JPS59176401A - Air-cooled gas turbine - Google Patents
Air-cooled gas turbineInfo
- Publication number
- JPS59176401A JPS59176401A JP4888683A JP4888683A JPS59176401A JP S59176401 A JPS59176401 A JP S59176401A JP 4888683 A JP4888683 A JP 4888683A JP 4888683 A JP4888683 A JP 4888683A JP S59176401 A JPS59176401 A JP S59176401A
- Authority
- JP
- Japan
- Prior art keywords
- cooling
- air
- blade
- stage rotor
- rotor blades
- 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
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
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【発明の詳細な説明】
〔発明の利用分野〕
本発明はガスタービンの動力発生部を構成する動翼及び
ロータの冷却構造に関する。DETAILED DESCRIPTION OF THE INVENTION [Field of Application of the Invention] The present invention relates to a cooling structure for moving blades and a rotor that constitute a power generating section of a gas turbine.
ガスタービンサイクルでは、タービン入口温度を高めて
熱落差を大きくするほど熱効率を向上し得る。このため
、通常のガスタービンでは、燃焼用圧縮空気の一部を冷
媒として抽気して翼等を冷却子ることによシ、作動ガス
の高温化を図っている。したがって、タービン全体の総
合効率を最大限に発揮するためには、極力抽気量を少な
くして有効に冷却する必要がある。In a gas turbine cycle, thermal efficiency can be improved by increasing the turbine inlet temperature and increasing the heat drop. For this reason, in a typical gas turbine, a portion of the compressed air for combustion is extracted as a refrigerant and a blade or the like is used as a cooler to raise the temperature of the working gas. Therefore, in order to maximize the overall efficiency of the entire turbine, it is necessary to reduce the amount of extracted air as much as possible for effective cooling.
以下、この種ガスタービンについて図を用いて説明する
。This type of gas turbine will be explained below using figures.
第1図は従来のガスタービンの動翼及びロータ部分の断
面構造を示す。10がロータ、20,30がそれぞれ第
1段動翼、第2段動翼である。同図において作動ガスは
矢印60で示す方向に流れ、静翼40,50と動420
.30を交互に通過し膨張を繰シ返えすことによって、
ロータ10の回転力として動力を発生する。作動ガスで
ある燃焼ガスの温度は、タービン効率を向上する目的か
ら翼材の耐熱温度をはるかに越えているため、特に、動
翼には強大な回転遠心力に対しても耐え得るように、そ
の内部に複雑な冷却流路21.31を設けた構造になっ
ておシ、冷媒としての空気をロータ10に直結して発生
動力の一部によシ駆動する圧縮機で生成した燃焼用空気
の一部を抽気し外部の熱交換器を介して矢印61に沿っ
て導入し、あるいは、圧縮機から直接ロータ10の中心
部を経由して矢印62に沿って導き冷却するようにガっ
ている。作動ガスの温度は後段にいくに従って低下する
ため、段毎に動翼の冷却熱負荷も異なり、高圧段側はど
複雑な冷却構造となっておシ、作動ガスの温度レベルに
よっても異なるが、伝熱を促進スるタービュレン子プロ
モータ22、ビンフィン23等で冷却強化され、冷却空
気の流量も、熱負荷に応じて、冷却流路抵抗を利用して
配分される。これらの冷却に供した空気は、吹出し孔2
4、あるいは、Wb翼30の先端部にある流路31の出
口から翼外に流出し作動ガス中に混入する。したがって
、冷却空気を抽気することは圧縮機に余分の動力を必要
とするほか、作動ガス中に混入して温度を低下させ、更
に、ロータ10内を通過する過程で冷却空気に回転エネ
ルギを吸収されるため、ガスタービン効¥低下の大きな
要因となり、作動ガスの温度を高めて効率向上をはかつ
ているにもかかわらず、光分にその効果を発揮し得ない
。FIG. 1 shows a cross-sectional structure of a rotor blade and a rotor portion of a conventional gas turbine. 10 is a rotor, and 20 and 30 are a first stage rotor blade and a second stage rotor blade, respectively. In the same figure, the working gas flows in the direction shown by the arrow 60, and the working gas flows between the stator vanes 40, 50 and the movable blade 420.
.. By passing through 30 alternately and repeating the expansion,
Power is generated as the rotational force of the rotor 10. The temperature of the combustion gas, which is the working gas, far exceeds the heat resistance temperature of the blade material in order to improve turbine efficiency. It has a structure in which complicated cooling channels 21 and 31 are provided inside it, and air as a refrigerant is directly connected to the rotor 10, and combustion air is generated by a compressor that is driven by a part of the generated power. A portion of the air is extracted and introduced along arrow 61 through an external heat exchanger, or alternatively, it is guided directly from the compressor through the center of rotor 10 along arrow 62 for cooling. There is. Since the temperature of the working gas decreases as it goes to the later stages, the cooling heat load on the rotor blades differs for each stage, and the high-pressure stage side has a complex cooling structure, which also varies depending on the temperature level of the working gas. Cooling is enhanced by the turbulent promoter 22, bin fins 23, etc. that promote heat transfer, and the flow rate of cooling air is also distributed using the cooling flow path resistance according to the heat load. The air used for these cooling is
4, or flows out from the outlet of the flow path 31 at the tip of the Wb blade 30 and mixes into the working gas. Therefore, extracting the cooling air requires extra power for the compressor, mixes it into the working gas and lowers its temperature, and also absorbs rotational energy into the cooling air as it passes through the rotor 10. This is a major factor in reducing the efficiency of the gas turbine, and even though efficiency has been improved by increasing the temperature of the working gas, the effect cannot be exerted on light components.
本発明の目的は、冷却空気を有効に活用してその流量を
少なくすることによシ、効率の高いガスタービンを提供
するにある。An object of the present invention is to provide a highly efficient gas turbine by effectively utilizing cooling air and reducing its flow rate.
本発明は、二段動翼内部の伝熱面積を拡大することによ
り、一段動翼の冷却に供したあとの排出空気条件で二段
動翼を冷却しうるεとに着目して、一段動翼内の冷却流
路出口がロータを介して第二段動翼内冷却流路の入口に
通じるようにし、一段動翼の排出空気で二段動翼を冷却
するようにしたことを特徴とする。The present invention focuses on the fact that by expanding the heat transfer area inside the second-stage rotor blade, the second-stage rotor blade can be cooled under the exhaust air conditions after cooling the first-stage rotor blade. The outlet of the cooling channel in the blade communicates with the inlet of the cooling channel in the second stage rotor blade via the rotor, so that the exhaust air of the first stage rotor blade cools the second stage rotor blade. .
以下、第2図によって本発明の動翼及び四−タ内の冷却
流路構成について説明する。第一段動翼については、そ
の■−■及びIV−IV断面を第3図及び第4図に示す
ように、隔壁83で仕切ることによって中心部の流路8
1と外周部の流路82に分離する。外周部の流路82は
、更に、動翼の外力に対する補強と共に、翼外面の熱負
荷に応じて冷却空気の流量を配分するだめに、リプ84
で複数個の流路に仕切っである。中心部流路81と外周
部流路82は、翼先端部の連通孔86あるいは熱負荷の
大きい前縁部及び後縁部では翼の半径方向に穿けた複数
個の吹出し孔87で連通しており、流路82の他端は更
に第4図に示すように翼のプラットホーム89に近い部
分のシャンク90の側面に開口する。この開口部は仕切
シ92とプラットホーム89及び従来のシャンク側壁に
よって囲まれ、翼同士が隣接することによってヘッダ9
2となっておシ、シャンク側壁の切欠き93によって翼
外に開口している。一方、第二段の動翼95はシャンク
の高圧側壁のプラットホームに近い部分にある流入口9
6からシャンク内部を経て複数個の冷却流路97に分岐
し、翼の先端部で作動ガス流路に開口する。この冷却流
路97は、従来よシも孔径を大きくして伝熱面及び流路
断面積を拡大しである。更に、ロータ70を形成してい
るスペーサ71の外側に、同スペーサ71と一体にガイ
ドリング72を形成して、その両端が一段及び二段動翼
側壁のプラットホーム近傍にかん合するようにし、スペ
ーサとガイドリンク間のスペース73を経て一段動翼と
二段動翼の冷却流路が連通ずるようにする。そこで、デ
ィスクに芽けた冷却空気の流入孔76からロータ70内
に流入した冷却空気は矢印75に沿ってディスク外周部
で第一段動翼に流入し、中心部流路81、連通孔86ま
だは吹出し孔85、外周部流路82、ヘッダ91の順に
動翼内部を循環して動翼を冷却したあと切欠き93から
翼部のスペース73内に流出し、同スペースを経てガイ
ドリング72の案内によシ流入口96から第二段動翼に
流入して冷却した後、翼先端部から作動ガス流路中に排
出される。スペーサ71を貫通している孔77を通る流
路は第二段動翼とディスクのかん合をなすダブテイル9
8を冷却するための流路であり、動翼を回転支持し最も
大きな応力が発生する部分なので、ロータ流人直後の冷
却空気によって冷却し、その後、スペース73内で一段
動翼排出空気中に混入する。Hereinafter, the rotor blade and cooling passage configuration in the quadrature according to the present invention will be explained with reference to FIG. As for the first stage rotor blade, as shown in FIGS. 3 and 4, the flow path in the center is partitioned by dividing the
1 and a flow path 82 on the outer periphery. The flow path 82 on the outer periphery is further provided with a lip 84 for reinforcing the rotor blade against external forces and distributing the flow rate of cooling air according to the thermal load on the outer surface of the blade.
It is partitioned into multiple flow channels. The center flow path 81 and the outer peripheral flow path 82 communicate with each other through a communication hole 86 at the tip of the blade or a plurality of blow holes 87 drilled in the radial direction of the blade at the leading and trailing edges where the heat load is large. The other end of the channel 82 further opens into the side of the shank 90 in a portion near the platform 89 of the wing, as shown in FIG. This opening is surrounded by the partition 92 and platform 89 and conventional shank sidewalls, and the adjoining wings allow the header 9
2 and is open to the outside of the wing through a notch 93 in the side wall of the shank. On the other hand, the second stage rotor blade 95 has an inlet 9 located near the platform on the high pressure side wall of the shank.
6, it branches into a plurality of cooling channels 97 through the inside of the shank, and opens into a working gas channel at the tip of the blade. This cooling flow path 97 has a larger hole diameter than the conventional cooling flow path to enlarge the heat transfer surface and the cross-sectional area of the flow path. Furthermore, a guide ring 72 is formed integrally with the spacer 71 on the outside of the spacer 71 forming the rotor 70, and both ends of the guide ring 72 are fitted near the platforms of the side walls of the first and second stage rotor blades. The cooling channels of the first-stage rotor blade and the second-stage rotor blade are communicated through a space 73 between the rotor blade and the guide link. Therefore, the cooling air that has flowed into the rotor 70 from the cooling air inflow holes 76 that have sprouted in the disk flows into the first stage rotor blades at the outer circumference of the disk along the arrow 75, and flows through the center flow passage 81 and the communication hole 86. After cooling the rotor blade by circulating inside the rotor blade in this order through the blow-off hole 85, the outer circumferential flow path 82, and the header 91, it flows out from the notch 93 into the space 73 of the blade portion, and passes through the space into the guide ring 72. The gas flows into the second-stage rotor blade through the inlet 96 under guidance, is cooled, and is then discharged from the blade tip into the working gas flow path. The flow path passing through the hole 77 penetrating the spacer 71 is a dovetail 9 that engages the second stage rotor blade and the disk.
8, which rotatably supports the rotor blades and generates the greatest stress, so it is cooled by the cooling air immediately after the rotor drifts, and then it is cooled by the cooling air immediately after the rotor drifts, and then in the first stage rotor blade exhaust air in the space 73. Mixed.
第一段動翼は、後縁から吹き出す従来の構造を回収式に
変更した点は異なるが、基本的な冷却法はほとんど類似
しておシ、また、冷却空気の流入条件も全く同じである
ため、従来と同様の冷却効果が得られる。第二段の動翼
は、第一段動翼冷却用の排出空気を利用するため、冷却
空気の温度レベルは上昇するが、流量変化に対して対流
冷却流路の断面積を比例的に拡大することによシ対流熱
伝達率が流量増加比の約0.4乗に比例して向上するた
め、従来に要した二段動翼冷却用空気流量の2倍以上の
流量を供給することによシ、翼全体の冷却効率は大巾に
向上し、従来の翼温度レベルを維持できる。一方、ダブ
テイルの冷却に要する空気流量は、間部が直接タービン
作動ガスにさらされることはなく、また、シャンクが牛
径方向に長いために、高温の翼部からダブテイルへの熱
流抵抗が大きく、熱負荷としては非常に小さいために、
少ない量で充分である。したがって、?1ぼ、従来に要
した二段動翼冷却空気流量に近い流量が節減できる。!
!、た、冷却空気流量が多いために、同空気の冷却時の
温度変化が少なく、翼面温度分布のむらが少なくなる利
点がある。更に、圧縮空気を外部の熱交換器を介した後
、ロータに供給する場合には、節約分の冷却空気流量を
第一段動翼に加算供給することによシ、第一段動翼の冷
却効率が向上し、排出空気の温度が下がるために、第二
段動翼の冷却効果も向上し、冷却空気流量を従来と同じ
にすれば、タービン入口温度の高温化をはかることがで
きる。一方、タービンの効率的には、第一段動翼で冷却
空気を排出しなくなった分だけ、第二段動翼で発生する
動力が減少するが、同排出空気が作動ガス中に混入しな
いことによる作動ガスの温度低下防止の方がよシ効果的
であシ、冷却空気の節約分に相幽する圧縮機の動力及び
冷却空気をロータと同一回転させる分の動力が軽減され
、タービンの総合効率を向上することができる。The first stage rotor blade differs in that the conventional structure in which air is blown out from the trailing edge has been changed to a recovery type, but the basic cooling method is almost the same, and the inflow conditions for cooling air are also exactly the same. Therefore, the same cooling effect as before can be obtained. Since the second stage rotor blades use the exhaust air for cooling the first stage rotor blades, the temperature level of the cooling air increases, but the cross-sectional area of the convection cooling channel increases proportionally to the change in flow rate. As a result, the convective heat transfer coefficient increases in proportion to the 0.4th power of the flow rate increase ratio, making it possible to supply a flow rate that is more than twice the conventional two-stage rotor blade cooling air flow rate. As a result, the cooling efficiency of the entire blade is greatly improved, and the conventional blade temperature level can be maintained. On the other hand, the air flow rate required to cool the dovetail is not directly exposed to the turbine working gas, and since the shank is long in the radial direction, there is a large resistance to heat flow from the high temperature blade section to the dovetail. Since the heat load is very small,
A small amount is sufficient. therefore,? It is possible to reduce the flow rate close to the two-stage rotor blade cooling air flow rate required in the past. !
! In addition, since the flow rate of cooling air is large, there is a small temperature change during cooling of the air, which has the advantage of reducing unevenness in the temperature distribution of the blade surface. Furthermore, when compressed air is supplied to the rotor after passing through an external heat exchanger, the saved cooling air flow rate can be added and supplied to the first stage rotor blades. Since the cooling efficiency is improved and the temperature of the exhaust air is lowered, the cooling effect of the second stage rotor blades is also improved, and if the cooling air flow rate is kept the same as before, the turbine inlet temperature can be raised. On the other hand, in terms of turbine efficiency, the power generated by the second stage rotor blades is reduced by the amount of cooling air no longer discharged by the first stage rotor blades, but this exhaust air does not mix into the working gas. It is more effective to prevent the temperature of the working gas from decreasing due to cooling air, and the power of the compressor and the power of rotating the cooling air at the same time as the rotor, which are offset by the savings in cooling air, are reduced, and the overall turbine Efficiency can be improved.
第5図は本発明の他の実施例を示す第1段動翼100の
縦断面を示し、第2図ないし第4図に示した動翼と異な
るのは矢印101で示すように、冷却空気を外周流路8
2から先に流して中心部流路81に合流して回収するこ
と、及び、第4図の仕切92を無くして孔i02から翼
部に排出するようにしたことであシ、以後の構成は前述
の実施例と全く同様である。この実施例によれば、第3
図に示した吹出孔85によるインピンジ冷却の効果はな
くなるが、仕切シ92を削除した分だけ構造が簡単にな
る利点がある。FIG. 5 shows a longitudinal section of a first stage rotor blade 100 showing another embodiment of the present invention, and the difference from the rotor blades shown in FIGS. 2 to 4 is that cooling air is The outer circumferential flow path 8
2 first, merges into the center channel 81, and collects it, and eliminates the partition 92 in FIG. 4 and discharges from the hole i02 to the wing section. This is exactly the same as the previous embodiment. According to this embodiment, the third
Although the effect of impingement cooling by the blow-off holes 85 shown in the figure is lost, there is an advantage that the structure is simplified by eliminating the partition 92.
なお、図中11は流入孔、80は第一段動翼、88はピ
ンフィン、90はシャンク、92は仕切シである。In the figure, 11 is an inflow hole, 80 is a first stage rotor blade, 88 is a pin fin, 90 is a shank, and 92 is a partition.
なお、第2図ないし第5図に示す第一段動翼の実施例で
は、冷却空気の全部を回収しているが、部分的に翼部面
への吹出し孔を設けて冷却を強化することも本発明の範
囲内である。In the embodiments of the first stage rotor blades shown in Figures 2 to 5, all of the cooling air is recovered, but it is possible to partially provide blow-off holes on the blade surface to strengthen cooling. Also within the scope of this invention.
本発明によれば、少ない冷却空気流量で冷却効率が向上
できるので、効率の高いガスタービンが得られる。According to the present invention, since cooling efficiency can be improved with a small flow rate of cooling air, a highly efficient gas turbine can be obtained.
第1図は従来のガスタービンの部分断面図、第2図は本
発明の冷却流路を構成したガスタービンの部分断面図、
第3図は第2図の■−■矢視断面図、第4図は第2図の
■−■矢視断面図、第5図は本発明の他の実施例の部分
断面図である。
72・・・ガイドリング、80.100・・・第1段動
翼、81.82..97・・・冷却流路、93・・・流
出孔、95・・・第2段動翼、96・・・流入孔。
第1図
一:?ハ
第2 図
第3図
n
環4図
2にFIG. 1 is a partial sectional view of a conventional gas turbine, and FIG. 2 is a partial sectional view of a gas turbine configured with a cooling flow path according to the present invention.
3 is a cross-sectional view taken along the line ■--■ in FIG. 2, FIG. 4 is a cross-sectional view taken along the line ■--■ in FIG. 2, and FIG. 5 is a partial cross-sectional view of another embodiment of the present invention. 72... Guide ring, 80.100... First stage rotor blade, 81.82. .. 97... Cooling channel, 93... Outflow hole, 95... Second stage rotor blade, 96... Inflow hole. Figure 1: ? Figure 2 Figure 3 n Ring 4 Figure 2
Claims (1)
部に冷却流路を設け、燃焼用圧縮空気の一部を前記冷却
流路に供給する空気冷却式のガスタービンにおいて、 前記第一段及び前記第二段動翼の前記冷却流路をシリー
ズに連通して前記第一段動翼の冷却に供給した空気の全
部おるいは一部を回収し、この回収空気を前記第二段動
翼の冷却用として供給するように構成したことを特徴と
する空気冷却式ガスタービン。 2、前記第一段動翼の翼内に形成した前E冷却流路を二
重構造とし、内側及び外側流路を翼先端部あるいは中間
部で連通ずる一方、いずれ力)の他端をダブテイル部分
に開口し、他の一つを前言上第二段動翼側のシャンクの
側面に開口し、自ff9己第二段動翼内の前記冷却流路
の入口を前記第一段動翼*Uのシャンクの側面に開口し
、更へ動翼を支持するディスクの間に在るスペーサの外
周にガイドリンクを設け、このガイドリングと前記スペ
ーサ間に形成されるスペースで前記両動翼の前記シャン
クの側面開口を連通ずるように構成したことを特徴とす
る特許請求の範囲第1項記載の空気冷却式ガスタービン
。[Claims] 1. The rotor blades are made up of multiple stages. 9. Cooling channels are provided inside the first and second stage rotor blades, and a portion of the compressed air for combustion is supplied to the cooling channels. In an air-cooled gas turbine, the cooling flow paths of the first stage and second stage rotor blades are connected in series so that all or part of the air supplied for cooling the first stage rotor blades is An air-cooled gas turbine characterized in that the air is recovered and the recovered air is supplied for cooling the second stage rotor blades. 2. The front E cooling channel formed in the blade of the first stage rotor blade has a double structure, and the inner and outer channels are connected at the tip or middle part of the blade, while the other end is dovetailed. The other one is opened on the side of the shank on the side of the second stage rotor blade, and the inlet of the cooling flow path in the second stage rotor blade is connected to the first stage rotor blade*U. A guide link is provided on the outer periphery of a spacer that opens on the side surface of the shank of the rotor blades and is located between the disks that support the rotor blades, and the space formed between the guide ring and the spacer connects the shanks of both rotor blades. 2. The air-cooled gas turbine according to claim 1, wherein the side openings of the air-cooled gas turbine are configured to communicate with each other.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
JP4888683A JPS59176401A (en) | 1983-03-25 | 1983-03-25 | Air-cooled gas turbine |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
JP4888683A JPS59176401A (en) | 1983-03-25 | 1983-03-25 | Air-cooled gas turbine |
Publications (1)
Publication Number | Publication Date |
---|---|
JPS59176401A true JPS59176401A (en) | 1984-10-05 |
Family
ID=12815756
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
JP4888683A Pending JPS59176401A (en) | 1983-03-25 | 1983-03-25 | Air-cooled gas turbine |
Country Status (1)
Country | Link |
---|---|
JP (1) | JPS59176401A (en) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP2017115881A (en) * | 2015-12-21 | 2017-06-29 | ゼネラル・エレクトリック・カンパニイ | Platform core feed for multi-wall blade |
-
1983
- 1983-03-25 JP JP4888683A patent/JPS59176401A/en active Pending
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP2017115881A (en) * | 2015-12-21 | 2017-06-29 | ゼネラル・エレクトリック・カンパニイ | Platform core feed for multi-wall blade |
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