JPS59160009A - Stationary blade for gas turbine - Google Patents
Stationary blade for gas turbineInfo
- Publication number
- JPS59160009A JPS59160009A JP3270483A JP3270483A JPS59160009A JP S59160009 A JPS59160009 A JP S59160009A JP 3270483 A JP3270483 A JP 3270483A JP 3270483 A JP3270483 A JP 3270483A JP S59160009 A JPS59160009 A JP S59160009A
- Authority
- JP
- Japan
- Prior art keywords
- cooling air
- blade
- hole
- platform
- head
- 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.)
- Granted
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
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/284—Selection of ceramic materials
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Ceramic Engineering (AREA)
- Materials Engineering (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【発明の詳細な説明】
本発明は主−すして高温ガスタービン等に使用されるガ
スタービンの静翼に関するものである近年、ガスタービ
ンは、性能向上および出力上昇のためますます高温化の
傾向にあり、このため、ガスタービンの翼は高温にさら
されることになるが、現在このような高温下で強度を有
する材料はないため、翼を冷却する方法が採用従来のガ
スタービンに使用される静翼（以下本説明では便宜上翼
と略称する）は、第１−Ａ図、第１−Ｂ図、第１−Ｃ図
及び第１−Ｄ図の例に示すように、翼１を中空に形成し
、冷却空気を導き、内部を対流冷却した第１−Ａ図に示
したもの、中空状の翼１内に中子４を設け、その中子４
内に冷却空気を導き、中子４先端の多数の細孔５より翼
内面に向けてその空気を吹出し、局所的に熱伝達を高め
、強制冷却した第１−Ｂ図に示したもの、さらに中空状
の翼１内に冷却空気を導き、翼前縁部の多数の細孔６よ
り真性に吹出し、翼１を冷却空気層でおおい、高温の燃
焼ガスから熱を遮断し、フィルム冷却した第１−Ｃ図に
示したもの等があり、ガスタービンが高温化するにつれ
て、これらを組合せて使用する第１−Ｄ図の翼１に至っ
ている。[Detailed Description of the Invention] The present invention primarily relates to stationary blades for gas turbines used in high-temperature gas turbines, etc.In recent years, gas turbines have tended to have higher temperatures in order to improve performance and increase output. Because of this, gas turbine blades are exposed to high temperatures, but there are currently no materials that have strength at such high temperatures, so a method of cooling the blades is adopted, which is used in conventional gas turbines. A stator blade (hereinafter abbreviated as a blade for convenience in this description) is a blade 1 made of a hollow blade, as shown in the examples of Fig. 1-A, Fig. 1-B, Fig. 1-C, and Fig. 1-D. A core 4 is provided in a hollow blade 1, and the core 4 is
The one shown in Figure 1-B in which cooling air is introduced into the core 4, and the air is blown out toward the inner surface of the blade through numerous pores 5 at the tip of the core 4 to locally increase heat transfer and force cooling. Cooling air is guided into the hollow blade 1, and is blown out from the numerous pores 6 at the leading edge of the blade, covering the blade 1 with a cooling air layer, insulating heat from the high-temperature combustion gas, and creating a film-cooled air layer. There are blades such as those shown in Fig. 1-C, and as gas turbines become hotter, a combination of these blades is used, leading to the blade 1 shown in Fig. 1-D.
また、第１−Ｅ図の従来例に示すごとく、プラットフォ
ーム１８およびシュラウド１９の通路面にも冷却空気吹
出し孔２４より冷却空気を吹出なお、上記第１−Ａ図か
ら第１−’Ｅ図において、同じ部品は同じ部品番号で示
している。In addition, as shown in the conventional example in Fig. 1-E, cooling air is also blown out from the cooling air blowing holes 24 on the passage surfaces of the platform 18 and the shroud 19. , the same parts are indicated by the same part numbers.
なお、上記第１−Ａ図から第１−Ｄ図において、同じ部
品は同じ部品番号で示している。In addition, in the said FIG. 1-A to FIG. 1-D, the same parts are shown by the same part number.
ここで、ガスタービンの翼１で燃焼ガスにさらされて最
も高温となるのは、主流ガスがせき止められる翼１の前
縁部であるので、この前線部の冷却が最も重要であり、
ガスタービンの高温化にともなってフィルム冷却を併用
し、また、この部分を冷却するのに必要な冷却空気の量
も多くなっている。Here, since the leading edge of the blade 1 of the gas turbine that is exposed to the combustion gas and reaches the highest temperature is the leading edge of the blade 1 where the mainstream gas is dammed up, cooling this front part is the most important.
As the temperature of gas turbines increases, film cooling is also used, and the amount of cooling air required to cool this part is also increasing.
しかしながら、翼１をフィルム冷却し、これに必要な冷
却空気の量が増加すれば、それだけ主流ガスに混合する
冷却空気の量が増し、平均主流ガス温度が低下し、この
ためガスタービンのサイクル効率は低下することになる
。However, if the blade 1 is film-cooled and the amount of cooling air required for this increases, the amount of cooling air mixed with the mainstream gas will increase, the average mainstream gas temperature will decrease, and this will reduce the cycle efficiency of the gas turbine. will decrease.
また、翼１を冷却する冷却空気は、ｌ常第２図の系統図
に示すように、ガスタービンのタービン部１ｏで駆動さ
れる圧縮機８で圧縮された空気を、燃焼器９前で抽気し
、ケーシングあるいは１、これに接続された配管等を通
って翼１内に供″′給される。In addition, the cooling air that cools the blades 1 is normally produced by extracting air compressed by a compressor 8 driven by the turbine section 1o of the gas turbine before the combustor 9, as shown in the system diagram in Fig. 2. The air is then supplied into the blade 1 through the casing or 1 and piping connected thereto.
このため、冷却空気量が増加すれば圧縮ｇ１．８で圧縮
するための所要動力が多くなり、この分だけガスタービ
ン１０の効率及び出力が低下することになる。For this reason, if the amount of cooling air increases, the power required for compression at compression g1.8 will increase, and the efficiency and output of the gas turbine 10 will decrease by this amount.
また、フィルム冷却を完全に行なうためには、主流ガス
の圧力に対する冷却空気の圧力差が適正である必要があ
り、この圧力差が小さいと局所的に吹出しが行なわれな
いのひならず、主流ガスが翼内部へ逆流することもあり
、冷却性能が損なわれ、逆に圧力差が大きすぎると、冷
却空気が勢いよく吹出し、翼面に対する吹出し角が大き
い場合、翼面に沿った冷却空気層が形成され難く、空力
性能までもが損なわれる。In addition, in order to achieve complete film cooling, the pressure difference between the cooling air and the pressure of the mainstream gas needs to be appropriate. Gas may flow back inside the blade, impairing cooling performance, and conversely, if the pressure difference is too large, the cooling air will blow out forcefully, and if the blowing angle to the blade surface is large, the cooling air layer along the blade surface will collapse. is difficult to form, and even aerodynamic performance is impaired.
一般に、主流ガスの圧力は、冷却空気の圧力よりわずか
に低いだけであるため、吹出しが完全に行なわれるより
に、主流ガス系の圧縮機８出口からガスタービンのター
ビン部１ｏの翼列に至るまでの間に絞り抵抗等を設け、
主流ガスのヰ′力を下げる場合もある。In general, the pressure of the mainstream gas is only slightly lower than the pressure of the cooling air, so the flow from the outlet of the compressor 8 in the mainstream gas system to the blade row of the turbine section 1o of the gas turbine takes place before the blowing is completed. In the meantime, provide aperture resistance, etc.
In some cases, the power of the mainstream gas may be lowered.
ノ＼１このように、主流ガスの圧力を下げることは、こ
の分が仕事に関与しないため、そのｉｔロス仁なり、出
力は低下する。＼1 In this way, lowering the pressure of the mainstream gas does not involve work, resulting in IT loss and the output decreases.
また、翼１の各部より冷却空気を吹出し、フｊ゛リレム
冷却を行なう場合には、翼面に沿って主流ガスに圧力分
布があり、それぞれの位置に所定量の冷却空気を吹出す
だめの翼構造は、複雑となっている。In addition, when cooling air is blown out from each part of the blade 1 to perform fringe cooling, there is a pressure distribution in the mainstream gas along the blade surface, and there is a need for a reservoir to blow out a predetermined amount of cooling air at each location. The wing structure is complex.
また、冷却空気の吹出し孔を設けることは、それだけ加
工の手間がかかり、コスト上昇をまねき、強度が低下し
、翼寿命は短かくなる。Further, providing cooling air blow-off holes requires a lot of processing time, increases costs, reduces strength, and shortens blade life.
以上のように、従来の冷却式の翼の構造では、ガスター
ビンの高温化にともない、翼前縁部からフィルム冷却を
行ない、これに必要な冷却空気量も多くなっているため
、主流ガス冷却によるガスタービン熱効率の低下と、圧
縮機所要動力にしめるロスが多くなり、また卑流ガス圧
カを下げるための出力像下等の問題があり、この対策が
強く望まれていた。As mentioned above, in the conventional cooling type blade structure, as the temperature of the gas turbine increases, film cooling is performed from the leading edge of the blade, and the amount of cooling air required for this increases, so the mainstream gas cooling There were problems such as a decrease in the thermal efficiency of the gas turbine due to this, a large loss in the required power of the compressor, and a decrease in the output image due to the lowering of the downstream gas pressure, so countermeasures against this problem were strongly desired.
そこで、本発明は前記従来の問題点を解消し、ガスター
ビンの効率向上を可能ならしめることを目的としてなさ
れたものである。SUMMARY OF THE INVENTION Therefore, the present invention has been made with the object of solving the above-mentioned conventional problems and making it possible to improve the efficiency of a gas turbine.
・ト即ち、本発明は、ガスタービンの静翼の頭部−に１
他の静翼構造部と分け、かつセラミックで構成すると共
に、該静翼のプラットフォーム及びシュラウドに穴ある
いは溝を設け、それらの−穴あるいは溝内に該頭部の上
下両端部を装着し、号がつそれらの穴あるいは溝の内面
に冷却空気吹、出し用の細孔を設けることにより構成さ
れる。・In other words, the present invention provides a structure in which:
It is separated from other stator blade structures and is made of ceramic, and holes or grooves are provided in the platform and shroud of the stator blade, and both upper and lower ends of the head are installed in these holes or grooves. It is constructed by providing pores for blowing and discharging cooling air on the inner surface of these holes or grooves.
以下、図面を参照して本発明の詳細な説明するが、第３
図は本発明の一実施例におけるガスタービンの静翼の翼
部断面図であり、第４図は第３図の静翼のキャンバ−ラ
インに沿った断面図で、第５図は第３図の数頭部の断面
図であり、第１−Ａから第１−Ｄ図に示す従来例と同じ
部品は同じ部品番号で示している。The present invention will be described in detail below with reference to the drawings.
The figure is a cross-sectional view of the stator blade of a gas turbine according to an embodiment of the present invention, FIG. 4 is a cross-sectional view of the stator blade shown in FIG. 3 along the camber line, and FIG. FIG. 1 is a cross-sectional view of several heads of the conventional example shown in FIGS. 1-A to 1-D.
まず、第２図の従来例で説明したと同様のガスタービン
のタービン部１０に適用される本発明の翼１において、
１２が頭部、１３が本体部、１４が中空の先端部、１５
が仕切、１６が冷却空気通路、１７が先端の冷却空気吹
出し孔、１８がプラットフォーム、１９がシュラウド、
そして２ｏがキャップである。First, in the blade 1 of the present invention applied to the turbine section 10 of a gas turbine similar to that described in the conventional example of FIG.
12 is the head, 13 is the main body, 14 is the hollow tip, 15
is a partition, 16 is a cooling air passage, 17 is a cooling air outlet at the tip, 18 is a platform, 19 is a shroud,
And 2o is the cap.
次に、この翼１では頭部１２がセラミックで形成されて
おり、本体部１６およびプラットフォーム１８とシュラ
ウド１９とは金属１、即ち、耐熱台ｔで形成されている
。Next, in this wing 1, the head 12 is made of ceramic, and the main body 16, platform 18, and shroud 19 are made of metal 1, that is, a heat-resistant base t.
頭部１２の範囲は、主流ガスがせき止められｂ１範囲、
あるいは、熱伝達率の高い範囲までとする。The range of the head 12 is the b1 range where the mainstream gas is dammed,
Alternatively, it should be within the range where the heat transfer coefficient is high.
また、頭部１２は本体部１３側が凸となるような曲線、
あるいは折線等でその分割線が翼列面；と接する角度が
大きくなるように本体部１６ と分けている。Further, the head 12 is curved so that the main body 13 side is convex,
Alternatively, it is separated from the main body portion 16 by a broken line or the like such that the angle at which the dividing line contacts the blade row surface is large.
また、本体部１３およびプラットフォーム１８とシュラ
ウド１９とは一体となっている。Further, the main body portion 13, the platform 18, and the shroud 19 are integrated.
更に、プラットフォーム１８に、頭部１２寸法よりやや
大きな穴２１を設け、シュラウド１９にも頭部１２の寸
法よりやや大きな溝２２を設け、頭部１２を穴２１を通
して溝２２にさし込み、穴２１にキャップ２０をし、キ
ャップ２０上端を全周溶接する。Further, the platform 18 is provided with a hole 21 that is slightly larger than the head 12, and the shroud 19 is also provided with a groove 22 that is slightly larger than the head 12. A cap 20 is attached to 21, and the upper end of the cap 20 is welded all around.
本体部１６には、仕切１５によって先端部１４と後縁部
２とに分けた中空部を設け、その先端に細孔の冷却空気
吹出し孔１７を多数穿設し、かつその外面、即ち、頭部
１２との合せ面には冷却空気通路１６を設け、後縁部２
の中空部は内部対流冷却構造とする。The main body part 16 is provided with a hollow part divided into a tip part 14 and a rear edge part 2 by a partition 15, and a number of small cooling air blowing holes 17 are bored at the tip, and the outer surface, that is, the head A cooling air passage 16 is provided on the mating surface with the rear edge portion 2.
The hollow part has an internal convection cooling structure.
プラントフオーム１８には、穴２１の内面に細孔の冷却
空気吹出し孔２４を穿設し、合せてキャツｌプ２０にも
その下端面に細孔の冷却空気吹出し孔２４を穿設し、ま
たシュラウド１９には、溝２２の内面に細孔の冷却空気
吹出し孔２４を穿設している。In the plant form 18, a small cooling air blowing hole 24 is formed in the inner surface of the hole 21, and in addition, a small cooling air blowing hole 24 is formed in the lower end surface of the cap 20. The shroud 19 is provided with fine cooling air blowing holes 24 in the inner surface of the grooves 22.
本発明は以上のように構成されており、プラットフォー
ム１８およびシュラウド１９双方に、冷却空気を導き、
冷却空気吹出し孔２４ より穴２１内面および溝２２内
面に吹出し、頭部１２からプラットフォーム１８および
シュラウド１９へ１の、熱伝達を防止し、プラットフォ
ーム１８側に導かれた冷却空気は、本体部１３の先端部
１４および後縁部２の中空部に導かれ、先端部１４の中
空部に導かれた冷却空気は、本体部１３先端の冷却空気
吹出し孔１７より頭部１２と本体部１６との間の冷却空
気通路１６に吹出され、その冷却空気通路１６を通って
翼列に吹出され、本体部１６を冷却空気層でおおい、フ
ィルム冷却する。The present invention is configured as described above, and guides cooling air to both the platform 18 and the shroud 19.
Cooling air is blown from the cooling air blowing holes 24 to the inner surfaces of the holes 21 and the grooves 22 to prevent heat transfer from the head 12 to the platform 18 and shroud 19, and the cooling air guided to the platform 18 side is The cooling air that is guided into the hollow portions of the tip portion 14 and the rear edge portion 2 is passed between the head portion 12 and the body portion 16 through the cooling air outlet 17 at the tip of the body portion 13. The main body 16 is covered with a layer of cooling air and film-cooled.
また、後縁部２の中空部に導かれた冷却空気は、本体部
１３の内部を対流冷却し、後縁の冷却空気吹出し孔３よ
り翼列に吹出される。Further, the cooling air guided into the hollow portion of the trailing edge portion 2 convects the inside of the main body portion 13 and is blown out from the cooling air blowing holes 3 at the trailing edge to the blade rows.
ここで、シュラウド１９に穴２１を設け、プラットフォ
ーム１８に溝２２を設けてもまた双方に穴を設けてもよ
い。Here, the shroud 19 may be provided with holes 21 and the platform 18 may be provided with grooves 22, or both may be provided with holes.
以上のごとく、本発明では翼１の頭部１２を、弛の翼構
造部、即ち、本体部１６、プラットフォーム１８、シュ
ラウド１９等と分けてあり、翼１の構造強度は後者でも
ち、頭部１２にかがる空気力も本体部でささえるため、
頭部１２は構造強度を必要としない。As described above, in the present invention, the head 12 of the wing 1 is divided into the loose wing structural parts, that is, the main body 16, the platform 18, the shroud 19, etc., and the structural strength of the wing 1 is maintained in the latter, and the head Since the main body also supports the aerodynamic force exerted on 12,
Head 12 does not require structural strength.
また、翼１はタービンケーシングの熱伸び等の影響を受
け、あるいは自からの熱伸び等によ、り一変形すること
もあるが、これらに頭部１２を収１付けるためのプラッ
トフォーム１８の穴２１と、ン１ニラウド１９の溝２２
とは頭部１２より大きく、縦部１２との間に間隙がある
ため、翼１が変形してもこの力が頭部１２に加わること
はない。In addition, the blade 1 may be deformed due to the thermal expansion of the turbine casing or its own thermal expansion, but the hole in the platform 18 for fitting the head 12 therein may 21, and the groove 22 of N1Niloud 19.
is larger than the head 12 and there is a gap between it and the vertical part 12, so even if the wing 1 deforms, this force will not be applied to the head 12.
即ち、翼１が変形していなけれず、頭部１２は空気力に
よりその後面が本体部１６先端、および穴２１と溝２２
の後面と接しており、翼１からは何ら力を受けていない
が、翼１が変形すれば穴２１と溝２２の中心がずれたり
、曲がったり、本体部、１６がせり出したりし、頭部１
２に力が作用する。That is, the wing 1 must be deformed, and the rear surface of the head 12 will be connected to the tip of the main body 16 and the holes 21 and grooves 22 due to the aerodynamic force.
Although it is in contact with the rear surface and does not receive any force from the wing 1, if the wing 1 deforms, the centers of the holes 21 and grooves 22 may shift or bend, and the main body 16 may protrude, causing the head 1
Force acts on 2.
ここで、穴２１と溝２２に間隙がなければ、翼１が変形
すれば、その力は全て頭部１２にも働くが、穴２１と溝
２２に間隙があるので、翼１が変形しても頭部１２は穴
２１と溝２２の中で移動し大きな力は働かない。Here, if there is no gap between the hole 21 and the groove 22, when the blade 1 deforms, all the force will also act on the head 12, but since there is a gap between the hole 21 and the groove 22, the blade 1 will deform. However, the head 12 moves within the hole 21 and groove 22 and no large force is applied.
従って、穴２１と溝２２の間隙は翼１の変形量より大き
いことが必要で、具体的には０．１〜０．１５調あれば
よい。Therefore, the gap between the hole 21 and the groove 22 needs to be larger than the amount of deformation of the blade 1, and specifically, it is sufficient if it is 0.1 to 0.15.
なお、熱伸びにより翼１全体が膨張する場合・、穴２１
と溝２２の中心線がずれたり、本体部６がせり出してく
ることもないので、翼１ の張に対する穴２１と溝２２
の間隙は考慮の必要稈重鴎
乙のだめ頭部１２に構造強度に対する信頼性が不１．ｔ
−分のため従来翼１を構成できなかったセラミ−ツクを
用いることができる。In addition, if the entire blade 1 expands due to thermal elongation, the holes 21
Since the center line of the groove 22 does not shift and the main body 6 does not protrude, the hole 21 and the groove 22 are
The gap needs to be taken into consideration.The reliability of the structural strength of the culm-juge-otsu nodame head 12 is insufficient. t
Therefore, ceramics, which could not be used to construct the blade 1 in the past, can be used.
なお、キャップ２０をプラットフォーム１８に全周溶接
したのは、主流ガスが穴２１の間隙を通って主流ガス通
路外にもれることを防止するためである。The reason why the cap 20 is welded to the platform 18 all the way around is to prevent the mainstream gas from leaking out of the mainstream gas passage through the gap between the holes 21.
このため、本発明では主流ガスがせき止められ、翼とし
て最も高温となるその前縁部、即ち頭部が金属より耐熱
性が高いセラミックとなっているので、金属より高い温
度で何月でき、金属のようにその前縁から冷却空気を吹
出し、フィルム冷却したりする必要はなくなる。For this reason, in the present invention, the mainstream gas is dammed up, and the leading edge, or head, which is the highest temperature for the blade, is made of ceramic, which has higher heat resistance than metal. There is no need for film cooling by blowing cooling air out from the leading edge.
また、金属の本体部、プラットフォーム、シュラウドは
冷却を必要とするが、本体部は翼として特に高温となる
部分ではなく、その先端はセラミックでできた頭部が燃
焼ガスの熱を遮断し、また頭部と本体部との間には、冷
却空気通路を冷却空気が流れており、頭部からの熱伝達
ｅ、〜防止しているため、従来の翼の前縁部のようｒ−
高温とはならず、本体部先端を冷却するため１衿冷却空
気量は少なくてすむ。In addition, although the metal main body, platform, and shroud require cooling, the main body is not a part that gets particularly hot as a wing, and the tip of the main body is a ceramic head that blocks the heat of combustion gas. Cooling air flows through the cooling air passage between the head and the main body, preventing heat transfer from the head.
Since the temperature does not reach high temperatures and the tip of the main body is cooled, the amount of cooling air per collar is small.
一方、本体部側面は燃焼ガスにさらされるた１め１、頭
部と本体部間の冷却空気通路より冷却空気１を吹出して
フィルム冷却を行なう。On the other hand, since the side surface of the main body is exposed to combustion gas, cooling air 1 is blown out from the cooling air passage between the head and the main body to perform film cooling.
ここで、プラットフォームおよびシュラウドは、頭部温
度が従来より高温となるため、その取付部である穴およ
び溝に冷却空気を吹出し、頭部からの熱伝達を防止する
。Here, since the head temperature of the platform and shroud is higher than before, cooling air is blown into the holes and grooves where the platform and shroud are attached to prevent heat transfer from the head.
このように頭部の冷却が不要になり、本体部を冷却する
冷却空気量も少なくなれば、翼全体として必要な冷却空
気量は少なくなる。In this way, if cooling of the head becomes unnecessary and the amount of cooling air used to cool the main body is reduced, the amount of cooling air required for the entire wing will be reduced.
冷却空気は、翼を冷却した後、あるいはフィルム冷却を
行なうため、主流ガス中に吹出すが、主流ガスと混合し
、これを冷却するため、冷却空気量が少なくなればそれ
だけ平均主流ガス温度の低下がおさえられ、ガスタービ
ンの熱効率が向上する。Cooling air is blown into the mainstream gas after cooling the blades or for film cooling, but it mixes with the mainstream gas and cools it, so the smaller the amount of cooling air, the lower the average mainstream gas temperature. The thermal efficiency of the gas turbine is improved.
また、冷却空気量が減れば、圧縮機で圧縮した空気を冷
却空気として抽気する量が減り、この分を出力として取
り出せるため、ガスタービン効率および出力が向上する
。Furthermore, if the amount of cooling air is reduced, the amount of air compressed by the compressor that is extracted as cooling air is reduced, and this amount can be extracted as output, thereby improving gas turbine efficiency and output.
また、本発明では、本体部をフィルム冷却する冷却空気
を、頭部と本体部の分割面にある冷ｌ空気通路より吹出
し、本体部側面をその前端事）らカバーするが、翼全体
としてみれば、前縁吹出しはなくなり、側面からの吹出
しとなる。In addition, in the present invention, the cooling air for film cooling the main body is blown out from the cold air passage in the dividing plane between the head and the main body, and covers the side surface of the main body from its front end. For example, there will be no leading edge airflow, and air will be emitted from the side.
翼前縁からフィルム冷却を行なう場合、翼前線には主流
ガスの動圧分が加わるため、冷却空ノ気の圧力はこれよ
り高いことが必要で、この圧力差を保つため、主流ガス
系の圧力−をわざと下げ、ることもあるが、翼後縁から
吹出す場合は、主流ガスが加速し、圧力は下っているた
め、主流ガスと冷却空気の圧力差は保たれることになり
、主流ガス系の圧力を下げる必要はなくなり、この分ガ
スタービンの効率が向上する。When performing film cooling from the leading edge of the blade, the dynamic pressure of the mainstream gas is applied to the blade front, so the pressure of the cooling air needs to be higher than this, and in order to maintain this pressure difference, the pressure of the mainstream gas system is Sometimes the pressure is intentionally lowered, but when blowing out from the trailing edge of the blade, the mainstream gas accelerates and the pressure decreases, so the pressure difference between the mainstream gas and the cooling air is maintained. There is no need to lower the pressure in the mainstream gas system, and the efficiency of the gas turbine improves accordingly.
また、本発明では頭部と本体部の分割線が翼列面と接す
る角度を大きくとっであるので、分割面にある冷却空気
通路を通って翼列に吹出す冷却空気は、翼後方に小さな
角度で吹出すことになる。In addition, in the present invention, the angle at which the dividing line between the head and the main body touches the blade row surface is set at a large angle, so that the cooling air blown out to the blade row through the cooling air passage in the dividing surface has a small space behind the blade. It will blow out at an angle.
このため、冷却空気の圧力が主流ガスの圧力より高くな
って勢よく吹出しても、翼面に沿って冷却空気層が形成
され、′冷却性能や空力性能が損なわれることはない。Therefore, even if the pressure of the cooling air becomes higher than the pressure of the mainstream gas and is blown out vigorously, a cooling air layer is formed along the blade surface, and the cooling performance and aerodynamic performance are not impaired.
また、本発明では、翼前縁からの冷却空気吹出しがなく
なり、翼側面および翼後縁からの吹出しとなり、冷却空
気を翼列に吹出す量は、冷却空気と主流ガスの圧力差に
応じて冷却空気吹出し孔の総断面積で規定するため、翼
前縁と翼ＱＵ面等から吹出しを行なう場合、主流ガスに
は翼面に沿った圧力分布があり、雪−れぞれの位置の冷
却空気吹出し量を所定の量にするための翼構造は複、雑
となっているが、主流ガスの動圧分゛を°受ける翼前縁
からの冷却空気吹出しがなくなり、主流ガスが加速し、
圧力の下がった翼側面および翼後縁からの吹出しとなれ
ば、翼面に沿った主流ガスの圧力分布に応じて冷却空気
を所定量吹出すための翼構造は簡単となる。In addition, in the present invention, cooling air is no longer blown out from the leading edge of the blade, but instead is blown out from the side surface and trailing edge of the blade, and the amount of cooling air blown out to the blade cascade depends on the pressure difference between the cooling air and the mainstream gas. Since it is defined by the total cross-sectional area of the cooling air outlet, when blowing from the leading edge of the blade, the QU surface, etc. of the blade, there is a pressure distribution in the mainstream gas along the blade surface, and the cooling of each position of the snow Although the blade structure for controlling the amount of air blown out to a predetermined level is complex, the cooling air blown out from the leading edge of the blade, which receives the dynamic pressure of the mainstream gas, is no longer blown out, and the mainstream gas accelerates.
If the air is blown out from the blade side surface and the trailing edge of the blade where the pressure has decreased, the blade structure for blowing out a predetermined amount of cooling air in accordance with the pressure distribution of the mainstream gas along the blade surface becomes simple.
また、本発明では翼を頭部と本体部に分けるとき、本体
部側が凸となるように分けであるため、頭部に働く空気
力の方向が変化してもこの力は有効に本体部でささえる
ことができる。In addition, in the present invention, when the wing is divided into the head and the main body, the main body side is convex, so even if the direction of the aerodynamic force acting on the head changes, this force is effectively transferred to the main body. I can support you.
また、頭部と本体部との組合せは、凹及び凸となり、頭
部が本体部とずれて段差ができ、翼面を流れる主流ガス
が剥離し、空力性能が低下することも防止できる。Further, the combination of the head and the main body is concave and convex, and it is possible to prevent the head from shifting from the main body, creating a step, causing separation of the mainstream gas flowing on the wing surface, and reducing aerodynamic performance.
また、何らかの原因でセラミックでできた頭部が破損し
ても、本体部は先端が凸形状の翼形をなしており、ある
程度の空力性能は保たれると共に、また頭部が破損して
も簡単、に取替える、ことができる。In addition, even if the ceramic head is damaged for some reason, the main body has an airfoil shape with a convex tip, so a certain level of aerodynamic performance is maintained, and even if the head is damaged, It can be easily replaced.
第１−Ａ図、第１−Ｂ図、第１−Ｃ図、第１−Ｄ図はそ
れぞれ異なる従来の冷却式の静翼の平断面図、第１−Ｅ
図は他の従来例の冷却式の゛静；臆の側断面図、第２図
はガスタービンの系統７図：くｊ第３図は本発明の一実
施例におけるガスタービン壱ンの静翼の翼部断面図であ
り、第４図は第３図の静翼のキャンバ−ラインに沿った
断面図で、第５図は第３図の数頭部の断面図である。
１・・・翼、１０・・・ガスタービンのタービン部、１
１・・・発電機、１２・・・頭部、１６・・・本体部、
１６・・・冷却空気通路、１８・・・プラットフォーム
、１９・・・シュラウド、２０・・・キャップ、２１・
・・穴、２２・・・溝、２４・・・冷却空気吹出し孔。
出願人 工業技術院長 石 坂 誠−
第１−ｋｌ 第１−Ｃ図
第１−Ｅ図
第 ３ 図Figure 1-A, Figure 1-B, Figure 1-C, and Figure 1-D are plan sectional views of different conventional cooling type stationary blades, and Figure 1-E.
The figure is a side sectional view of another conventional cooling type cooling system, and Figure 2 is a diagram of a gas turbine system. FIG. 4 is a cross-sectional view of the stator vane of FIG. 3 along the camber line, and FIG. 5 is a cross-sectional view of several heads of the stator blade of FIG. 3. 1... Blade, 10... Turbine part of gas turbine, 1
1... Generator, 12... Head, 16... Main body,
16... Cooling air passage, 18... Platform, 19... Shroud, 20... Cap, 21...
...hole, 22...groove, 24...cooling air outlet. Applicant Makoto Ishizaka, Director General of the Agency of Industrial Science and Technology - 1-kl Figure 1-C Figure 1-E Figure 3
Claims (1)
かつセラミックで形成すると共に、該静翼のプラットフ
ォーム及びシュラウドに穴あるいは溝を設け、それらの
穴あるいは溝内に該頭部の上下両端部を装着し、かつそ
れらの穴あるいは溝の内面に冷却空気吹出し用の細孔を
設けたことを特徴とするガスタービンの静翼。Separate the head of the gas turbine stator blade from the other stator blade structure,
The platform and shroud of the stationary vane are formed with holes or grooves, the upper and lower ends of the head are installed in the holes or grooves, and the inner surfaces of the holes or grooves are provided with cooling air. A stator blade for a gas turbine characterized by having pores for blowing out.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP3270483A JPS59160009A (en) | 1983-03-01 | 1983-03-01 | Stationary blade for gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP3270483A JPS59160009A (en) | 1983-03-01 | 1983-03-01 | Stationary blade for gas turbine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPS59160009A true JPS59160009A (en) | 1984-09-10 |
| JPS6242122B2 JPS6242122B2 (en) | 1987-09-07 |
Family
ID=12366228
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP3270483A Granted JPS59160009A (en) | 1983-03-01 | 1983-03-01 | Stationary blade for gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS59160009A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1333160C (en) * | 2002-10-16 | 2007-08-22 | 三菱重工业株式会社 | Gas turbine |
Families Citing this family (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH04330Y2 (en) * | 1986-07-21 | 1992-01-08 | ||
| JPH0411506U (en) * | 1990-05-22 | 1992-01-30 |
-
1983
- 1983-03-01 JP JP3270483A patent/JPS59160009A/en active Granted
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1333160C (en) * | 2002-10-16 | 2007-08-22 | 三菱重工业株式会社 | Gas turbine |
Also Published As
| Publication number | Publication date |
|---|---|
| JPS6242122B2 (en) | 1987-09-07 |
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