JPH0953406A - Impingement cooling device and gas turbine blade - Google Patents
Impingement cooling device and gas turbine bladeInfo
- Publication number
- JPH0953406A JPH0953406A JP20698595A JP20698595A JPH0953406A JP H0953406 A JPH0953406 A JP H0953406A JP 20698595 A JP20698595 A JP 20698595A JP 20698595 A JP20698595 A JP 20698595A JP H0953406 A JPH0953406 A JP H0953406A
- Authority
- JP
- Japan
- Prior art keywords
- cooling
- blade
- gas turbine
- impingement
- gap
- 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
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【０００１】[0001]
【発明の属する技術分野】本発明は、インピンジメント
冷却装置及びこの冷却装置を用いたガスタービン翼に関
する。TECHNICAL FIELD The present invention relates to an impingement cooling device and a gas turbine blade using this cooling device.
【０００２】[0002]
【従来の技術】ガスタービン冷却翼は、圧縮空気の一部
を用いて冷却されている。ガスタービン冷却翼では、翼
入口ガス温度に応じて、一般に対流冷却、インピンジメ
ント冷却、フィルム冷却といった冷却が単独又は組合せ
て用いられている。2. Description of the Related Art Gas turbine cooling blades are cooled by using a part of compressed air. In a gas turbine cooling blade, cooling such as convection cooling, impingement cooling, and film cooling is generally used alone or in combination depending on the blade inlet gas temperature.
【０００３】その内の一であるインピンジメント冷却
は、翼の流通面に冷却空気を吹き出して翼面を覆うこと
によって冷却効率を向上させるもので、高温ガスタービ
ン空冷静翼においては欠かすことのできない技術であ
る。Impingement cooling, which is one of them, improves cooling efficiency by blowing cooling air to the flow surface of the blade to cover the blade surface, and is essential for high temperature gas turbine air-cooled stationary blades. It is a technology.
【０００４】図３はインピンジメント冷却を施した従来
のガスタービン冷却翼の断面図である。翼１１の内部
に、内部が中空部１２ａとなっている多孔板のインサー
ト１２を挿入し、翼内部に鋳出しされているシールダム
１３にインサート１２を密着させて固定している。な
お、１２ｂは前記多孔板のインサート１２に設けられた
孔である。FIG. 3 is a cross-sectional view of a conventional gas turbine cooling blade that has been impingement cooled. An insert 12 of a perforated plate having a hollow portion 12a inside is inserted into the blade 11, and the insert 12 is fixed in close contact with a seal dam 13 cast inside the blade. In addition, 12b is a hole provided in the insert 12 of the perforated plate.
【０００５】冷却空気は、インサート１２の内部の中空
部１２ａから多孔板の孔１２ｂを通ってインサート１２
と翼内面とで形成される隙間１５内に流出し、翼１１に
設けられたフィルム冷却穴１４を通って翼冷却面に均一
かつ効果的に吹き出される。The cooling air passes from the hollow portion 12a inside the insert 12 through the holes 12b of the perforated plate to the insert 12.
It flows out into a gap 15 formed by the inner surface of the blade and the inner surface of the blade, and is uniformly and effectively blown out to the blade cooling surface through the film cooling hole 14 provided in the blade 11.
【０００６】図２は、従来のインピンジメント冷却翼の
部分斜視図であり、インサート１２と翼内面１１とで形
成される隙間１５は、隙間内を流れる冷却空気流れの上
流側と下流側とも同じ大きさＺ3 になっている。FIG. 2 is a partial perspective view of a conventional impingement cooling blade, and the gap 15 formed between the insert 12 and the blade inner surface 11 is the same on the upstream side and the downstream side of the cooling air flow flowing in the gap. The size is Z 3 .
【０００７】[0007]
【発明が解決しようとする課題】インピンジメント冷却
においては、インサートと翼内面とで形成される隙間を
流れる冷却空気流れ（クロスフロー）の上流側から下流
に向うに従って図２に示されるように流量Ｗc が増加す
るため、下流側では冷却空気流れの流速が大きくなる。In the impingement cooling, as shown in FIG. 2, the cooling air flow (cross flow) flowing through the gap formed between the insert and the blade inner surface flows from the upstream side to the downstream side as shown in FIG. Since W c increases, the flow velocity of the cooling air flow increases on the downstream side.
【０００８】本発明者の研究の結果、インサートと翼内
面の隙間の大きさがある一定範囲以内であれば、インピ
ンジメント冷却は冷却空気流れの流速に大きく影響さ
れ、冷却空気流れの流速が大きいときにはインピンジメ
ント冷却の効率が低下することが判明した。As a result of the research by the present inventor, if the size of the gap between the insert and the inner surface of the blade is within a certain range, impingement cooling is greatly affected by the flow velocity of the cooling air flow, and the flow velocity of the cooling air flow is large. It has been found that the efficiency of impingement cooling is sometimes reduced.
【０００９】従って、インサートと翼内面で形成される
隙間を流れる冷却空気流れの流速が下流に向うに従って
大きくなる前記の従来のガスタービン冷却翼では、冷却
空気流れの下流側でインピンジメント冷却を効率よく行
うことができないという問題点がある。Therefore, in the conventional gas turbine cooling blade described above, in which the flow velocity of the cooling air flow flowing through the gap formed between the insert and the blade inner surface increases toward the downstream side, the impingement cooling is efficiently performed on the downstream side of the cooling air flow. There is a problem that it cannot be done well.
【００１０】本発明は、以上の問題点を解決するために
なされたものである。The present invention has been made to solve the above problems.
【００１１】[0011]
（１）本発明のインピンジメント冷却装置は、冷却空気
が内部の中空部に供給される多孔板のインサートと被冷
却物内面との隙間を冷却空気流れ下流側に向うにしたが
って大きくなるように形成したことを特徴とする。(1) In the impingement cooling device of the present invention, the gap between the insert of the perforated plate and the inner surface of the object to be cooled, in which cooling air is supplied to the hollow portion inside, is formed so as to increase toward the downstream side of the cooling air flow. It is characterized by having done.
【００１２】（２）また本発明のガスタービン翼は、前
記本発明（１）のインピンジメント冷却装置をガスター
ビン翼の中空冷却部分に適用したことを特徴とする。(2) A gas turbine blade of the present invention is characterized in that the impingement cooling device of the present invention (1) is applied to a hollow cooling portion of the gas turbine blade.
【００１３】本発明では、インサートとガスタービン翼
等の被冷却物内面とで形成される隙間を、同隙間内を流
れる冷却空気流れの下流側に向うにしたがって大きくし
たことによって、下流側において冷却空気流れの流路面
積が大きくなる。従って、冷却空気流れの流量が下流側
で増加しても流速は上流側と同様に保たれる。それによ
って下流側でもインピンジメント冷却の効果が低減する
ことはない。According to the present invention, the gap formed between the insert and the inner surface of the object to be cooled such as the gas turbine blade is enlarged toward the downstream side of the cooling air flow flowing in the gap, so that cooling is performed on the downstream side. The flow path area of the air flow is increased. Therefore, even if the flow rate of the cooling air flow increases on the downstream side, the flow velocity is maintained similar to that on the upstream side. This does not reduce the effect of impingement cooling even on the downstream side.
【００１４】[0014]
【発明の実施の形態】本発明の実施の一形態を、図１に
よって説明する。図１はインピンジメント冷却を施した
ガスタービン冷却翼の部分斜視図である。本実施の形態
は、図３に示されるインピンジメント冷却を施したガス
タービン冷却翼の翼の内面と多孔板のインサートで形成
される隙間を次のように構成したものである。BEST MODE FOR CARRYING OUT THE INVENTION An embodiment of the present invention will be described with reference to FIG. FIG. 1 is a partial perspective view of a gas turbine cooling blade subjected to impingement cooling. In this embodiment, the gap formed by the inner surface of the impingement-cooled gas turbine cooling blade shown in FIG. 3 and the insert of the perforated plate is configured as follows.
【００１５】即ち、翼１１の内面と多孔板のインサート
１２とで形成される隙間１５を、隙間１５を流れる冷却
空気流れ（クロスフロー）の上流側と下流側とで異なら
せていて、下流側の隙間の大きさＺ2 を上流側の隙間の
大きさＺ1 より大きくしている。That is, the gap 15 formed by the inner surface of the blade 11 and the insert 12 of the perforated plate is different between the upstream side and the downstream side of the cooling air flow (cross flow) flowing through the gap 15, and the downstream side. The size Z 2 of the gap is larger than the size Z 1 of the gap on the upstream side.
【００１６】隙間１５を流れる冷却空気流れの流量Ｗc
は下流に向うに従って増加するが、下流側の方が隙間１
５が大きく流路面積が大きいため、流速Ｖc は上流側と
同じに保たれる。これによって、下流側においてインピ
ンジメント冷却の効果が低減することを防止することが
できる。Flow rate W c of the cooling air flowing through the gap 15
Increases as it goes downstream, but the gap 1 is on the downstream side.
Since 5 is large and the flow passage area is large, the flow velocity V c is kept the same as that on the upstream side. This can prevent the effect of impingement cooling on the downstream side from decreasing.
【００１７】なお、前記本発明の実施の一形態は、イン
ピンジメント冷却を施したガスタービン冷却翼に係るも
のであるが、本発明は、ガスタービン翼のシュラウド等
の高温な被冷却物のインピンジメント冷却装置に広く適
用することができる。The embodiment of the present invention relates to a gas turbine cooling blade that is impingement-cooled, but the present invention is to impinge a high temperature object to be cooled, such as a shroud of a gas turbine blade. It can be widely applied to the ment cooling device.
【００１８】[0018]
【発明の効果】本発明のインピンジメント冷却装置及び
同インピンジメント冷却装置を中空冷却部分に適用した
ガスタービン翼によれば、多孔板のインサートと被冷却
物の内面とで形成される隙間を冷却空気流れの下流側で
大きくし、その流速を上流側と下流側とも同様に保つこ
とができる。従って、流速の変化に伴う下流側における
インピンジメント冷却効率の低下を防止することができ
る。According to the impingement cooling device of the present invention and the gas turbine blade in which the impingement cooling device is applied to the hollow cooling portion, the gap formed between the insert of the perforated plate and the inner surface of the object to be cooled is cooled. It is possible to increase the flow rate on the downstream side of the air flow and maintain the same flow rate on the upstream side and the downstream side. Therefore, it is possible to prevent the impingement cooling efficiency on the downstream side from being lowered due to the change of the flow velocity.
【図１】本発明の実施の一形態に係るガスタービン冷却
翼のインピンジメント冷却装置の部分斜視図である。FIG. 1 is a partial perspective view of an impingement cooling device for a gas turbine cooling blade according to an embodiment of the present invention.
【図２】従来のインピンジメント冷却を施したガスター
ビン冷却翼の部分斜視図である。FIG. 2 is a partial perspective view of a conventional gas turbine cooling blade subjected to impingement cooling.
【図３】従来のインピンジメント冷却を施したガスター
ビン冷却翼の断面図である。FIG. 3 is a cross-sectional view of a gas turbine cooling blade that has been subjected to conventional impingement cooling.
１１ 翼 １２ インサート １２ａ 中空部 １２ｂ 孔 １５ 隙間 11 blade 12 insert 12a hollow part 12b hole 15 gap
Claims (2)
孔板のインサートと被冷却物内面との間の隙間を冷却空
気流れ下流側に向うにしたがって大きくなるように形成
したことを特徴とするインピンジメント冷却装置。1. A gap between an insert of a perforated plate, into which cooling air is supplied to a hollow portion inside, and an inner surface of an object to be cooled is formed so as to become larger toward a downstream side of the cooling air flow. Impingement cooling system.
ガスタービン翼の中空冷却部分に適用したことを特徴と
するガスタービン翼。2. A gas turbine blade, wherein the impingement cooling device according to claim 1 is applied to a hollow cooling portion of a gas turbine blade.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP20698595A JPH0953406A (en) | 1995-08-14 | 1995-08-14 | Impingement cooling device and gas turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP20698595A JPH0953406A (en) | 1995-08-14 | 1995-08-14 | Impingement cooling device and gas turbine blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH0953406A true JPH0953406A (en) | 1997-02-25 |
Family
ID=16532275
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP20698595A Pending JPH0953406A (en) | 1995-08-14 | 1995-08-14 | Impingement cooling device and gas turbine blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH0953406A (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2008111441A (en) * | 2006-10-30 | 2008-05-15 | Snecma | Turbomachine turbine shroud sector |
| JP2010038141A (en) * | 2008-08-08 | 2010-02-18 | Mitsubishi Heavy Ind Ltd | Device for cooling high temperature part of gas turbine |
| US11220916B2 (en) | 2020-01-22 | 2022-01-11 | General Electric Company | Turbine rotor blade with platform with non-linear cooling passages by additive manufacture |
| US11242760B2 (en) | 2020-01-22 | 2022-02-08 | General Electric Company | Turbine rotor blade with integral impingement sleeve by additive manufacture |
| US11248471B2 (en) | 2020-01-22 | 2022-02-15 | General Electric Company | Turbine rotor blade with angel wing with coolant transfer passage between adjacent wheel space portions by additive manufacture |
| US11492908B2 (en) | 2020-01-22 | 2022-11-08 | General Electric Company | Turbine rotor blade root with hollow mount with lattice support structure by additive manufacture |
-
1995
- 1995-08-14 JP JP20698595A patent/JPH0953406A/en active Pending
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2008111441A (en) * | 2006-10-30 | 2008-05-15 | Snecma | Turbomachine turbine shroud sector |
| JP2010038141A (en) * | 2008-08-08 | 2010-02-18 | Mitsubishi Heavy Ind Ltd | Device for cooling high temperature part of gas turbine |
| US11220916B2 (en) | 2020-01-22 | 2022-01-11 | General Electric Company | Turbine rotor blade with platform with non-linear cooling passages by additive manufacture |
| US11242760B2 (en) | 2020-01-22 | 2022-02-08 | General Electric Company | Turbine rotor blade with integral impingement sleeve by additive manufacture |
| US11248471B2 (en) | 2020-01-22 | 2022-02-15 | General Electric Company | Turbine rotor blade with angel wing with coolant transfer passage between adjacent wheel space portions by additive manufacture |
| US11492908B2 (en) | 2020-01-22 | 2022-11-08 | General Electric Company | Turbine rotor blade root with hollow mount with lattice support structure by additive manufacture |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| JP3808211B2 (en) | Equipment for ventilation of turbine stator rings | |
| JP4513002B2 (en) | Cooling system for platform edge of nozzle segment | |
| US3807892A (en) | Cooled guide blade for a gas turbine | |
| US5480281A (en) | Impingement cooling apparatus for turbine shrouds having ducts of increasing cross-sectional area in the direction of post-impingement cooling flow | |
| US6435813B1 (en) | Impigement cooled airfoil | |
| EP3124745B1 (en) | Turbo-engine component with film cooled wall | |
| US7334992B2 (en) | Turbine blade cooling system | |
| JPH01232102A (en) | Air-cooling gas turbine blade | |
| JP2001317302A (en) | Film cooling for closed loop cooled airfoil | |
| JPH10220203A (en) | Stationary blade for gas turbine | |
| JPH1089005A (en) | High temperature member cooling device | |
| JPH11148303A (en) | Segment unit for platform | |
| KR20010109466A (en) | Steam exit flow design for aft cavities of an airfoil | |
| JPH09189203A (en) | Gas turbine stator | |
| JPH11257003A (en) | Impingement cooling device | |
| JP4112942B2 (en) | Turbine shroud cooling hole diffuser and associated method | |
| US6328532B1 (en) | Blade cooling | |
| JPH0953406A (en) | Impingement cooling device and gas turbine blade | |
| JPH0777061A (en) | Method and equipment for cooling structural section | |
| JP2003129803A (en) | Gas turbine | |
| JPS62153504A (en) | Shrouding segment | |
| JPH0463901A (en) | Gas turbine cooling blade | |
| JPS61142399A (en) | Method of preventing adhesion of dust to static blade | |
| JP4416387B2 (en) | Methods for improving component life in gas flow | |
| JP3615907B2 (en) | Gas turbine cooling blade |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| A02 | Decision of refusal |
Free format text: JAPANESE INTERMEDIATE CODE: A02 Effective date: 20000919 |