止部と
接触が生じても、フィンの剛性を小さく設計しておくこ
とによって、シンニング９部の損傷を減らすことができ
る。更に、シンニング９はフィン構造であって一体でな
いため、たとえシンニング９部と動翼本体３の間に大き
な温度差が生じたとしても、フィン部に作用する熱応力
は非常に小さいものとなり、クラック等を生じる恐れが
無くなる。
【００１５】なお、本発明に基づいてガスタービン動翼
に設けるフィンは、従来のシンニング（図６）から放電
加工で作ってもよいし、また、精密鋳造で翼部と同時に
製作してもよい。或いはまた、別途フィンを製作したも
のをロー付け等で動翼先端に取り付けても良い。
【００１６】次に、本発明のガスタービン動翼に設けた
シンニング部を構成するフィンの具体例を図４及び図５
により説明する。図４は丸棒型のフィンを設けたもの
で、ａは０．１mm，ｂは０．５mmである。また、図５は
方型のフィンを設けたもので、ｃは０．１mm，ｄは１mm
である。
【００１７】以上、本発明を図示した実施形態に基づい
て具体的に説明したが、本発明がこれらの実施形態に限
定されず特許請求の範囲に示す本発明の範囲内で、その
具体的構造に種々の変更を加えてよいことはいうまでも
ない。例えば、上記実施の形態ではシンニングを構成す
るフィンを２列設けているが、これはシンニング部の厚
さに応じて適宜の数を選択してよい。
【００１８】
【発明の効果】以上説明したように、本発明によるガス
タービン動翼では、動翼先端部にお互いの間に小隙をも
って形成されたフィン構造のシンニング部を設けている
ので、これによって翼腹側から翼背側への主流高温ガス
のリークを効率的に防止する。また、本発明は、接触に
対して安全で、かつ、冷却され易く酸化減肉の少ないク
ラックの入らない信頼性の高いシンニングをもつガスタ
ービン動翼を提供することができる。Description: BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine rotor blade provided with a thinning portion at the tip of the rotor blade. 2. Description of the Related Art FIGS. 6 and 7 show a tip structure of a conventional turbine blade. FIG. 6 is an external perspective view of a thinning portion at the blade tip, and FIG. 7 is a cross-sectional view taken along line BB of FIG. As shown in FIGS. 6 and 7, the blade tip of the gas turbine blade has an airfoil shape so that the blade body 3 is not damaged even if the stationary part 2 makes contact with the blade. However, the thinning 7 is formed by making the inside thinner by the hollow portion 4. On the other hand, high-temperature mainstream gas 1 leaks from blade moving side 5 to blade moving side 6 as shown in the figure. Blade body 3
Is cooled from the inner surface by the cooling air, and the cooling air is blown out from the cooling holes 10 to also cool the thinning portion. However, the thinning 7 at the tip of the blade substantially corresponds to the portion connected to the blade body 3. It is only cooled by heat conduction, and is easily heated to high temperature because it is surrounded by the high temperature mainstream gas 1 on three sides. If this temperature exceeds the oxidation limit of the wing material, the thinning 7 will oxidize and wear out due to fluid forces, reducing its height and reducing its role in sealing the leak of hot mainstream gas 1. In the thinning of a conventional gas turbine rotor blade, the thinning portion has a high temperature, while the blade portion is cooled and the metal temperature is low. Due to this large temperature difference, tensile thermal stress is applied to the thinning portion. There was a risk of working and cracking. [0005] Therefore, as described above, the air for cooling the blade body 3 is blown out from the cooling holes 10 and the thinning portion 7 is covered with cool cooling air to prevent oxidation. However, the cooling air is mixed with the mainstream high-temperature gas 1 at the dent portion at the tip, and it is difficult to effectively cool the tip of the blade, even though a large amount of cooling air flows. According to the present invention, as seen in a conventional gas turbine blade, the tip of the blade becomes hot and oxidizes and wears, thereby preventing the mainstream gas from leaking from the blade aft side to the blade back side. An object of the present invention is to provide a gas turbine rotor blade capable of effectively cooling a blade tip. According to the present invention, in order to solve the above-mentioned problems in a gas turbine rotor blade, a thinning portion having a fin formed with a small gap between each other is provided at the tip of the rotor blade. Provided is a gas turbine rotor blade having a configuration as described above. The fins provided on the gas turbine rotor blade of the present invention may be formed by electric discharge machining from the conventional thinning described with reference to FIGS. 6 and 7, or may be manufactured simultaneously with the blade portion by precision casting. Is also good. Alternatively, a separately manufactured fin may be attached to the blade tip by brazing or the like. According to the present invention, the fin provided at the tip of the gas turbine rotor blade depends on the thickness of the thinning portion.
It is preferable to adopt a structure in which rows are provided. Also, the fins are formed in rows so as to have a small gap between each other. Note that the shape of the fin alone is not limited. In the gas turbine rotor blade according to the present invention, since the thinning portion having two or three rows of fins is provided at the tip of the rotor blade, when the mainstream gas leaks from the blade vent side to the blade back side, the fluid resistance is reduced. Becomes large and it becomes difficult to leak. The cooling air blown out from the ventral side of the wing flows through the fin-shaped thinning portion to effectively cool the side surface of the fin. As described above, in the gas turbine rotor blade of the present invention, it is possible to prevent a situation in which the tip of the blade becomes high temperature, oxidizes and wears, and a main gas leaks from the blade vent side to the blade rear side. can do. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS Hereinafter, a gas turbine blade according to the present invention will be described in detail with reference to the embodiments shown in FIGS. In the following embodiment,
6 and 7 are denoted by the same reference numerals for the sake of simplicity. FIG. 1 is an external perspective view of a thinning portion of a gas turbine rotor blade according to an embodiment of the present invention, FIG. 2 is a cross-sectional view taken along line AA of FIG. 1, and FIG. The schematic diagram seen from arrow C is shown. In the gas turbine blade according to the present embodiment, a thinning 9 having a two-row fin structure is provided at the tip of the blade.
The thinning 9 having the fin structure is cooled by the cooling air 8 blown out from the film cooling holes 10. Thus, the fin-shaped thinning 9 can cool the side surface of the fin with the film cooling air. Even if the thinning 9 portion comes into contact with the stationary portion, damage to the thinning 9 portion can be reduced by designing the fin to have a small rigidity. Further, since the thinning 9 has a fin structure and is not integrated, even if a large temperature difference occurs between the thinning 9 portion and the rotor blade body 3, the thermal stress acting on the fin portion becomes very small, and cracks occur. And the like. The fins provided on the gas turbine rotor blade according to the present invention may be manufactured by electric discharge machining from conventional thinning (FIG. 6) or may be manufactured simultaneously with the blade by precision casting. . Alternatively, a separately manufactured fin may be attached to the blade tip by brazing or the like. Next, specific examples of the fins constituting the thinning portion provided in the gas turbine rotor blade of the present invention will be described with reference to FIGS.
This will be described below. FIG. 4 shows a case in which round bar fins are provided, wherein a is 0.1 mm and b is 0.5 mm. FIG. 5 shows a case in which rectangular fins are provided, where c is 0.1 mm and d is 1 mm.
It is. As described above, the present invention has been specifically described based on the illustrated embodiments. However, the present invention is not limited to these embodiments, and specific structures within the scope of the present invention are set forth in the appended claims. It is needless to say that various changes may be made to. For example, in the above embodiment, two rows of fins constituting thinning are provided, but an appropriate number of fins may be selected according to the thickness of the thinning portion. As described above, in the gas turbine rotor blade according to the present invention, since the thinning portion of the fin structure formed with a small gap between the rotor blade tips is provided at the tip of the rotor blade. This effectively prevents the mainstream hot gas from leaking from the blade ventral side to the blade rear side. Further, the present invention can provide a gas turbine blade having a thinning which is safe against contact, is easy to be cooled, has little oxidation thinning and does not contain cracks, and has high reliability.
【図面の簡単な説明】
【図１】本発明の実施の一形態に係るガスタービン動翼
におけるシンニング部の外観斜視図。
【図２】図１のＡ−Ａ線に沿う拡大断面図。
【図３】図１のＣ矢視図。
【図４】本発明によるガスタービン動翼におけるシンニ
ング部の具体例の平面図。
【図５】本発明によるガスタービン動翼におけるシンニ
ング部の他の具体例の平面図。
【図６】従来のガスタービン動翼におけるシンニング部
を示す外観斜視図。
【図７】図６のＢ−Ｂ線に沿う拡大断面図。
【符号の説明】
１ 主流高温ガス
２ 静止部
３ 動翼本体
４ 動翼先端くりぬき部
５ 動翼腹側
６ 動翼背側
７ シンニング部
８ 冷却空気
９ フィン構造のシンニング
１０ フィルム冷却孔BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is an external perspective view of a thinning portion in a gas turbine rotor blade according to an embodiment of the present invention. FIG. 2 is an enlarged sectional view taken along the line AA of FIG. FIG. 3 is a view taken in the direction of the arrow C in FIG. 1; FIG. 4 is a plan view of a specific example of a thinning portion in the gas turbine bucket according to the present invention. FIG. 5 is a plan view of another specific example of a thinning portion in the gas turbine bucket according to the present invention. FIG. 6 is an external perspective view showing a thinning portion in a conventional gas turbine rotor blade. FIG. 7 is an enlarged sectional view taken along line BB of FIG. 6; [Description of Signs] 1 Mainstream high-temperature gas 2 Stationary portion 3 Moving blade body 4 Blade tip hollowing portion 5 Rotating blade vent side 6 Rotating blade backside 7 Thinning portion 8 Cooling air 9 Finning thinning 10 Film cooling hole
フロントページの続き (56)参考文献 特開 平７−229403（ＪＰ，Ａ) 特開 平７−243303（ＪＰ，Ａ) 特開 昭62−186004（ＪＰ，Ａ) 特開 平７−293202（ＪＰ，Ａ) 特開 平６−229204（ＪＰ，Ａ) 特開 昭57−13201（ＪＰ，Ａ) 特開 昭63−41603（ＪＰ，Ａ) 実開 昭57−107802（ＪＰ，Ｕ) (58)調査した分野(Int.Cl.7，ＤＢ名) F01D 1/00 - 11/10 Continuation of the front page (56) References JP-A-7-229403 (JP, A) JP-A-7-243303 (JP, A) JP-A-62-186004 (JP, A) JP-A-7-293202 (JP) JP-A-6-229204 (JP, A) JP-A-57-13201 (JP, A) JP-A-63-41603 (JP, A) JP-A-57-107802 (JP, U) (58) Field surveyed (Int.Cl. 7 , DB name) F01D 1/00-11/10
Claims (1)
にお互いの間に小隙をもって形成されたフィンを有する
シンニング部を設けたことを特徴とするガスタービン動
翼。(1) A gas turbine moving blade, wherein a thinning portion having a fin formed with a small gap between the moving blade tips is provided at the moving blade tip. Bucket.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP386096A JP3422611B2 (en) | 1996-01-12 | 1996-01-12 | Gas turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP386096A JP3422611B2 (en) | 1996-01-12 | 1996-01-12 | Gas turbine blade |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPH09195704A JPH09195704A (en) | 1997-07-29 |
| JP3422611B2 true JP3422611B2 (en) | 2003-06-30 |
Family
ID=11568956
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP386096A Expired - Fee Related JP3422611B2 (en) | 1996-01-12 | 1996-01-12 | Gas turbine blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP3422611B2 (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9273561B2 (en) | 2012-08-03 | 2016-03-01 | General Electric Company | Cooling structures for turbine rotor blade tips |
-
1996
- 1996-01-12 JP JP386096A patent/JP3422611B2/en not_active Expired - Fee Related
Also Published As
| Publication number | Publication date |
|---|---|
| JPH09195704A (en) | 1997-07-29 |
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