JPS6310284B2 - - Google Patents
Info
- Publication number
- JPS6310284B2 JPS6310284B2 JP4588383A JP4588383A JPS6310284B2 JP S6310284 B2 JPS6310284 B2 JP S6310284B2 JP 4588383 A JP4588383 A JP 4588383A JP 4588383 A JP4588383 A JP 4588383A JP S6310284 B2 JPS6310284 B2 JP S6310284B2
- Authority
- JP
- Japan
- Prior art keywords
- blade
- head
- cooling air
- main body
- gas
- 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
Links
- 238000001816 cooling Methods 0.000 description 57
- 239000007789 gas Substances 0.000 description 48
- 238000007664 blowing Methods 0.000 description 8
- 230000007423 decrease Effects 0.000 description 6
- 239000000567 combustion gas Substances 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 230000001771 impaired effect Effects 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 238000005192 partition Methods 0.000 description 2
- 239000011148 porous material Substances 0.000 description 2
- 239000007779 soft material Substances 0.000 description 2
- 239000000956 alloy Substances 0.000 description 1
- 229910045601 alloy Inorganic materials 0.000 description 1
- 239000000919 ceramic Substances 0.000 description 1
- 229910010293 ceramic material Inorganic materials 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 230000013011 mating Effects 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 238000000926 separation method Methods 0.000 description 1
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
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Description
【発明の詳細な説明】
本発明は主と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。DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a stationary blade of a gas turbine mainly used in a high-temperature gas turbine or the like.
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されてい
る。 In recent years, gas turbines have become increasingly hotter in order to improve performance and increase output, and as a result, gas turbine blades are exposed to high temperatures. Since there is no material available, the method used is to cool the wings.
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、ここに冷却空気
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より翼外に吹出し、
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、フイルム冷却した第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に至つている。 Stator blades (hereinafter referred to as blades for convenience in this explanation) used in conventional gas turbines are shown in Figure 1-A,
As shown in the examples of Figures 1-B, 1-C, and 1-D, the blade 1 is formed hollow, cooling air is introduced here, and the inside is cooled by convection. What is shown is that a core 4 is provided inside a hollow blade 1, and cooling air is introduced into the core 4, and the air is blown out toward the inner surface of the blade through a number of pores 5 at the tip of the core 4, and is locally cooled. In addition, cooling air is introduced into the hollow blade 1 and blown out of the blade through numerous pores 6 at the leading edge of the blade,
There are blades 1 covered with a cooling air layer to block heat from high-temperature combustion gas and film-cooled, as shown in Figure 1-C. This leads to wing 1 in Figure 1-D.
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、この前縁部の冷
却が最も重要であり、ガスタービンの高温化にと
もなつてフイルム冷却を併用し、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つて
いる。 Here, the leading edge of the gas turbine blade 1 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, so cooling this leading edge is the most important. As the temperature of the turbine 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をフイルム冷却し、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 accordingly, and the average mainstream gas temperature will decrease, thereby reducing the cycle efficiency of the gas turbine. will decrease.
また、翼１を冷却する冷却空気は、通常第２図
の系統図に示すように、ガスタービンのタービン
部１０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、
これに接続された配管等を通つて翼１内に供給さ
れる。 Cooling air for cooling the blades 1 is usually obtained by extracting air compressed by a compressor 8 driven by a turbine section 10 of a gas turbine before a combustor 9, as shown in the system diagram in FIG. , casing or
It is supplied into the blade 1 through piping etc. connected to this.
このため、冷却空気量が増加すれば圧縮機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。 Therefore, if the amount of cooling air increases, the power required for compression by the compressor 8 will increase, and the efficiency and output of the gas turbine 10 will decrease by this amount.
また、フイ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みならず、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、翼面に沿つ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 completely cool the film,
The pressure difference between the cooling air and the pressure of the mainstream gas needs to be appropriate. If this pressure difference is small, not only will blowing not occur locally, but the mainstream gas may flow back into the blade, which will affect the cooling performance. On the other hand, if the pressure difference is too large, the cooling air will be blown out forcefully, and if the blowing angle to the blade surface is large, it will be difficult to form a cooling air layer along the blade surface, and even aerodynamic performance will be 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
行なわれるように、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１０の翼列に至るま
での間に絞り抵抗等を設け、主流ガスの圧力を下
げる場合もある。 Generally, the pressure of the mainstream gas is only slightly lower than the pressure of the cooling air, so in order to ensure complete blowing, the main stream gas is routed from the outlet of the compressor 8 in the mainstream gas system to the blade row of the turbine section 10 of the gas turbine. In some cases, a throttle resistor or the like is provided between these steps to lower the pressure of the mainstream gas.
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、そのままロスと
なり、出力は低下する。 In this way, lowering the pressure of the mainstream gas is
Since this amount is not involved in work, it becomes a loss and the output decreases.
また、翼１の各部より冷却空気を吹出し、フイ
ルム冷却を行なう場合には、翼面に沿つて主流ガ
スに圧力分布があり、それぞれの位置に所定量の
冷却空気を吹出すための翼構造は複雑となつてい
る。 In addition, when cooling air is blown out from each part of the blade 1 to perform film cooling, there is a pressure distribution in the mainstream gas along the blade surface, and the blade structure is required to blow out a predetermined amount of cooling air to each location. It's getting complicated.
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
フイルム冷却を行ない、これに必要な冷却空気量
も多くなつているため、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、また主流ガス圧力を下げるた
めの出力低下等の問題があり、この対策が強く望
まれていた。 As mentioned above, in the structure of conventional cooled blades,
As the temperature of gas turbines increases, film cooling is performed from the leading edge of the blade, and the amount of cooling air required for this is also increasing, resulting in a decrease in gas turbine thermal efficiency due to mainstream gas cooling and a loss in the required power of the compressor. In addition, there were problems such as a reduction in output due to lowering the mainstream gas pressure, and 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。 Therefore, the present invention solves the above-mentioned conventional problems,
This was done with the aim of making it possible to improve the efficiency of gas turbines.
即ち、本発明はガスタービンの静翼の頭部と本
体部とを別体に形成すると共に、その静翼のプラ
ツトフオーム及びシユラウドに該頭部寸法よりや
や大きい寸法を有し、かつその内面に突起を有し
た穴及び溝内に、該頭部の上下両端部を装着する
ことにより構成される。 That is, the present invention forms the head and main body of a stator vane of a gas turbine separately, and has a platform and shroud of the stator vane with dimensions slightly larger than the head size, and an inner surface of the stator vane. It is constructed by fitting the upper and lower ends of the head into holes and grooves that have protrusions.
以下図面を参照して本発明の実施例を説明する
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
の静翼のキヤンバーラインに沿つた断面図で、第
５図は第３図の翼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。 Embodiments of the present invention will be described below with reference to the drawings. FIG. 3 is a sectional view of a stator blade of a gas turbine in an embodiment of the present invention, and FIG. 4 is a sectional view of the stator blade of FIG. 3. Figure 5 is a cross-sectional view of the wing head in Figure 3, along the camber line of Figure 1-A.
The same parts as in the conventional example shown in FIGS. 1-D are indicated by the same part numbers.
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
ツトフオーム、１９がシユラウド、そして２０が
キヤツプである。 First, in the blade 1 of the present invention applied to the turbine section 10 of a gas turbine similar to that explained in the conventional example of FIG.
is a 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 20 is a cap.
次に、この翼１では頭部１２と本体部１３とが
別体に形成されており、頭部１２は、本体部１
３、プラツトフオーム１８、シユラウド１９と同
じ耐熱合金で形成することも、または本体部１３
とは異なるセラミツク材で形成しても良い。 Next, in this wing 1, the head 12 and the main body 13 are formed separately, and the head 12 is formed from the main body 1.
3. The platform 18 and the shroud 19 can be made of the same heat-resistant alloy, or the main body 13
It may be formed from a ceramic material different from the above.
頭部１２の範囲は、主流ガスがせき止められる
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 range where the mainstream gas is blocked or the range where the heat transfer coefficient is high.
また、頭部１２は本体部１３側で凸となるよう
な曲線、あるいは折線等でその分割線が翼外面と
接する角度が大きくなるように本体部１３と分け
ている。 Further, the head 12 is separated from the main body 13 by a convex curve on the main body 13 side or by a broken line so that the angle at which the dividing line contacts the outer surface of the wing is large.
また、本体部１３およびプラツトフオーム１８
とシユラウド１９とは１体となつている。 In addition, the main body 13 and the platform 18
and Shroud 19 are one body.
更に、頭部１２は上下両端部２ケ所に溝を設
け、軟かい材料のブツシユ２４を取付けている。 Furthermore, the head 12 is provided with grooves at two locations at both upper and lower ends, and bushes 24 made of soft material are attached thereto.
プラツトフオーム１８には、頭部１２寸法より
やや大きな穴２１を設け、シユラウド１９には、
頭部１２の寸法よりやや大きな溝２２を設け、穴
２１と溝２２の内面には、環状の、あるいは環状
に複数個設けた突起２３を設け、頭部１２を穴２
１を通して溝２２にはめ込み、穴２１にキヤツプ
２０をかぶせ、そのキヤツプ上端を全周溶接す
る。 The platform 18 is provided with a hole 21 that is slightly larger than the head 12, and the shroud 19 is provided with a hole 21 that is slightly larger than the head 12.
A groove 22 that is slightly larger than the size of the head 12 is provided, and an annular or plural annular projections 23 are provided on the inner surfaces of the hole 21 and the groove 22.
1 and fit into the groove 22, the cap 20 is placed over the hole 21, and the upper end of the cap is welded all around.
このとき、穴２１と溝２２の突起２３は、頭部
１２のブツシユ２４に接し、頭部１２を固定す
る。 At this time, the hole 21 and the protrusion 23 of the groove 22 contact the bush 24 of the head 12 and fix the head 12.
また、本体部１３には、仕切１５によつて先端
部１４と後縁部２に分けた中空部を設け、その先
端に冷却空気吹出し孔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。 Further, the main body part 13 is provided with a hollow part divided into a tip part 14 and a rear edge part 2 by a partition 15, and a large number of cooling air blowing holes 17 are bored in the tip part, and the outer surface of the hollow part is divided into a tip part 14 and a rear edge part 2. A cooling air passage 16 is provided on the mating surface with 12, and the hollow portion of the trailing edge 2 has an internal convection cooling structure.
本発明の静翼は、以上のように構成されてお
り、本体部１３の先端部１４および後縁部２に冷
却空気を導き、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１
７より頭部１２と本体部１３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を通つ
て翼外に吹出され、本体部１３を冷却空気層でお
おい、フイルム冷却する。 The stator vane of the present invention is configured as described above, and the cooling air is guided to the tip 14 and the trailing edge 2 of the main body 13, and the cooling air guided to the hollow part of the tip 14 is directed to the main body 13. 13 Tip cooling air outlet 1
7 into the cooling air passage 16 between the head 12 and the main body 13, and is blown out of the blade through the cooling air passage 16, covering the main body 13 with a cooling air layer and cooling the film.
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より翼外に吹出される。 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 of the blade from the cooling air blowing hole 3 at the trailing edge.
なお、ここで、シユラウド１９に穴２１を設
け、プラツトフオーム１８に溝２２を設けても、
または双方に穴を設けても良い。 Note that even if the hole 21 is provided in the shroud 19 and the groove 22 is provided in the platform 18,
Alternatively, holes may be provided on both sides.
以上のごとく、本発明では翼１の頭部１２を、
他の翼構造部、即ち、本体部１３、プラツトフオ
ーム１８、シユラウド１９等と分けてあり、翼１
の構造強度は後者でもち、頭部１２にかかる空気
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
It is separated from other wing structural parts, namely, the main body part 13, the platform 18, the shroud 19, etc., and the wing 1
The structural strength of the head 12 is maintained by the latter, and the aerodynamic force applied to the head 12 is also supported by the main body, so the 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変
形することもあるが、頭部１２の上下両端部の側
面には軟かい材質のブツシユ２４が取付けられて
おり、穴２１および溝２２の突起２３と、線ある
いは点で接触しているため、翼１が変形しても、
ブツシユ２４が変形することにより、頭部１２に
大きな力が働くことはない。 In addition, although the blade 1 may be deformed due to thermal expansion of the turbine casing or its own thermal expansion, bushes 24 made of soft material are attached to the sides of the upper and lower ends of the head 12. Since the blade 1 is in contact with the protrusion 23 of the hole 21 and the groove 22 in a line or point, even if the blade 1 is deformed,
Due to the deformation of the bush 24, no large force is applied to the head 12.
ただし、このためにはブツシユ２４に頭部１２
の構造強度より弱い材料を選ぶ必要があり、また
翼１の変形量と同等量が頭部１２の構造強度より
弱い力で変形する必要があり、その必要変形量は
具体的には0.1〜0.15mmである。 However, for this purpose, the head 12 is attached to the bush 24.
It is necessary to select a material whose structural strength is weaker than that of the head 12, and it is necessary to deform the wing 1 by an amount equivalent to the structural strength of the head 12 with a force weaker than the structural strength of the head 12. Specifically, the required deformation is 0.1 to 0.15. mm.
なお、熱伸びにより翼１全体が膨張する場合
は、頭部１２に力が働くことはない。 Note that when the entire blade 1 expands due to thermal elongation, no force is applied to the head 12.
このため頭部１２に、構造強度に対する信頼性
が不十分のため従来翼１を構成できなかつたセラ
ミツクを用いることもできる。 For this reason, it is also possible to use ceramic for the head 12, which could not conventionally be used to construct the wing 1 due to insufficient reliability in terms of structural strength.
なお、キヤツプ２０をプラツトフオーム１８に
全周溶接したのは、主流ガスが穴２１の間隙を通
つ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.
従つて、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前縁部の頭部が本体部と
は別体に形成されているので、頭部が高温により
膨張しても本体部には影響を与えることがなく、
翼全体としての構造強度を十分に維持することが
できる。 Therefore, in the present invention, the mainstream gas is dammed,
The head of the leading edge, which is the hottest part of the wing, is formed separately from the main body, so even if the head expands due to high temperature, it will not affect the main body.
The structural strength of the wing as a whole can be maintained sufficiently.
特に本発明では、静翼のプラツトフオーム及び
シユラウドに設けた頭部寸法よりやや大きな穴及
び溝内に、突起を介して頭部の上下両端部を装着
しており、頭部が膨張してもそれを十分吸収し、
その支持強度を維持できる。 In particular, in the present invention, both upper and lower ends of the head are mounted via protrusions in holes and grooves that are slightly larger than the head size provided in the stator vane platform and shroud, so that the head expands. also fully absorbs it,
The support strength can be maintained.
また、本発明では、頭部と本体部との分割面に
冷却空気通路を設け、その冷却空気通路より冷却
空気を本体部側面に吹出し、本体部をフイルム冷
却することができ、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, a cooling air passage is provided in the dividing plane between the head and the main body, and cooling air is blown out from the cooling air passage to the side of the main body, so that the main body can be film-cooled, and the blade as a whole can be cooled. For example, there will be no leading edge airflow, and air will be emitted from the side.
翼前縁からフイルム冷却を行なう場合、翼前縁
には主流ガスの動圧分が加わるため、冷却空気の
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、圧力は下つているため、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 leading edge of the blade, so the pressure of the cooling air needs to be higher than this.In order to maintain this pressure difference, the pressure of the mainstream gas system must be increased. Sometimes this 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, and the mainstream gas system There is no need to lower the pressure, which increases the efficiency of the gas turbine.
また、上記の翼では、頭部と本体部との分割線
が翼外面と接する角度を大きくとることができる
ので、分割面にある冷却空気通路を通つて翼外に
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 above wing, the angle at which the parting line between the head and the main body touches the outer surface of the wing can be set large, so that the cooling air blown out of the wing through the cooling air passage in the parting surface is It will blow out at a small angle backwards.
このため、冷却空気の圧力が主流ガスの圧力よ
り高くなつて勢よく吹出しても、翼面に沿つて冷
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
となる。 Further, 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 of the blade and the trailing edge of the blade.
冷却空気を翼内から翼外に吹出す量は、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、翼前縁と翼側面等か
ら吹出しを行なう場合、主流ガスには翼面に沿つ
た圧力分布があり、そ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
つているが、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、圧力の下がつた翼側面および翼後縁からの吹
出しとなれば、翼面に沿つ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。 The amount of cooling air blown from the inside of the blade to the outside of the blade is determined by the total cross-sectional area of the cooling air outlet, depending on the pressure difference between the cooling air and the mainstream gas, so when blowing from the leading edge of the blade and the side of the blade, etc. , the mainstream gas has a pressure distribution along the blade surface, and the blade structure is complicated to make the amount of cooling air blown out at each position a predetermined amount. If the cooling air is no longer blown out from the leading edge, the mainstream gas is accelerated, and the air is blown out from the lower pressure side of the blade and the trailing edge of the blade, the cooling air will be distributed according to the pressure distribution of the mainstream gas along the blade surface. The blade structure for blowing out a predetermined amount 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てあるため、
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 dividing the wing into the head and the main body, the wings are separated so that the main body side is convex.
Even if the direction of the aerodynamic force acting on the head changes, this force can be effectively supported by the main body.
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.
また別体に形成した頭部が、何らかの原因で破
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 separately formed 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, can also be easily replaced.
第１−Ａ図、第１−Ｂ図、第１−Ｃ図及び第１
−Ｄ図は、それぞれ異なる従来の冷却式の静翼の
断面図、第２図はガスタービンの系統図、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
ヤンバーラインに沿つた断面図で、第５図は第３
図の翼頭部の断面図である。
１……翼、１０……ガスタービンのタービン
部、１１……発電機、１２……頭部、１３……本
体部、１８……プラツトフオーム、１９……シユ
ラウド、２０……キヤツプ、２１……穴、２２…
…溝、２３……突起。
Figure 1-A, Figure 1-B, Figure 1-C, and Figure 1
-D is a sectional view of different conventional cooling type stator vanes, Fig. 2 is a system diagram of a gas turbine, and Fig. 3 is a sectional view of a stator blade of a gas turbine according to an embodiment of the present invention. There is a
FIG. 3 is a cross-sectional view of the wing head shown in the figure. DESCRIPTION OF SYMBOLS 1... Blade, 10... Turbine part of gas turbine, 11... Generator, 12... Head, 13... Main body part, 18... Platform, 19... Shroud, 20... Cap, 21 ...hole, 22...
...Groove, 23...Protrusion.
Claims (1)
に形成すると共に、該静翼のプラツトフオーム及
びシユラウドに該頭部寸法よりやや大きい寸法を
有し、かつその内面に突起を有した穴及び溝内
に、該頭部の上下両端部を装着したことを特徴と
するガスタービンの静翼。1 The head and main body of the stator blade of a gas turbine are formed separately, and the platform and shroud of the stator blade have dimensions slightly larger than the head size, and have protrusions on the inner surface. A stationary blade for a gas turbine, characterized in that both upper and lower ends of the head are mounted in the hole and groove.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP4588383A JPS59173503A (en) | 1983-03-22 | 1983-03-22 | Stationary blade of gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP4588383A JPS59173503A (en) | 1983-03-22 | 1983-03-22 | Stationary blade of gas turbine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPS59173503A JPS59173503A (en) | 1984-10-01 |
| JPS6310284B2 true JPS6310284B2 (en) | 1988-03-05 |
Family
ID=12731628
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP4588383A Granted JPS59173503A (en) | 1983-03-22 | 1983-03-22 | Stationary blade of gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS59173503A (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP2295722B1 (en) | 2009-09-09 | 2019-11-06 | Ansaldo Energia IP UK Limited | Blade of a turbine |
-
1983
- 1983-03-22 JP JP4588383A patent/JPS59173503A/en active Granted
Also Published As
| Publication number | Publication date |
|---|---|
| JPS59173503A (en) | 1984-10-01 |
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