JP2007211618A - Gas turbine - Google Patents

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JP2007211618A
JP2007211618A JP2006030202A JP2006030202A JP2007211618A JP 2007211618 A JP2007211618 A JP 2007211618A JP 2006030202 A JP2006030202 A JP 2006030202A JP 2006030202 A JP2006030202 A JP 2006030202A JP 2007211618 A JP2007211618 A JP 2007211618A
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gas turbine
hole
cavity
blade
moving blade
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Susumu Wakazono
進 若園
Yoichi Iwasaki
洋一 岩崎
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Mitsubishi Heavy Industries Ltd
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Mitsubishi Heavy Industries Ltd
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Abstract

<P>PROBLEM TO BE SOLVED: To provide a gas turbine improved in product reliability. <P>SOLUTION: In the gas turbine 1, a turbine rotor blade 41 includes cooling passages 413, 414 which allow cooling air to flow therethrough so as to cool the rotor blade 41. The cooling passages 413, 414 are composed of a cavity portion 413 formed within the rotor blade 41 and holes 414 extending from the cavity portion 413 in the longitudinal direction of the rotor blade 41. As viewed cross-sectionally in the direction of thickness of the rotor blade 41, the width w of the cavity portion 413 is expanded in a boundary portion 416 between the cavity portion and the holes 414 (w1<w2). <P>COPYRIGHT: (C)2007,JPO&INPIT

Description

この発明は、ガスタービンに関し、さらに詳しくは、製品の信頼性を向上させ得るガスタービンに関する。   The present invention relates to a gas turbine, and more particularly to a gas turbine that can improve the reliability of a product.

近年のガスタービンでは、タービンの動翼に冷却通路が形成されており、この冷却通路に冷却空気が流通することにより動翼が冷却される。かかる構成を採用する従来のガスタービン（ガスタービン動翼）には、特許文献１に記載される技術が知られている。従来のガスタービンは、翼先端にシュラウドを有し、同翼内部に基部から先端に向けて冷却通路を設けて冷却空気を流して前記シュラウド内に導き、同シュラウドの周囲より流出させるガスタービン動翼において、前記冷却通路は基部側が一体の空洞と同空洞に設けた多数のピンフィン、先端側が先端に向かう多数の細穴からなり、同先端側の細穴の長さは翼の全長に対して１／２以下としている。   In recent gas turbines, cooling passages are formed in the moving blades of the turbine, and the moving blades are cooled by circulating cooling air through the cooling passages. For a conventional gas turbine (gas turbine blade) employing such a configuration, a technique described in Patent Document 1 is known. A conventional gas turbine has a shroud at the tip of the blade, a cooling passage is provided in the blade from the base to the tip, and the cooling air flows to guide the gas into the shroud and flow out of the shroud. In the blade, the cooling passage is composed of a plurality of pin fins provided on the base side in the same cavity and a plurality of pin fins on the base side, and a number of fine holes on the tip side toward the tip. 1/2 or less.

特開平１１−２２９８０８号公報Japanese Patent Laid-Open No. 11-229808

ここで、従来のガスタービンでは、動翼１００の冷却通路が、動翼１００の基部に形成された空洞部１０１と、この空洞部１０１から動翼１００の先端部に向かう多数の細穴（孔）１０２とから成る（図１２参照）。また、この冷却通路は、通常、空洞部１０１が動翼１００本体と共に鋳造により一時に成形され、その後に動翼１００の先端部側からドリル加工により細穴１０２が開けられる。   Here, in the conventional gas turbine, the cooling passage of the moving blade 100 has a hollow portion 101 formed at the base of the moving blade 100 and a large number of small holes (holes) from the hollow portion 101 toward the tip of the moving blade 100. ) 102 (see FIG. 12). In addition, in this cooling passage, the cavity 101 is usually formed at a time together with the main body of the moving blade 100 by casting, and then a narrow hole 102 is drilled from the tip end side of the moving blade 100 by drilling.

しかしながら、かかる構成では、動翼１００の肉厚方向にかかる空洞部１０１の幅ｗ１が狭いため、細穴１０２の加工時に位置ズレが生じて細穴１０２と空洞部１０１との連通が不完全となるおそれがある。すると、この位置にて冷却不良などが生じて製品の信頼性が低下するおそれがある。   However, in such a configuration, since the width w1 of the hollow portion 101 in the thickness direction of the rotor blade 100 is narrow, a positional shift occurs when the fine hole 102 is processed, and the communication between the fine hole 102 and the hollow portion 101 is incomplete. There is a risk. Then, in this position, a cooling failure or the like may occur and the reliability of the product may be reduced.

そこで、この発明は、上記に鑑みてされたものであって、製品の信頼性を向上させ得るガスタービンを提供することを目的とする。   Accordingly, the present invention has been made in view of the above, and an object thereof is to provide a gas turbine capable of improving the reliability of a product.

上記目的を達成するため、この発明にかかるガスタービンは、タービンの動翼に冷却通路が形成されると共に前記冷却通路に冷却空気を流通させて前記動翼の冷却を行うガスタービンであって、前記冷却通路が、前記動翼の内部に形成された空洞部と、前記空洞部から前記動翼の長手方向に延在する孔とを含み、且つ、前記動翼の肉厚方向の断面視にて、前記空洞部の幅が前記孔との境界部にて拡張されていることを特徴とする。   In order to achieve the above object, a gas turbine according to the present invention is a gas turbine in which a cooling passage is formed in a moving blade of a turbine and cooling air is circulated through the cooling passage to cool the moving blade. The cooling passage includes a cavity formed inside the blade, and a hole extending from the cavity in the longitudinal direction of the blade, and in a cross-sectional view in the thickness direction of the blade. The width of the cavity is expanded at the boundary with the hole.

このガスタービンでは、動翼の肉厚方向の断面視にて、空洞部の幅が孔との境界部にて拡張されているので、幅狭な（拡幅されていない）空洞部に対して動翼の長手方向から孔が開けられる構成と比較して、孔の加工が容易である。これにより、加工時の位置ズレによる孔の連通不良などが低減されるので、製品の信頼性が向上する利点がある。   In this gas turbine, since the width of the cavity is expanded at the boundary with the hole in the cross-sectional view of the rotor blade in the thickness direction, the rotor is moved against the narrow (not widened) cavity. Compared with a configuration in which a hole is formed from the longitudinal direction of the blade, the hole is easily processed. As a result, poor hole communication due to misalignment during processing is reduced, and thus there is an advantage that the reliability of the product is improved.

また、この発明にかかるガスタービンは、前記動翼の肉厚方向の断面視にて、前記境界部がラウンド形状を有する。   In the gas turbine according to the present invention, the boundary portion has a round shape in a cross-sectional view of the moving blade in the thickness direction.

このガスタービンでは、境界部がスクエア形状を有する構成と比較して、孔の開口部付近における動翼の断面形状（肉厚）が緩やかに変化する。これにより、孔の開口部付近における応力集中が緩和されて動翼の破壊（クリープ亀裂などの発生）が抑制されるので、製品の信頼性が向上する利点がある。   In this gas turbine, the cross-sectional shape (thickness) of the moving blade in the vicinity of the opening of the hole gradually changes as compared with a configuration in which the boundary portion has a square shape. As a result, the stress concentration in the vicinity of the opening of the hole is alleviated and the destruction of the moving blade (occurrence of creep cracks, etc.) is suppressed, so that there is an advantage that the reliability of the product is improved.

また、この発明にかかるガスタービンは、タービンの動翼に冷却通路が形成されると共に前記冷却通路に冷却空気を流通させて前記動翼の冷却を行うガスタービンであって、前記冷却通路が、前記動翼の内部に形成された空洞部と、前記空洞部から前記動翼の長手方向に延在する孔とを含み、且つ、前記動翼の肉厚方向の断面視にて前記空洞部の幅が前記孔との境界部にて減縮されると共に前記境界部がラウンド形状を有することを特徴とする。   The gas turbine according to the present invention is a gas turbine in which a cooling passage is formed in a moving blade of a turbine and cooling air is circulated through the cooling passage to cool the moving blade. A cavity portion formed inside the rotor blade, and a hole extending from the cavity portion in the longitudinal direction of the rotor blade, and the cavity portion in a cross-sectional view in the thickness direction of the rotor blade. The width is reduced at the boundary with the hole, and the boundary has a round shape.

このガスタービンでは、（１）境界部がラウンド形状を有するので、孔の開口部付近における動翼の断面形状が緩やかに変化する。これにより、孔の開口部付近における応力集中が緩和されて動翼の破壊が抑制される。また、（２）空洞部の幅が境界部にて減縮されているので、空洞部の幅が拡張されている構成と比較して、孔の開口部付近における動翼の肉厚が確保される。これにより、動翼が補強されて動翼の破壊がさらに抑制される。これらにより、製品の信頼性がさらに向上する利点がある。   In this gas turbine, (1) since the boundary portion has a round shape, the cross-sectional shape of the moving blade in the vicinity of the opening of the hole gradually changes. Thereby, the stress concentration in the vicinity of the opening of the hole is relaxed and the destruction of the moving blade is suppressed. Also, (2) since the width of the cavity is reduced at the boundary, the thickness of the rotor blade near the opening of the hole is ensured as compared with the configuration in which the width of the cavity is expanded. . Thereby, a moving blade is reinforced and destruction of a moving blade is further suppressed. As a result, there is an advantage that the reliability of the product is further improved.

また、この発明にかかるガスタービンは、前記境界部の前記孔側の壁面が平坦に形成される。   In the gas turbine according to the present invention, the hole-side wall surface of the boundary portion is formed flat.

このガスタービンでは、境界部の孔側の壁面が平坦なので、孔の加工位置がズレた場合にも冷却通路の形状が良好に確保される利点がある。   In this gas turbine, since the wall surface on the hole side of the boundary portion is flat, there is an advantage that the shape of the cooling passage is satisfactorily ensured even when the hole processing position is shifted.

この発明にかかるガスタービンでは、動翼の肉厚方向の断面視にて、空洞部の幅が孔との境界部にて拡張されているので、幅狭な空洞部に対して動翼の長手方向から孔が開けられる構成と比較して、孔の加工が容易である。これにより、加工時の位置ズレによる孔の連通不良などが低減されるので、製品の信頼性が向上する利点がある。   In the gas turbine according to the present invention, the width of the cavity is expanded at the boundary with the hole in the sectional view in the thickness direction of the rotor blade. Compared with a configuration in which a hole is formed from a direction, the hole is easily processed. As a result, poor hole communication due to misalignment during processing is reduced, and thus there is an advantage that the reliability of the product is improved.

以下、この発明につき図面を参照しつつ詳細に説明する。なお、この実施例によりこの発明が限定されるものではない。また、この実施例の構成要素には、当業者が置換可能かつ容易なもの、或いは実質的同一のものが含まれる。また、この実施例に記載された複数の変形例は、当業者自明の範囲内にて任意に組み合わせが可能である。   Hereinafter, the present invention will be described in detail with reference to the drawings. Note that the present invention is not limited to the embodiments. The constituent elements of this embodiment include those that can be easily replaced by those skilled in the art or those that are substantially the same. In addition, a plurality of modifications described in this embodiment can be arbitrarily combined within a range obvious to those skilled in the art.

図１は、この発明の実施例にかかるガスタービンを示す全体構成図である。図２〜図５は、図１に記載したガスタービンの動翼を示す縦断面図（図２）、要部拡大図（図３）、Ａ−Ａ視断面図（図４）およびＢ−Ｂ視断面図（図５）である。図６〜図１１は、図１に記載したガスタービンの変形例を示す説明図である。図１２は、従来のガスタービンの動翼を示す説明図である。   FIG. 1 is an overall configuration diagram showing a gas turbine according to an embodiment of the present invention. 2 to 5 are a longitudinal sectional view (FIG. 2), a main part enlarged view (FIG. 3), a cross-sectional view taken along line AA (FIG. 4), and a line BB showing the moving blade of the gas turbine shown in FIG. 1. FIG. 6 is a sectional view (FIG. 5). 6-11 is explanatory drawing which shows the modification of the gas turbine described in FIG. FIG. 12 is an explanatory view showing a moving blade of a conventional gas turbine.

［ガスタービン］
このガスタービン１は、圧縮機２と、燃焼器３と、タービン４とを有する（図１参照）。圧縮機２は、空気取込口から取り込まれた空気を圧縮して圧縮空気を生成する。燃焼器３は、この圧縮空気に燃料を噴射して高温・高圧の燃焼ガスを発生させる。タービン４は、この燃焼ガスの熱エネルギーをロータ５の回転エネルギーに変換して駆動力を発生させる。そして、この駆動力がロータ５に連結された発電機（図示省略）などに伝達される。 [gas turbine]
The gas turbine 1 includes a compressor 2, a combustor 3, and a turbine 4 (see FIG. 1). The compressor 2 compresses the air taken in from the air intake and generates compressed air. The combustor 3 injects fuel into the compressed air to generate high-temperature and high-pressure combustion gas. The turbine 4 generates the driving force by converting the thermal energy of the combustion gas into the rotational energy of the rotor 5. The driving force is transmitted to a generator (not shown) connected to the rotor 5.

また、タービン４は、複数の動翼４１および複数の静翼４２を有する。これらの動翼４１および静翼４２は、ロータ５の外周かつ軸方向に交互に配列される。そして、これらの動翼４１および静翼４２間を燃料ガスが膨張しつつ通過することにより動翼４１と共にロータ５が回転して駆動力が発生する。なお、このタービン４では、圧縮空気の一部が冷却空気として圧縮機２から抽出され、この冷却空気により動翼４１、静翼４２、プラットフォーム、静翼４２の内側シュラウドおよび外側シュラウドなどが冷却される。   Further, the turbine 4 includes a plurality of moving blades 41 and a plurality of stationary blades 42. The rotor blades 41 and the stationary blades 42 are alternately arranged on the outer periphery and the axial direction of the rotor 5. Then, when the fuel gas expands and passes between the moving blade 41 and the stationary blade 42, the rotor 5 rotates together with the moving blade 41 to generate a driving force. In the turbine 4, a part of the compressed air is extracted as cooling air from the compressor 2, and the moving blade 41, the stationary blade 42, the platform, the inner shroud and the outer shroud of the stationary blade 42 are cooled by the cooling air. The

タービン４の動翼４１は、その基部４１１にてロータ５に取り付けられ、その先端部４１２をロータ５の径方向に向けて配置される（図２参照）。また、動翼４１は、冷却空気を流通させるための冷却通路４１３、４１４を有する。この冷却通路４１３、４１４は、空洞部４１３と、複数の孔（マルチホール）４１４とにより構成される。空洞部４１３は、動翼４１の基部４１１側に形成される。また、空洞部４１３の内部には、複数のピンフィン４１５が形成される。孔４１４は、空洞部４１３から動翼４１の長手方向に延びる細孔であり、動翼４１の先端部４１２側に形成される。   The rotor blade 41 of the turbine 4 is attached to the rotor 5 at its base 411, and the tip portion 412 thereof is arranged in the radial direction of the rotor 5 (see FIG. 2). Further, the moving blade 41 has cooling passages 413 and 414 for circulating cooling air. The cooling passages 413 and 414 include a hollow portion 413 and a plurality of holes (multiholes) 414. The cavity 413 is formed on the base 411 side of the rotor blade 41. In addition, a plurality of pin fins 415 are formed inside the cavity 413. The hole 414 is a fine hole extending from the cavity portion 413 in the longitudinal direction of the rotor blade 41, and is formed on the tip portion 412 side of the rotor blade 41.

かかる構成では、タービン稼働時にて、冷却空気が冷却通路（空洞部４１３および複数の孔４１４）を流通することにより熱交換が行われて動翼４１が冷却される。具体的には、冷却空気が動翼４１の基部４１１側から空洞部４１３に流入し、複数の孔４１４を通って動翼４１の先端部４１２側に流れる。これにより、動翼４１全体の冷却が行われる。また、空洞部４１３内にて冷却空気がピンフィン４１５に衝突することにより、熱交換が促進されて冷却効率が向上する。   In such a configuration, when the turbine is in operation, the cooling air flows through the cooling passages (the hollow portion 413 and the plurality of holes 414), whereby heat exchange is performed and the moving blade 41 is cooled. Specifically, the cooling air flows from the base portion 411 side of the moving blade 41 into the cavity portion 413 and flows through the plurality of holes 414 to the tip portion 412 side of the moving blade 41. Thereby, the moving blade 41 is cooled as a whole. Further, when the cooling air collides with the pin fins 415 in the cavity 413, heat exchange is promoted and the cooling efficiency is improved.

［空洞部と孔との境界部］
また、この動翼４１では、動翼４１の肉厚方向の断面視にて、空洞部４１３の幅ｗが孔４１４との境界部４１６にて拡張される（図５参照）。言い換えると、空洞部４１３と孔４１４との境界部４１６には、空洞部４１３の幅ｗを拡張して成る拡幅部（キャビティ）が形成される。また、この境界部４１６では、孔４１４が配置される範囲（動翼４１の前縁側から後縁側に至る領域）の略全域に渡って、空洞部４１３の幅ｗが拡張される（図４参照）。そして、この拡幅された空洞部４１３（境界部４１６）に対して孔４１４が開口する。 [Boundary between cavity and hole]
Further, in the moving blade 41, the width w of the cavity 413 is expanded at the boundary portion 416 with the hole 414 in a cross-sectional view of the moving blade 41 in the thickness direction (see FIG. 5). In other words, a widened portion (cavity) formed by expanding the width w of the cavity 413 is formed at the boundary 416 between the cavity 413 and the hole 414. Further, in the boundary portion 416, the width w of the cavity portion 413 is expanded over substantially the entire region where the hole 414 is disposed (region extending from the front edge side to the rear edge side of the moving blade 41) (see FIG. 4). ). And the hole 414 opens with respect to this widened cavity part 413 (boundary part 416).

また、これらの空洞部４１３および孔４１４は、以下の工程により成形加工される。まず、空洞部４１３が動翼４１本体と共に一時に鋳造される。このとき、境界部４１６が空洞部４１３の一部として併せて成形される。次に、孔４１４が切削加工（ドリル加工）により動翼４１の先端部４１２側から空洞部４１３に向かって開けられる。   Further, the cavity 413 and the hole 414 are formed by the following process. First, the cavity 413 is cast at a time together with the rotor blade 41 main body. At this time, the boundary portion 416 is molded together as a part of the cavity portion 413. Next, the hole 414 is opened from the tip 412 side of the rotor blade 41 toward the cavity 413 by cutting (drilling).

かかる構成では、動翼４１の肉厚方向の断面視にて、空洞部４１３の幅ｗが孔４１４との境界部４１６にて拡張されている（ｗ１＜ｗ２）ので、幅狭な（拡幅されていない）空洞部に対して動翼４１の長手方向から孔が開けられる構成（図１２参照）と比較して、孔４１４の加工が容易である。これにより、加工時の位置ズレによる孔４１４の連通不良などが低減されるので、製品の信頼性が向上する利点がある。   In such a configuration, the width w of the cavity portion 413 is expanded at the boundary portion 416 with the hole 414 (w1 <w2) in the cross-sectional view in the thickness direction of the rotor blade 41, so that the width is narrow (widened). The hole 414 can be easily processed as compared with a configuration (see FIG. 12) in which a hole is formed in the cavity from the longitudinal direction of the rotor blade 41. As a result, poor communication of the holes 414 due to misalignment during processing is reduced, which has the advantage of improving product reliability.

なお、このガスタービン１では、冷却通路の空洞部４１３および境界部４１６が隣接かつ連続する空間であるため、便宜上、境界部４１６が空洞部４１３の一部を構成するものとして説明した。しかし、これに限らず、境界部４１６と空洞部４１３とが別個の構成要素として扱われても良い。すなわち、従来のガスタービンが有する空洞部４１３に対して新たに境界部４１６が形成されたものと考えても良い。   In the gas turbine 1, since the cavity 413 and the boundary 416 of the cooling passage are adjacent and continuous spaces, the boundary 416 is described as a part of the cavity 413 for convenience. However, the present invention is not limited to this, and the boundary portion 416 and the hollow portion 413 may be treated as separate components. That is, it may be considered that a boundary part 416 is newly formed with respect to the cavity part 413 of the conventional gas turbine.

［変形例１］
また、このガスタービン１では、動翼４１の肉厚方向の断面視にて、境界部４１６がラウンド形状を有することが好ましい（図６参照）。具体的には、動翼４１の肉厚方向の断面視にて、空洞部４１３の壁面のうち孔４１４の開口部を有する側の壁面が、孔４１４側に凸となる（空洞部４１３側に中心を有する）略円弧形状を有することが好ましい。例えば、動翼４１の肉厚方向の断面視にて、空洞部４１３が孔４１４との境界部４１６で略円形状あるいは略楕円形状に膨らんで拡幅された形状を有する。 [Modification 1]
Moreover, in this gas turbine 1, it is preferable that the boundary part 416 has a round shape in sectional view of the moving blade 41 in the thickness direction (see FIG. 6). Specifically, in the cross-sectional view of the moving blade 41 in the thickness direction, the wall surface on the side having the opening of the hole 414 out of the wall surface of the cavity portion 413 is convex toward the hole 414 side (to the cavity portion 413 side) It preferably has a substantially arc shape (having a center). For example, in a cross-sectional view of the moving blade 41 in the thickness direction, the cavity 413 has a shape that is expanded and widened in a substantially circular shape or a substantially elliptical shape at the boundary portion 416 with the hole 414.

かかる構成では、境界部４１６がスクエア形状を有する構成（図５参照）と比較して、孔４１４の開口部付近における動翼４１の断面形状（肉厚）が緩やかに変化する（図６参照）。すなわち、孔４１４の開口部付近では、境界部４１６がラウンド形状を有するので、境界部４１６により区画された動翼４１の前縁側および後縁側の肉厚が冷却通路の上流側から下流側（孔４１４側）に向かうに連れて徐々に増加する。また、動翼４１の肉厚方向の断面視にて、孔４１４の開口部の縁が鈍角となる（境界部４１６と孔４１４とが鈍角にて緩やかに接続される）。これにより、孔４１４の開口部付近における応力集中が緩和されて動翼４１の破壊（クリープ亀裂などの発生）が抑制されるので、製品の信頼性が向上する利点がある。   In such a configuration, the cross-sectional shape (thickness) of the moving blade 41 in the vicinity of the opening of the hole 414 changes more slowly than the configuration in which the boundary portion 416 has a square shape (see FIG. 5) (see FIG. 6). . That is, since the boundary portion 416 has a round shape in the vicinity of the opening portion of the hole 414, the thickness of the front edge side and the rear edge side of the moving blade 41 partitioned by the boundary portion 416 is reduced from the upstream side to the downstream side (hole 414 side) and gradually increase. Moreover, the edge of the opening part of the hole 414 becomes an obtuse angle in the cross-sectional view of the moving blade 41 in the thickness direction (the boundary part 416 and the hole 414 are gently connected at an obtuse angle). Thereby, the stress concentration in the vicinity of the opening of the hole 414 is alleviated and the destruction of the moving blade 41 (occurrence of creep cracks or the like) is suppressed, so that there is an advantage that the reliability of the product is improved.

［変形例２］
また、変形例１の構成は、動翼４１の肉厚方向の断面視にて空洞部４１３の幅ｗが孔４１４との境界部４１６にて拡張される構成（図６参照）に限らず、空洞部４１３の幅ｗが孔４１４との境界部４１６にて徐々に減少する構成に対して採用されても良い（図７参照）。すなわち、この変形例２のガスタービン１では、動翼４１の肉厚方向の断面視にて空洞部４１３の幅ｗが孔４１４との境界部４１６にて減縮され、且つ、境界部４１６がラウンド形状を有する。 [Modification 2]
The configuration of Modification 1 is not limited to the configuration in which the width w of the cavity 413 is expanded at the boundary portion 416 with the hole 414 in the cross-sectional view in the thickness direction of the rotor blade 41 (see FIG. 6). The width w of the cavity 413 may be adopted for a configuration in which the width w gradually decreases at the boundary 416 with the hole 414 (see FIG. 7). That is, in the gas turbine 1 of the second modification, the width w of the cavity 413 is reduced at the boundary 416 with the hole 414 in the cross-sectional view of the moving blade 41 in the thickness direction, and the boundary 416 is round. Has a shape.

かかる構成では、（１）境界部４１６がラウンド形状を有するので、孔４１４の開口部付近における動翼４１の断面形状が緩やかに変化する。これにより、孔４１４の開口部付近における応力集中が緩和されて動翼４１の破壊が抑制される。また、（２）空洞部４１３の幅ｗが境界部４１６にて減縮されているので、空洞部４１３の幅ｗが拡張されている構成（図５参照）と比較して、孔４１４の開口部付近における動翼４１の肉厚が確保される。これにより、動翼４１が補強されて動翼４１の破壊がさらに抑制される。これらにより、製品の信頼性がさらに向上する利点がある。   In this configuration, (1) since the boundary portion 416 has a round shape, the cross-sectional shape of the moving blade 41 in the vicinity of the opening portion of the hole 414 changes gently. Thereby, the stress concentration in the vicinity of the opening of the hole 414 is relaxed, and the destruction of the moving blade 41 is suppressed. Further, (2) since the width w of the cavity 413 is reduced at the boundary 416, the opening of the hole 414 is compared with the configuration in which the width w of the cavity 413 is expanded (see FIG. 5). The thickness of the moving blade 41 in the vicinity is secured. Thereby, the moving blade 41 is reinforced and the destruction of the moving blade 41 is further suppressed. As a result, there is an advantage that the reliability of the product is further improved.

［変形例３］
また、上記のガスタービン１では、境界部４１６の孔４１４側の壁面（下流側の壁面）が平坦に形成されることが好ましい（図８参照）。すなわち、この変形例３のガスタービン１では、境界部４１６がラウンド形状を有し、且つ、境界部４１６の孔４１４側の壁面が平坦に形成される。具体的には、動翼４１の肉厚方向の断面視にて、境界部４１６の孔４１４側の壁面が孔４１４の延在方向（冷却空気の流れ方向）に対して略垂直となるように形成される。 [Modification 3]
Moreover, in said gas turbine 1, it is preferable that the wall surface (downstream wall surface) by the side of the hole 414 of the boundary part 416 is formed flatly (refer FIG. 8). That is, in the gas turbine 1 of the third modification, the boundary portion 416 has a round shape, and the wall surface on the hole 414 side of the boundary portion 416 is formed flat. Specifically, in the cross-sectional view of the moving blade 41 in the thickness direction, the wall surface on the hole 414 side of the boundary portion 416 is substantially perpendicular to the extending direction of the hole 414 (cooling air flow direction). It is formed.

かかる構成では、境界部４１６の孔４１４側の壁面が平坦なので、孔４１４の加工位置がズレた場合にも冷却通路の形状が良好に確保される利点がある。例えば、動翼４１の肉厚方向の断面視にて境界部４１６がラウンド形状を有する構成では、境界部４１６孔４１４側の壁面がシャープエッジ（ヘアピン状に湾曲した形状）Ｓとなる（図１０参照）。すると、孔４１４の加工位置がズレたときに、このシャープエッジＳが残って冷却通路の形状に不備が生じる。この点において、この変形例３にかかるガスタービン１では、孔４１４の加工位置がズレた場合にも、かかるシャープエッジＳが生じない点で好ましい（図９参照）。なお、上記の構成は、図６に記載した構成においても同様に適用可能である（図１１参照）。   In such a configuration, since the wall surface on the hole 414 side of the boundary portion 416 is flat, there is an advantage that the shape of the cooling passage is ensured well even when the processing position of the hole 414 is shifted. For example, in a configuration in which the boundary portion 416 has a round shape in a cross-sectional view in the thickness direction of the moving blade 41, the wall surface on the boundary portion 416 hole 414 side is a sharp edge (a shape curved in a hairpin shape) S (FIG. 10). reference). Then, when the machining position of the hole 414 is shifted, this sharp edge S remains, and the shape of the cooling passage is deficient. In this respect, the gas turbine 1 according to the third modification is preferable in that the sharp edge S does not occur even when the processing position of the hole 414 is shifted (see FIG. 9). Note that the above configuration is also applicable to the configuration illustrated in FIG. 6 (see FIG. 11).

以上のように、本発明にかかるガスタービンは、製品の信頼性を向上させ得る点で有用である。   As described above, the gas turbine according to the present invention is useful in that the reliability of the product can be improved.

この発明の実施例にかかるガスタービンを示す全体構成図である。1 is an overall configuration diagram showing a gas turbine according to an embodiment of the present invention. 図１に記載したガスタービンの動翼を示す縦断面図である。It is a longitudinal cross-sectional view which shows the moving blade of the gas turbine described in FIG. 図１に記載したガスタービンの動翼を示す要部拡大図である。It is a principal part enlarged view which shows the moving blade of the gas turbine described in FIG. 図１に記載したガスタービンの動翼を示すＡ−Ａ視断面図である。FIG. 2 is a cross-sectional view taken along line AA showing a moving blade of the gas turbine described in FIG. 1. 図１に記載したガスタービンの動翼を示すＢ−Ｂ視断面図である。It is a BB sectional view showing the moving blade of the gas turbine described in FIG. 図１に記載したガスタービンの変形例を示す説明図である。It is explanatory drawing which shows the modification of the gas turbine described in FIG. 図１に記載したガスタービンの変形例を示す説明図である。It is explanatory drawing which shows the modification of the gas turbine described in FIG. 図１に記載したガスタービンの変形例を示す説明図である。It is explanatory drawing which shows the modification of the gas turbine described in FIG. 図１に記載したガスタービンの変形例を示す説明図である。It is explanatory drawing which shows the modification of the gas turbine described in FIG. 図１に記載したガスタービンの変形例を示す説明図である。It is explanatory drawing which shows the modification of the gas turbine described in FIG. 図１に記載したガスタービンの変形例を示す説明図である。It is explanatory drawing which shows the modification of the gas turbine described in FIG. 従来のガスタービンの動翼を示す説明図である。It is explanatory drawing which shows the moving blade of the conventional gas turbine.

符号の説明Explanation of symbols

１ ガスタービン
２ 圧縮機
３ 燃焼器
４ タービン
５ ロータ
４１ 動翼
４２ 静翼
４１１ 基部
４１２ 先端部
４１３ 空洞部
４１４ 孔
４１５ ピンフィン
４１６ 境界部 DESCRIPTION OF SYMBOLS 1 Gas turbine 2 Compressor 3 Combustor 4 Turbine 5 Rotor 41 Rotor blade 42 Static blade 411 Base part 412 Tip part 413 Cavity part 414 Hole 415 Pin fin 416 Boundary part

Claims (4)

タービンの動翼に冷却通路が形成されると共に前記冷却通路に冷却空気を流通させて前記動翼の冷却を行うガスタービンであって、
前記冷却通路が、前記動翼の内部に形成された空洞部と、前記空洞部から前記動翼の長手方向に延在する孔とを含み、且つ、前記動翼の肉厚方向の断面視にて、前記空洞部の幅が前記孔との境界部にて拡張されていることを特徴とするガスタービン。 A gas turbine in which a cooling passage is formed in a moving blade of a turbine and cooling air is circulated through the cooling passage to cool the moving blade,
The cooling passage includes a cavity formed inside the blade, and a hole extending from the cavity in the longitudinal direction of the blade, and in a cross-sectional view in the thickness direction of the blade. The gas turbine is characterized in that the width of the cavity is expanded at the boundary with the hole. 前記動翼の肉厚方向の断面視にて、前記境界部がラウンド形状を有する請求項１に記載のガスタービン。   The gas turbine according to claim 1, wherein the boundary portion has a round shape in a cross-sectional view in the thickness direction of the moving blade. タービンの動翼に冷却通路が形成されると共に前記冷却通路に冷却空気を流通させて前記動翼の冷却を行うガスタービンであって、
前記冷却通路が、前記動翼の内部に形成された空洞部と、前記空洞部から前記動翼の長手方向に延在する孔とを含み、且つ、前記動翼の肉厚方向の断面視にて前記空洞部の幅が前記孔との境界部にて減縮されると共に前記境界部がラウンド形状を有することを特徴とするガスタービン。 A gas turbine in which a cooling passage is formed in a moving blade of a turbine and cooling air is circulated through the cooling passage to cool the moving blade,
The cooling passage includes a cavity formed inside the blade, and a hole extending from the cavity in the longitudinal direction of the blade, and in a cross-sectional view in the thickness direction of the blade. The width of the cavity is reduced at the boundary with the hole, and the boundary has a round shape. 前記境界部の前記孔側の壁面が平坦に形成される請求項２または３に記載のガスタービン。   The gas turbine according to claim 2, wherein a wall surface on the hole side of the boundary portion is formed flat.
JP2006030202A 2006-02-07 2006-02-07 Gas turbine Withdrawn JP2007211618A (en)

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JP2009066659A (en) * 2007-09-10 2009-04-02 Snecma Use of activating flux for tig welding of metal parts
JP2009167934A (en) * 2008-01-17 2009-07-30 Mitsubishi Heavy Ind Ltd Gas turbine moving blade and gas turbine
US20110217181A1 (en) * 2010-03-03 2011-09-08 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
US20110217180A1 (en) * 2010-03-03 2011-09-08 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
JP2011202656A (en) * 2010-03-25 2011-10-13 General Electric Co <Ge> Airfoil cooling hole flag region

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JP2009066659A (en) * 2007-09-10 2009-04-02 Snecma Use of activating flux for tig welding of metal parts
JP2009167934A (en) * 2008-01-17 2009-07-30 Mitsubishi Heavy Ind Ltd Gas turbine moving blade and gas turbine
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US8827646B2 (en) 2010-03-03 2014-09-09 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
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US20110217181A1 (en) * 2010-03-03 2011-09-08 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
JP5200189B2 (en) * 2010-03-03 2013-05-15 三菱重工業株式会社 Gas turbine rotor blade, method for manufacturing the same, and gas turbine using the rotor blade
US8506251B2 (en) 2010-03-03 2013-08-13 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
US20130209271A1 (en) * 2010-03-03 2013-08-15 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
US20110217180A1 (en) * 2010-03-03 2011-09-08 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
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EP2543821B1 (en) * 2010-03-03 2015-11-04 Mitsubishi Hitachi Power Systems, Ltd. Rotor blade for gas turbine, method for manufacturing same, and gas turbine using rotor blade
JP2011202656A (en) * 2010-03-25 2011-10-13 General Electric Co <Ge> Airfoil cooling hole flag region

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