JPS6380004A - Gas turbine stator blade - Google Patents
Gas turbine stator bladeInfo
- Publication number
- JPS6380004A JPS6380004A JP22199486A JP22199486A JPS6380004A JP S6380004 A JPS6380004 A JP S6380004A JP 22199486 A JP22199486 A JP 22199486A JP 22199486 A JP22199486 A JP 22199486A JP S6380004 A JPS6380004 A JP S6380004A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- core plug
- gas turbine
- cooling
- stator blade
- 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
- 238000005192 partition Methods 0.000 claims abstract description 28
- 238000001816 cooling Methods 0.000 abstract description 66
- 239000007789 gas Substances 0.000 description 20
- 230000000694 effects Effects 0.000 description 7
- 239000002184 metal Substances 0.000 description 7
- 230000008646 thermal stress Effects 0.000 description 6
- 230000001737 promoting effect Effects 0.000 description 2
- 239000000112 cooling gas Substances 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 238000007599 discharging Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 238000002955 isolation Methods 0.000 description 1
- 238000000465 moulding Methods 0.000 description 1
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【発明の詳細な説明】
［産業上の利用分野〕
本発明は、冷却構造を改良したガスタービン静翼に関す
るものである。DETAILED DESCRIPTION OF THE INVENTION [Field of Industrial Application] The present invention relates to a gas turbine stationary blade with an improved cooling structure.
ガスタービン静翼の冷却においては、インビンリメント
冷却１強制対流冷却及びフィルム冷却のいずれかを組合
せて行うのが一般的である。In cooling gas turbine stationary blades, it is common to use a combination of invincment cooling, forced convection cooling, and film cooling.
第６１！ｌは従来のガスタービン静翼の冷却構造の１例
を示す斜視図であって、この従来例はインピンジメント
冷却と強制対流冷却とを併用した構造である。61st! 1 is a perspective view showing an example of a conventional cooling structure for gas turbine stator blades, and this conventional example has a structure that uses both impingement cooling and forced convection cooling.
円弧状をなす静翼セグメント（図示せず）のチップ側エ
ンドウオール（図示せず）と、ハブ側エンドウオール（
図示せず）との間に中空の静翼本体１が固着される。該
中空静翼本体１の空洞内にはコアプラグ２が、翼内壁面
４に対向離間して設置され該コアプラグ２にはインピン
ジ孔３が配列されている。５は、コアプラグ２と翼内壁
面４とに挟まれた空間である。The tip-side end wall (not shown) of the arc-shaped stator vane segment (not shown) and the hub-side end wall (
(not shown), a hollow stator blade main body 1 is fixed between the two. In the cavity of the hollow stator blade body 1, a core plug 2 is installed facing and spaced apart from the inner wall surface 4 of the blade, and impingement holes 3 are arranged in the core plug 2. 5 is a space sandwiched between the core plug 2 and the blade inner wall surface 4.
静翼本体１の後縁部７には、冷却空気を主流の中へ放出
するための後縁流路８が設けられている。A trailing edge passage 8 is provided at the trailing edge 7 of the stator blade body 1 for discharging cooling air into the mainstream.
コアプラグ２の内部空間６に導入された低温の冷却空気
は、インピンジ孔３を通り噴流となって翼内壁４に衝突
し、インビンジメント冷却を行う。The low-temperature cooling air introduced into the internal space 6 of the core plug 2 passes through the impingement hole 3, becomes a jet stream, and collides with the blade inner wall 4, thereby performing impingement cooling.
その後、冷却空気はコアプラグ２と翼内壁４とで形成さ
れる空間５を通路として、次第に流量を増加させながら
後縁側へコード方向に、流れ、後縁流路８から後縁部７
の強制対流冷却を行い、主流中へと排出される。なお、
この種の装置として関連するものには例えば特開昭５３
−９０５０９号等が挙げられる。Thereafter, the cooling air flows through the space 5 formed by the core plug 2 and the blade inner wall 4 in the chord direction toward the trailing edge while gradually increasing the flow rate, and flows from the trailing edge flow path 8 to the trailing edge 7.
is subjected to forced convection cooling and discharged into the mainstream. In addition,
Related devices of this type include, for example, Japanese Unexamined Patent Publication No. 53
-90509 etc. are mentioned.
前記の従来技術においては、冷却空気の１部はインピン
ジ孔３を通過し、噴流となって翼内壁面４に衝突し、空
間５を冷却空気通路として、流量を増加させながら後縁
流路８に至る。この空間５を流れる空気流は、インピン
ジ孔３からの噴出空気流に対してほぼ直角に交わる（ク
ロスフローする。）
この現象により、インビンジメント冷却が妨げられて冷
却効率が低下する。In the prior art described above, a part of the cooling air passes through the impingement hole 3, becomes a jet, collides with the blade inner wall surface 4, and uses the space 5 as a cooling air passage to increase the flow rate and flow into the trailing edge flow passage 8. leading to. The airflow flowing through this space 5 intersects (crossflows) the airflow ejected from the impingement hole 3 at a substantially right angle (crossflow).This phenomenon impedes impingement cooling and reduces cooling efficiency.
一般に、主流ガスの半径方向温度分布は、翼の中央部を
ピークとする二次曲線状となっており。Generally, the radial temperature distribution of the mainstream gas has a quadratic curve shape with a peak at the center of the blade.
主流ガス温度は翼の中央部で高く、外径側部分及び内径
側部分（本第６図において上端付近及び下端付近）では
比較的低い。The mainstream gas temperature is high at the center of the blade, and relatively low at the outer and inner diameter portions (near the upper end and lower end in FIG. 6).
インピンジ孔３の配列ピッチはＳ／ｄ≦２０の範囲内で
（ｄは孔径、Ｓはピッチ寸法、単位ｎａ）外部熱負荷に
応じて定められ、翼の中央部では密に、外径側及び内径
側部分では粗として、翼メタル温度が均一となるように
冷却設計される。従って、外部熱負荷条件の厳しい翼中
央部で多量の冷却空気を消費することになり、この中央
部でのクロスフローによる影響が問題となる。The arrangement pitch of the impingement holes 3 is determined within the range of S/d≦20 (d is the hole diameter, S is the pitch dimension, unit: na) according to the external heat load, and is dense in the center of the blade, and densely arranged on the outer diameter side and Cooling is designed to be rough on the inner diameter side so that the temperature of the blade metal is uniform. Therefore, a large amount of cooling air is consumed in the center of the blade where external heat load conditions are severe, and the influence of crossflow in this center becomes a problem.
前記インピンジ孔の配列ピッチを、Ｓ／ｄ≦２０ｍの範
囲内で最も疎に設定しても、翼外径側付近及び翼内後側
付近は、外部熱負荷が小さいので翼メタル温度は適冷気
味となる。このため、静翼内での温度分布が不均一にな
り、大きい熱応力を生じてタービン強度の信頼性が低下
する。Even if the arrangement pitch of the impingement holes is set to be the sparsest within the range of S/d≦20m, the external heat load is small near the outer diameter side of the blade and near the inner rear side of the blade, so the temperature of the blade metal remains appropriately cool. It feels a little weird. As a result, the temperature distribution within the stator blade becomes non-uniform, causing large thermal stress and reducing the reliability of the turbine strength.
本発明は上記の事情に鑑みて為されたもので。The present invention has been made in view of the above circumstances.
翼中央部での冷却効率を高め、所要冷却空気量の低減（
必然的に熱効率を上昇させる）、及び、温度分布の均一
化（当然に熱応力を低減させる）を図ったガスタービン
静翼を提供するものである。Increased cooling efficiency in the center of the blade, reducing the amount of cooling air required (
The purpose of the present invention is to provide a gas turbine stationary blade that achieves a uniform temperature distribution (which naturally increases thermal efficiency) and a uniform temperature distribution (which naturally reduces thermal stress).
上記目的を達成するために創作した本発明のガスタービ
ン静翼について、先ず、その基本的原理を略述すると次
の如くである。First, the basic principle of the gas turbine stationary blade of the present invention created to achieve the above object will be briefly described as follows.
前記目的は、翼内壁とコアプラグで形成される空間部の
翼中央部位置に、半径方向に延びる隔壁をインピンジ孔
とインピンジ孔の間にコード方向に複数個配置して前記
空間部を区画することにより達成される。The purpose is to divide the space by arranging a plurality of radially extending partition walls in the chord direction between the impingement holes at the center of the space in the space formed by the inner wall of the wing and the core plug. This is achieved by
又、外部熱負荷の大きい翼中央部をインピンジメント冷
却、比較的、外部熱負荷の小さい翼の外径側及び内径側
部分を強制対流冷却すると、より好適な冷却が行われる
。Further, more suitable cooling can be achieved by impingement cooling the center portion of the blade, which has a large external heat load, and by forced convection cooling the outer diameter side and inner diameter side portions of the blade, which have a relatively small external heat load.
上記の原理に基いて、これを実際のガスタービン静翼の
適用するための具体的な構成として１本発明は、静翼本
体内に形成された空洞の内壁面に対向離間せしめてコア
プラグを固定すると共に。Based on the above principle, the present invention has a concrete configuration for applying this to an actual gas turbine stator blade.The present invention has a core plug fixed to the inner wall surface of a cavity formed in the stator blade body so as to be spaced apart from each other. Along with.
該コアプラグにインピンジ孔を配列すると共に。While arranging impingement holes in the core plug.
前記空洞と連通せしめて静翼後縁流路を設けたガスター
ビン静翼において、前記コアプラグと静翼本体の内壁面
とに挟まれた空間をコード方向に仕切る複数の隔壁を設
けたことを特徴とする。The gas turbine stator blade is provided with a stator blade trailing edge flow path communicating with the cavity, characterized in that a plurality of partition walls are provided to partition a space sandwiched between the core plug and the inner wall surface of the stator blade body in the chord direction. shall be.
〔作用〕
上記のように構成した静翼においては、翼中央部でのイ
ンピンジ孔を通り翼内壁に衝突した冷却空気は、半径方
向に延びる隔壁により区画された空間部を通路として半
径方向に流れ、隔壁の途絶えた翼の外径側及び内径側部
分でコード方向に流れを変えて後縁流路に到達する。そ
れによって。[Operation] In the stationary blade configured as described above, the cooling air that has passed through the impingement hole in the center of the blade and collided with the inner wall of the blade flows in the radial direction through the space partitioned by the partition wall extending in the radial direction. The flow changes in the chord direction at the outer diameter side and inner diameter side portions of the blade where the partition wall is interrupted, and reaches the trailing edge flow path. Thereby.
翼中央部でのコード方向のクロスフローの影響を解消し
冷却効率を向上する。更に、全流量が通過する翼の外径
側及び内径側部分でのクロスフローの影響を大きくシ、
翼の外径側及び内径側部分ててのインビンジメント冷却
の効果を抑制することができる。Improves cooling efficiency by eliminating the influence of cross flow in the chord direction at the center of the blade. Furthermore, the effect of cross flow on the outer diameter side and inner diameter side of the blade through which the entire flow passes is greatly reduced.
The effect of impingement cooling on the outer diameter side and the inner diameter side portion of the blade can be suppressed.
以下、本発明の第１の実施例を第１図及び第２図により
説明する。A first embodiment of the present invention will be described below with reference to FIGS. 1 and 2.
第１図は本実施例の部分断面斜視図、第２図はそのＡ−
Ａ断面図である。Fig. 1 is a partial cross-sectional perspective view of this embodiment, and Fig. 2 is its A-
It is an A sectional view.
本実施例は第６図の従来例に本発明を適用して改良した
１例であって、第６図と同一の図面参照番号を付したも
のは前記従来例におけると同様乃至は類似の構成部分で
ある。This embodiment is an example in which the present invention is applied and improved to the conventional example shown in FIG. 6, and the same drawing reference numbers as in FIG. It is a part.
本例の静翼本体１は、翼内壁４．とコアプラグ２で形成
される空間５の翼中央部をコアプラグ２にロー付けされ
た半径方向に同一長さの円柱状の隔壁９で軸方向に区画
している。しかるに、本発明のガスタービン静翼におい
て、ガスタービンの運転により静翼を冷却する場合、翼
中央部ではコアプラグ２の内部空間６に導入された冷却
空気は。The stator blade main body 1 of this example has a blade inner wall 4. The blade center portion of the space 5 formed by the core plug 2 and the core plug 2 is partitioned in the axial direction by a cylindrical partition wall 9 having the same length in the radial direction and brazed to the core plug 2. However, in the gas turbine stationary blade of the present invention, when the stationary blade is cooled by operation of the gas turbine, the cooling air introduced into the internal space 6 of the core plug 2 at the center of the blade.
コアプラグ２に、密に配列されたインピンジ孔３を通り
噴流となって、翼内壁４に衝突しインピンジメント冷却
を行う、しかるのち、隔壁９により区画された空間５を
、冷却空気通路としてそれぞれ、翼の外径側及び内径側
へと半径方向に流れ、外径側及び内径側の隔壁９の途絶
えた所でコード方向に流れを変えて後縁部７の後縁流路
８に到り。The core plug 2 passes through impingement holes 3 arranged densely, becomes a jet, collides with the blade inner wall 4, and performs impingement cooling.Then, the space 5 divided by the partition wall 9 is used as a cooling air passage, respectively. It flows in the radial direction toward the outer diameter side and the inner diameter side of the blade, and at the point where the partition walls 9 on the outer diameter side and the inner diameter side are interrupted, the flow changes in the chord direction and reaches the trailing edge flow path 8 of the trailing edge portion 7.
後縁部７を強制対流冷却しながら主流中へと排出される
。The trailing edge 7 is discharged into the mainstream while being cooled by forced convection.
本発明において外径側とはガスタービンの外径側の意（
本第１図において上方）であり、内径側゛とはガスター
ビンの内径側（図の下方）の意である。In the present invention, the outer diameter side means the outer diameter side of the gas turbine (
1), and the term "inner diameter side" means the inner diameter side (lower part in the figure) of the gas turbine.
一方、翼の外径側及び内径側部分では、粗に配列された
インピンジ孔３を介し冷却空気は同様にインビンジメン
ト冷却後、直接、コード方向へ流れて同様に後縁流路８
から主流中へ排出される。On the other hand, in the outer diameter side and inner diameter side portions of the blade, the cooling air similarly flows directly in the chord direction after impingement cooling through the roughly arranged impingement holes 3, and similarly flows into the trailing edge flow path 8.
is discharged into the mainstream.
従って、外部熱負荷の大きい翼中央部に対しては、コー
ド方向でのクロスインピンジメントフローの影響を、−
切、受けることなく冷却ができるので冷却効率が向上し
、従来に比べ、冷却空気の消費量を低減できる。尚、隔
壁９により区画された空間５の半径方向でのクロスフロ
ーの影響を考慮する必要がある。インピンジ孔を通過し
た冷却空気は、外径側及び内径側へ流れるので、空間５
を通過する流量自体、半減されている。しかも、前述し
たように主流ガスの半径方向温度分布は二次曲線状であ
り、半径方向に外部熱負荷は緩和されるので真中央部間
での外径側及び内径側部分でクロスフローによる冷却効
率の低下があっても、その影響は小さく問題のないレベ
ルである。Therefore, for the center of the blade, which has a large external heat load, the influence of cross impingement flow in the chord direction can be reduced to -
Cooling can be done without being cut or exposed, improving cooling efficiency and reducing cooling air consumption compared to conventional systems. Note that it is necessary to consider the influence of cross flow in the radial direction of the space 5 partitioned by the partition wall 9. The cooling air that has passed through the impingement hole flows to the outer diameter side and the inner diameter side, so the space 5
The flow rate passing through the tube itself has been halved. Moreover, as mentioned above, the radial temperature distribution of the mainstream gas is quadratic, and the external heat load is alleviated in the radial direction, so cooling is achieved by cross flow on the outer and inner diameter sides between the true centers. Even if there is a decrease in efficiency, the effect is small and at a level that poses no problem.
翼の外径側及び内径側部分に対しては、翼中央部を冷却
した空気が流れこむためクロスフローによる影響が大き
くなる。Since the air that has cooled the center of the blade flows into the outer diameter side and inner diameter side portions of the blade, the influence of crossflow becomes large.
しかし、翼の外径側及び内径側部分では主流ガス温度が
低く、しかも、前述したようにインピンジ孔３の配列ピ
ッチの制約から、メタル温度が翼中央部に比べ低下する
傾向ある。従って、翼の外径側及び内径側部分でのイン
ピンジメント冷却の冷却効果をクロスフローの影響によ
って抑制するので、翼全体にわたっての温度不均一を低
減し熱応力を低下できる。However, the mainstream gas temperature is low in the outer diameter side and inner diameter side portions of the blade, and furthermore, as described above, due to the restriction on the arrangement pitch of the impingement holes 3, the metal temperature tends to be lower than that in the center portion of the blade. Therefore, the cooling effect of impingement cooling on the outer diameter side and the inner diameter side portion of the blade is suppressed by the influence of the cross flow, so that temperature non-uniformity over the entire blade can be reduced and thermal stress can be reduced.
尚１本実施例（第１図、第２図）では１円柱状の隔壁９
をコアプラグ２にロー付けして設けた場合について述べ
たが、例えば第３図に示すように。In this embodiment (FIGS. 1 and 2), one cylindrical partition wall 9 is used.
The case where the core plug 2 is brazed to the core plug 2 has been described, for example, as shown in FIG.
隔壁９を翼内壁４側への一体鋳造部材として設ける場合
や、逆に第４図に示すように隔壁９をコアプラグ２を成
形して一体物として設置しても適用できることは勿論で
ある。更に、上記隔壁９は第２図乃至第４図に示した断
面形状に拘束されないのは勿論である。It goes without saying that the partition wall 9 may be provided as an integrally cast member on the side of the blade inner wall 4, or conversely, the partition wall 9 may be installed as an integral part by molding the core plug 2 as shown in FIG. Furthermore, it goes without saying that the partition wall 9 is not limited to the cross-sectional shape shown in FIGS. 2 to 4.
本実施例によれば、インピンジメント冷却において翼中
央部での冷却効率を向上させ、所要冷却空気流量を低減
できる。更に、翼全体にわたっての温度不均一を小さく
シ、熱応力を緩和することができる。According to this embodiment, in impingement cooling, the cooling efficiency at the blade center can be improved and the required cooling air flow rate can be reduced. Furthermore, temperature non-uniformity over the entire blade can be reduced and thermal stress can be alleviated.
第５図は１本発明の第２の実施例を示している図におい
て、第１図ないし第６図と同符号のものは同じもの、も
しくは相当するものを示している。FIG. 5 shows a second embodiment of the present invention, in which the same reference numerals as in FIGS. 1 to 6 indicate the same or equivalent components.
前記実施例では、隔壁９の半径方向長さを同一として、
コード方向に配列したが、この実施例では、前縁側から
後縁側にかけ隔壁９が次第に短くなるように配列されて
いる。更に、前記実施例では、静翼の全面にわたりイン
ピンジメント冷却を行っているが、この実施例では、前
縁を除き半径方向に対し隔壁９で囲まれた空間のみをイ
ンピンジメント冷却として、囲まれない空間部を強制対
流冷却としている。In the embodiment, the radial length of the partition wall 9 is the same,
Although the partition walls 9 are arranged in the cord direction, in this embodiment, the partition walls 9 are arranged so as to become gradually shorter from the leading edge side to the trailing edge side. Furthermore, in the embodiment described above, impingement cooling is performed over the entire surface of the stationary blade, but in this embodiment, impingement cooling is performed only in the space surrounded by the partition wall 9 in the radial direction, excluding the leading edge. Forced convection cooling is applied to the space where there is no space.
翼中央部でのインビンジメント冷却後の冷却空気は、隔
壁９により仕切られた空間５を通路として内外半径方向
に流れ、翼の外径側及び内径側部分でコード方向に流れ
を変える。この翼の外径側及び内径側部分の空間５を通
過する冷却空気は、前縁側から後縁側に流れるにしたが
い１次第に流量を増加させる。隔壁９によって囲まれな
い部分の翼内壁４には、突起状の乱流促進体１０が設け
られており、前記冷却空気は乱流となって、この間を強
制対流冷却しながら後縁流路８に到る。The cooling air after impingement cooling at the center of the blade flows in the inner and outer radial directions through the space 5 partitioned by the partition wall 9 as a passage, and changes its flow in the chord direction at the outer diameter side and the inner diameter side of the blade. The cooling air passing through the spaces 5 on the outer diameter side and the inner diameter side of the blade gradually increases in flow rate as it flows from the leading edge side to the trailing edge side. A protruding turbulence promoting body 10 is provided on the blade inner wall 4 in a portion not surrounded by the partition wall 9, and the cooling air becomes a turbulent flow, and the cooling air flows through the trailing edge flow path 8 while being cooled by forced convection. reach.
本実施例では、全冷却空気流量を熱的条件の厳しい翼中
央部のインピンジメント冷却に使用できるので、主流ガ
ス量と冷却空気流量の割合である冷却空気流量比が、第
１の実施例と同一とすれば翼中央部のメタル温度を第１
の実施例よりも低下できる。更に、翼の外径側及び内径
側部分では、インビンジメント冷却に比べ冷却効率の低
い強制対流冷却方式を採用しているため、翼メタル温度
は第１の実施例よりも高く、翼中央部とのメタル温度差
を小さくできる。又、隔壁９は、前縁側から後縁側にか
けて、半径方向長さが短くなるようにコード方向に配置
し、この隔！！９で囲まれた空間のみをインピンジメン
ト冷却している。これは、前縁側ではインビンジメント
冷却後の、隔壁９で囲まれない空間を通過する冷却空気
が少ないためであり、後縁側では次第に各隔離ブロック
からの冷却空気が増加されて流れこむので強制対流冷却
面を大きくできることになる０本実施例によれば熱応力
の緩和効果を第１の実施例より発揮することができると
いう利点がある。In this embodiment, the entire cooling air flow rate can be used for impingement cooling in the center of the blade, which has severe thermal conditions, so the cooling air flow rate ratio, which is the ratio of the mainstream gas amount to the cooling air flow rate, is different from that in the first embodiment. If they are the same, the metal temperature at the center of the blade is the first.
This can be lower than that of the embodiment. Furthermore, since the forced convection cooling method, which has lower cooling efficiency than impingement cooling, is adopted for the outer diameter side and inner diameter side of the blade, the temperature of the blade metal is higher than in the first embodiment, and the temperature of the blade metal is higher than that of the first embodiment. The difference in metal temperature can be reduced. Moreover, the partition wall 9 is arranged in the cord direction so that the radial length becomes shorter from the leading edge side to the trailing edge side. ! Only the space surrounded by 9 is impingement cooled. This is because on the leading edge side, after impingement cooling, there is less cooling air passing through the space not surrounded by the partition wall 9, and on the trailing edge side, the cooling air from each isolation block gradually increases and flows in, resulting in forced convection. According to this embodiment, which allows the cooling surface to be enlarged, there is an advantage that the effect of alleviating thermal stress can be exerted more effectively than in the first embodiment.
尚１本実施例では突起状の乱流促進体１０を翼内Ｍ４に
設けた場合について述べたが１本発明は乱流促進体１０
をコアプラグ２側へ設けても適用できる。また、第５図
に示す形状に拘束されないのは勿論である。更に、必要
に応じて翼内壁４及びコアプラグ２を平滑面として強制
対流冷却しても本発明を適用できる。In this embodiment, the case where the protruding turbulence promoting body 10 is provided in the blade M4 is described;
It can also be applied by providing it on the core plug 2 side. Furthermore, it goes without saying that the shape shown in FIG. 5 is not restrictive. Furthermore, if necessary, the present invention can be applied even if the blade inner wall 4 and the core plug 2 are made smooth and forced convection cooling is performed.
以上詳述したように、本発明を適用するとガスタービン
静翼の中央部での冷却効率を高め、これによって冷却所
要空気量を低減して当該ガスタービンの熱効率を向上せ
しめるとともに、静翼の温度分布を均一ならしめて熱応
力、熱歪を軽減し、耐久性向上に貢献し得るという優れ
た実用的効果を奏する。As described in detail above, when the present invention is applied, the cooling efficiency at the center of the gas turbine stator blade is increased, thereby reducing the amount of air required for cooling, improving the thermal efficiency of the gas turbine, and increasing the temperature of the stator blade. It has an excellent practical effect of making the distribution uniform, reducing thermal stress and thermal strain, and contributing to improved durability.
第１図は本発明の第１の実施例を示すガスタービン静翼
の部分断面斜視図、第２図は第１図のＡ−Ａ断面図であ
る。第３図及び第４図はそれぞれ上記第１の実施例の変
形例の説明図である。第５図は本発明の第２の実施例を
示すガスタービン静翼の部分断斜視面図、第６図は従来
のガスタービン静翼の１例を示す部分断面斜視図である
。
１・・・静翼、２・・・コアプラグ、３・・・インピン
ジ孔、４・・・翼内壁、５・・・空間、９・・・隔壁、
１０・・・乱流促進体。FIG. 1 is a partially sectional perspective view of a gas turbine stationary blade showing a first embodiment of the present invention, and FIG. 2 is a sectional view taken along line AA in FIG. FIGS. 3 and 4 are explanatory diagrams of modified examples of the first embodiment, respectively. FIG. 5 is a partially sectional perspective view of a gas turbine stationary blade according to a second embodiment of the present invention, and FIG. 6 is a partially sectional perspective view showing an example of a conventional gas turbine stationary blade. DESCRIPTION OF SYMBOLS 1... Stationary blade, 2... Core plug, 3... Impingement hole, 4... Blade inner wall, 5... Space, 9... Partition wall,
10...Turbulence promoter.
Claims (1)
しめてコアプラグを固定して、該コアプラグにインピン
ジ孔を配列すると共に、前記空洞と連通せしめて静翼後
続流路を設けたガスタービン静翼において、前記コアプ
ラグと静翼本体の内壁面とに挟まれた空間をコード方向
に仕切る複数の隔壁を設けたことを特徴とするガスター
ビン静翼。 ２、前記複数個の隔壁はガスタービンロータの半径方向
に設置されたものであり、かつ、前縁側の端に配設した
隔壁の半径方向が最も長く、後縁側に向けて順次に短く
なつていること特徴とする特許請求の範囲第１項に記載
のガスタービン静翼。 ３、前記コアプラグと静翼本体と挟まれた空間の、翼長
方向に関して両端付近に位置せしめて、乱流促進体を設
けたことを特徴とする特許請求の範囲第１項又は同第２
項に記載のガスタービン静翼。[Scope of Claims] 1. A core plug is fixed to the inner wall surface of a cavity formed in the main body of the stator vane so as to be spaced apart from each other, and impingement holes are arranged in the core plug, and the impingement holes are arranged in the core plug, and the core plug is communicated with the cavity to form a rear part of the stator vane. A gas turbine stator blade provided with a flow path, characterized in that the gas turbine stator blade is provided with a plurality of partition walls that partition a space sandwiched between the core plug and an inner wall surface of a stator blade main body in a chord direction. 2. The plurality of partition walls are installed in the radial direction of the gas turbine rotor, and the partition wall installed at the leading edge side is longest in the radial direction, and gradually becomes shorter toward the trailing edge side. The gas turbine stationary blade according to claim 1, characterized in that: 3. Turbulence promoters are provided near both ends of the space between the core plug and the stationary blade body in the blade length direction, as claimed in claim 1 or 2.
The gas turbine stator blade described in .
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP22199486A JPS6380004A (en) | 1986-09-22 | 1986-09-22 | Gas turbine stator blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP22199486A JPS6380004A (en) | 1986-09-22 | 1986-09-22 | Gas turbine stator blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPS6380004A true JPS6380004A (en) | 1988-04-11 |
Family
ID=16775420
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP22199486A Pending JPS6380004A (en) | 1986-09-22 | 1986-09-22 | Gas turbine stator blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS6380004A (en) |
Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4946346A (en) * | 1987-09-25 | 1990-08-07 | Kabushiki Kaisha Toshiba | Gas turbine vane |
| JPH02241902A (en) * | 1989-03-13 | 1990-09-26 | Toshiba Corp | Cooling blade of turbine and combined generating plant utilizing gas turbine equipped with this blade |
| JP2001041003A (en) * | 1999-07-16 | 2001-02-13 | General Electric Co <Ge> | Prestressed gas turbine nozzle |
| JP2006017119A (en) * | 2004-06-30 | 2006-01-19 | Snecma Moteurs | Improved cooling stationary turbine blade |
| US7217095B2 (en) | 2004-11-09 | 2007-05-15 | United Technologies Corporation | Heat transferring cooling features for an airfoil |
| JP2012052513A (en) * | 2010-09-03 | 2012-03-15 | Ihi Corp | Turbine blade |
| WO2013089250A1 (en) * | 2011-12-15 | 2013-06-20 | 株式会社Ihi | Impingement cooling mechanism, turbine blade and combustor |
| US20140290257A1 (en) * | 2011-12-15 | 2014-10-02 | Ihi Corporation | Impingement cooling mechanism, turbine blade and cumbustor |
| WO2015109040A1 (en) * | 2014-01-15 | 2015-07-23 | Siemens Aktiengesellschaft | Internal cooling system with corrugated insert forming nearwall cooling channels for gas turbine airfoil |
| WO2015157780A1 (en) * | 2014-04-09 | 2015-10-15 | Siemens Aktiengesellschaft | Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil including heat dissipating ribs |
| WO2016036366A1 (en) * | 2014-09-04 | 2016-03-10 | Siemens Aktiengesellschaft | Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil |
| WO2016036367A1 (en) * | 2014-09-04 | 2016-03-10 | Siemens Aktiengesellschaft | Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil |
| JP2016540150A (en) * | 2013-11-07 | 2016-12-22 | シーメンス アクチエンゲゼルシヤフトＳｉｅｍｅｎｓ Ａｋｔｉｅｎｇｅｓｅｌｌｓｃｈａｆｔ | Investment casting for the vane segment of gas turbine engines. |
-
1986
- 1986-09-22 JP JP22199486A patent/JPS6380004A/en active Pending
Cited By (26)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4946346A (en) * | 1987-09-25 | 1990-08-07 | Kabushiki Kaisha Toshiba | Gas turbine vane |
| JPH02241902A (en) * | 1989-03-13 | 1990-09-26 | Toshiba Corp | Cooling blade of turbine and combined generating plant utilizing gas turbine equipped with this blade |
| JP2001041003A (en) * | 1999-07-16 | 2001-02-13 | General Electric Co <Ge> | Prestressed gas turbine nozzle |
| JP4546334B2 (en) * | 2004-06-30 | 2010-09-15 | スネクマ | Turbine stator blades with improved cooling |
| JP2006017119A (en) * | 2004-06-30 | 2006-01-19 | Snecma Moteurs | Improved cooling stationary turbine blade |
| US7819169B2 (en) | 2004-11-09 | 2010-10-26 | United Technologies Corporation | Heat transferring cooling features for an airfoil |
| US7217095B2 (en) | 2004-11-09 | 2007-05-15 | United Technologies Corporation | Heat transferring cooling features for an airfoil |
| JP2012052513A (en) * | 2010-09-03 | 2012-03-15 | Ihi Corp | Turbine blade |
| WO2013089250A1 (en) * | 2011-12-15 | 2013-06-20 | 株式会社Ihi | Impingement cooling mechanism, turbine blade and combustor |
| JP2013124630A (en) * | 2011-12-15 | 2013-06-24 | Ihi Corp | Impingement cooling mechanism, turbine blade, and combustor |
| CN103975129A (en) * | 2011-12-15 | 2014-08-06 | 株式会社Ihi | Impingement cooling mechanism, turbine blade and combustor |
| US20140290257A1 (en) * | 2011-12-15 | 2014-10-02 | Ihi Corporation | Impingement cooling mechanism, turbine blade and cumbustor |
| US9957812B2 (en) * | 2011-12-15 | 2018-05-01 | Ihi Corporation | Impingement cooling mechanism, turbine blade and cumbustor |
| US9771809B2 (en) | 2011-12-15 | 2017-09-26 | Ihi Corporation | Impingement cooling mechanism, turbine blade and combustor |
| JP2016540150A (en) * | 2013-11-07 | 2016-12-22 | シーメンス アクチエンゲゼルシヤフトＳｉｅｍｅｎｓ Ａｋｔｉｅｎｇｅｓｅｌｌｓｃｈａｆｔ | Investment casting for the vane segment of gas turbine engines. |
| WO2015109040A1 (en) * | 2014-01-15 | 2015-07-23 | Siemens Aktiengesellschaft | Internal cooling system with corrugated insert forming nearwall cooling channels for gas turbine airfoil |
| WO2015157780A1 (en) * | 2014-04-09 | 2015-10-15 | Siemens Aktiengesellschaft | Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil including heat dissipating ribs |
| WO2016036367A1 (en) * | 2014-09-04 | 2016-03-10 | Siemens Aktiengesellschaft | Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil |
| CN106661945A (en) * | 2014-09-04 | 2017-05-10 | 西门子公司 | Internal Cooling System With Insert Forming Nearwall Cooling Channels In An Aft Cooling Cavity Of A Gas Turbine Airfoil |
| CN106795771A (en) * | 2014-09-04 | 2017-05-31 | 西门子公司 | Inner cooling system with the insert that nearly wall cooling duct is formed in cooling chamber in the middle part of the wing chord of gas turbine aerofoil profile |
| CN107075955A (en) * | 2014-09-04 | 2017-08-18 | 西门子公司 | Include the inner cooling system of cooling fin with the insert that nearly wall cooling duct is formed in the rear portion cooling chamber of combustion gas turbine airfoil |
| WO2016036366A1 (en) * | 2014-09-04 | 2016-03-10 | Siemens Aktiengesellschaft | Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil |
| JP2017532483A (en) * | 2014-09-04 | 2017-11-02 | シーメンス アクチエンゲゼルシヤフトＳｉｅｍｅｎｓ Ａｋｔｉｅｎｇｅｓｅｌｌｓｃｈａｆｔ | Internal cooling system having an insert forming a near-wall cooling passage in the rear cooling cavity of a gas turbine blade |
| JP2017532482A (en) * | 2014-09-04 | 2017-11-02 | シーメンス アクチエンゲゼルシヤフトＳｉｅｍｅｎｓ Ａｋｔｉｅｎｇｅｓｅｌｌｓｃｈａｆｔ | Internal cooling system having an insert that forms a near-wall cooling passage within a chord central cooling cavity of a gas turbine blade |
| US9840930B2 (en) | 2014-09-04 | 2017-12-12 | Siemens Aktiengesellschaft | Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil |
| US9863256B2 (en) | 2014-09-04 | 2018-01-09 | Siemens Aktiengesellschaft | Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of an airfoil usable in a gas turbine engine |
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