JPH0544691A - Axial flow turbomachinery blade - Google Patents
Axial flow turbomachinery bladeInfo
- Publication number
- JPH0544691A JPH0544691A JP22225091A JP22225091A JPH0544691A JP H0544691 A JPH0544691 A JP H0544691A JP 22225091 A JP22225091 A JP 22225091A JP 22225091 A JP22225091 A JP 22225091A JP H0544691 A JPH0544691 A JP H0544691A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- fillet
- curvature
- flow
- point
- 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.)
- Withdrawn
Links
Landscapes
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
Description
【０００１】[0001]
【産業上の利用分野】本発明は、軸流型のファン、圧縮
機、タービン、ポンプ等の軸流ターボ機械の動翼及び静
翼に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a moving blade and a stationary blade of an axial flow turbomachine such as an axial flow type fan, compressor, turbine and pump.
【０００２】[0002]
【従来の技術】従来の軸流ターボ機械の翼の例を図５及
び図６に示す。図５は翼の斜視図、図６はその側面図で
ある。本例は動翼に関する図であるが、静翼の場合も同
様であるので、動翼についてのみ説明する。これらの図
において、１は翼、２ははダブテール、３はダブテール
２のプラットフォーム、４は翼１における翼背面５は翼
１における翼腹面、６は翼１の後縁、７は翼１の前縁、
８は翼背面４とプラットフォーム３とを滑らかに接続す
るフィレット（丸み）をそれぞれ示している。これが動
翼の通常の形態であり、静翼の場合にはプラットフォー
ム３がケーシング壁面又はシュラウド環の面に相当す
る。2. Description of the Related Art An example of a blade of a conventional axial flow turbomachine is shown in FIGS. FIG. 5 is a perspective view of the wing, and FIG. 6 is a side view thereof. Although this example is a diagram relating to a moving blade, the same applies to the case of a stationary blade, so only the moving blade will be described. In these figures, 1 is a wing, 2 is a dovetail, 3 is a platform of a dovetail 2, 4 is a wing back surface of wing 1, 5 is a wing vent surface of wing 1, 6 is a trailing edge of wing 1, and 7 is a front of wing 1. edge,
Reference numerals 8 respectively indicate fillets (roundness) that smoothly connect the wing rear surface 4 and the platform 3. This is a normal form of a moving blade, and in the case of a stationary blade, the platform 3 corresponds to the casing wall surface or the surface of the shroud ring.
【０００３】ここで着目しているのはフィレット８であ
る。フィレット８の役目は、翼１を応力集中が発生せず
に保持することと、コーナ部に発生する渦流れを緩和し
て二次損失の一部を減らすことにある。The focus here is on the fillet 8. The roles of the fillet 8 are to hold the blade 1 without stress concentration and to reduce a part of the secondary loss by relaxing the vortex flow generated at the corners.
【０００４】[0004]
【発明が解決しようとする課題】圧縮機において動翼や
静翼によりなされる仕事のうち、ある部分は損失として
消費される。その例を、段効率における割合として図７
に示す。その損失の内訳は、大まかには、プロファイル
損失、アニュラス損失、及びそれ以外の損失に分けられ
る。プロファイル損失とは、主要部分の流線に沿う翼断
面で生ずる摩擦損失等、二次元的な損失である。また、
アニュラス損失は、ケーシングとハブとの各面の摩擦損
失に相当する。一方、それ以外の損失をまとめて二次損
失と呼んでいる。これは翼端すきまからの漏れ流れやコ
ーナ部で発生する渦、流路の曲りにより発生する流路渦
等、主として渦となってその運動エネルギを回収できな
い流れ損失である。この図から分かるように、二次損失
は全損失の１／２〜１／３程度の大きな割合を占めてい
る。A part of the work performed by the moving blades and the stationary blades in the compressor is consumed as a loss. An example of this is shown in FIG.
Shown in. The breakdown of the loss is roughly divided into profile loss, annulus loss, and other loss. The profile loss is a two-dimensional loss such as a friction loss generated in a blade section along the streamline of the main part. Also,
The annulus loss corresponds to the friction loss on each surface of the casing and the hub. On the other hand, the other losses are collectively called the secondary loss. This is a flow loss in which kinetic energy cannot be recovered mainly as vortices such as leakage flow from the blade tip clearance, vortices generated at the corners, and flow channel vortices generated by bending of the flow channel. As can be seen from this figure, the secondary loss occupies a large proportion of about 1/2 to 1/3 of the total loss.
【０００５】軸流圧縮機の動翼における翼間あるいは翼
と壁間との間の流れの特徴を図８に示す。ここに示され
た流れの細部には各々それなりの名称がつけられている
が、設計流れからの逸脱を示す点で、ここではそれらを
まとめて二次流れと呼んでいる。FIG. 8 shows the characteristics of the flow between the blades or between the blades and the walls in the moving blades of the axial compressor. Each of the details of the flow shown here is given a proper name, but they are collectively referred to as a secondary flow here because they show deviations from the design flow.
【０００６】最近では、翼の断面プロファイルの性態は
ほぼ計画どおりに達成されているが、これら二次流れは
まだ制御し難い点が多い。ターボ機械の効率を更に向上
させるには、この二次損失の低減を図ることが必要にな
りつつある。Recently, the characteristics of the blade cross-sectional profile have been achieved almost as planned, but these secondary flows are still difficult to control. In order to further improve the efficiency of turbomachines, it is becoming necessary to reduce this secondary loss.
【０００７】一方、最近では、流れの測定や可視化の技
術が向上し、二次流れに関するデータが蓄積され、その
イメージが把握されてきている。また、数値計算流体力
学等、理論計算法の高度化などにより、詳しい検討も可
能になりつつある。これらにより、二次流れを制御し、
その損失を低減する試みが始められつつある。On the other hand, recently, techniques for flow measurement and visualization have been improved, and data concerning secondary flows have been accumulated and the image thereof has been grasped. Further, due to the sophistication of theoretical calculation methods such as computational fluid dynamics, detailed examination is becoming possible. These control the secondary flow,
Attempts are being made to reduce that loss.
【０００８】現在行われている二次損失低減の手法は、
軸流タービンの例で見ると、翼断面の積重ね方を工夫
し、翼高さ方向の静圧分布を制御して二次流れを減らす
かアニュラス壁面の絞り方を工夫して翼面の間のコーナ
の渦を抑制するもので、主として静翼に適用されること
が多い。[0008] The currently used secondary loss reduction method is as follows.
Looking at an example of an axial flow turbine, the blade stacking method is devised to control the static pressure distribution in the blade height direction to reduce the secondary flow, or the annulus wall surface is throttled to improve the It suppresses corner vortices and is often applied mainly to vanes.
【０００９】図８に示される各種の二次的な流れの低減
のため、様々なアプローチを行っている。ここで、特
に、ハブコーナ渦について着目してみると、ハブコーナ
渦の発生の原因は、図９に示す如く、翼腹面５の上側で
の高い圧力と、翼背面４の上側での低い圧力のため、プ
ラットフォーム３の上を通る流れがプラットフォーム３
を経て、破線９で示す如く、翼背面４の上に巻き上がっ
て行くことによる。この流れ９は翼背面４上の境界層と
干渉して流れを劣化させ、大きな圧力損失をハブ付近に
発生させるのである。Various approaches have been taken to reduce the various secondary flows shown in FIG. Here, paying particular attention to the hub corner vortex, the cause of the hub corner vortex is that the high pressure on the upper side of the blade surface 5 and the low pressure on the upper side of the blade rear surface 4 as shown in FIG. , The flow passing over the platform 3 is the platform 3
And by rolling up over the blade back surface 4 as indicated by the broken line 9. This flow 9 interferes with the boundary layer on the blade back surface 4 to deteriorate the flow and causes a large pressure loss near the hub.
【００１０】本発明は二次流れのうち、ハブコーナ渦の
抑制と、それに起因する損失の低減を図った翼を提供す
ることを目的とする。It is an object of the present invention to provide a blade that suppresses hub corner vortex in the secondary flow and reduces loss caused by it.
【００１１】[0011]
【課題を解決するための手段】上記目的に対し、本発明
によれば、翼とプラットフォーム又はケーシング壁面と
の接合部にてこれらを翼背面で滑らかに接続するフィレ
ットを、翼の前縁から背面最大反り点を越したある点ま
では曲率半径が大きく、そのある点から翼の後縁までは
曲率半径が小さく、かつそのある点における曲率半径は
急激に変化させてなる軸流ターボ機械翼が提供される。For the above object, according to the present invention, a fillet that smoothly connects the blade and the platform or the casing wall surface at the joint between the blade and the blade is provided from the leading edge to the back of the blade. The radius of curvature is large up to a certain point beyond the maximum warp point, the radius of curvature is small from that point to the trailing edge of the blade, and the radius of curvature at that point is abruptly changed. Provided.
【００１２】[0012]
【作用】上記手段によれば、曲率半径の急変更点は下流
向きのステップであるため、ここで流れは剥離し、急変
更点より後縁までのフィレットに局部的な負圧が発生す
る。ハブコーナ渦に伴う壁面沿いの流れはその負圧部分
に取り込まれて吸収されることになる。According to the above-mentioned means, since the point of sudden change of the radius of curvature is the step in the downstream direction, the flow is separated at this point, and a local negative pressure is generated in the fillet from the point of sudden change to the trailing edge. The flow along the wall surface due to the hub corner vortex is taken in by the negative pressure portion and absorbed.
【００１３】[0013]
【実施例】以下、図１ないし図４を参照し、本発明の好
適な実施例について詳述する。図１は本発明による翼の
斜視図、図２はその平面図、図３はその側面図、図４は
流れの説明図であり、これらの図において、図５及び図
６に示したものと同一の要素については同一の符号を付
してある。なお、符号１０は背面前部付根フィレット、
１１は背面後部付根フィレット、１２は背面前部付根フ
ィレット１０の丸み上端、１３はその丸み下端、１４は
背面後部付根フィレット１１の丸み上端、１５はその丸
み下端、１６は背面前部付根フィレット１０と背面後部
付根フィレット１１との境界の急変更点、１７は負圧
面、１８及び１９は流れを示している。DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT A preferred embodiment of the present invention will be described in detail below with reference to FIGS. FIG. 1 is a perspective view of a wing according to the present invention, FIG. 2 is a plan view thereof, FIG. 3 is a side view thereof, and FIG. 4 is an explanatory view of a flow. The same elements are designated by the same reference numerals. In addition, reference numeral 10 is a fillet having a root on the back side,
Reference numeral 11 is a back rear root fillet, 12 is a rounded upper end of the back front root fillet 10, 13 is a rounded lower end thereof, 14 is a rounded upper end of the back rear root fillet 11, 15 is a rounded lower end thereof, 16 is a back front root fillet 10. Shows a sudden change in the boundary between the rear root joint fillet 11 and the rear part, 17 is a suction surface, and 18 and 19 are flows.
【００１４】翼背面プロファイルのフィレットの前半部
分である背面前部付根フィレット１０は曲率半径ｒ1を
大きくとり、後半部分の背面後部付根フィレット１１は
曲率半径ｒ2を小さくとり、かつ急変更点１６にてｒ1か
らｒ2への変化を急に行っている。The back front root fillet 10, which is the front half of the fillet of the blade back profile, has a large radius of curvature r 1, and the back rear root fillet 11 of the latter half has a small radius of curvature r 2 and has a sudden change point 16. The change from r 1 to r 2 is suddenly made.
【００１５】急変更点１６は、応力集中のため剥離を起
こす可能性のある背面最大反り点のごく近くを避けてそ
の最大反り点と後縁６との間で決定され、最も適切な位
置は実験的に最適化される。また、曲率半径ｒ1及びｒ2
の大きさも実験的に最適に設定される。The abrupt change point 16 is determined between the maximum warp point and the trailing edge 6 avoiding the vicinity of the maximum warp point of the back surface which may cause separation due to stress concentration, and the most suitable position is Optimized experimentally. Also, the radii of curvature r 1 and r 2
The size of is also experimentally set optimally.
【００１６】作用について説明すると、図１〜３に示す
背面前部付根フィレット１０及び背面後部付根フィレッ
ト１１の急変更点１６は下流向きのステップであるた
め、ここでは流れは剥離され、急変更点１６の面に接し
て、逆流を伴なう強く乱れた流れが発生し、大きな負圧
が局部的に発生する。The operation will be described. Since the abrupt change point 16 of the back front root fillet 10 and the back rear root fillet 11 shown in FIGS. 1 to 3 is a downstream step, the flow is separated here and the abrupt change point is shown. A strongly turbulent flow accompanied by backflow is generated in contact with the surface 16 and a large negative pressure is locally generated.
【００１７】この負圧は周囲の圧力より低いため、図８
に示すハブコーナ渦に伴う壁面沿いの流れはこの急変更
点１６の下流の負圧面１７に取り込まれ、吸収されて、
翼背面４上によじ登って行き難くなる。即ち、図４にお
ける流れ１８，１９の如く、プラットフォーム３の面と
翼面に沿って負圧面２５に引込まれる。Since this negative pressure is lower than the ambient pressure, FIG.
The flow along the wall surface due to the hub corner vortex shown in is taken in by the suction surface 17 downstream of the sudden change point 16 and absorbed,
It becomes difficult to climb on the back surface 4 of the wing. That is, like the flows 18 and 19 in FIG. 4, the suction surface 25 is drawn along the surface of the platform 3 and the blade surface.
【００１８】したがって、翼背面付根には、細い負圧部
分が発生するが、これが障壁となって本来のハブコーナ
渦はくい止められ（図９の流れ９と図４の流れ１９とは
翼背面４上では逆方向となっている）、それと翼背面４
上の境界層との干渉によって生ずる翼面流れの劣化と、
それに起因する大きな圧力損失は抑止される。すなわ
ち、二次損失の幾許かは低減される。この作用はある点
でボルテックスゼネレータの作用に似ている。Therefore, a thin negative pressure portion is generated at the root of the blade rear surface, and this serves as a barrier to block the original hub corner vortex (the flow 9 in FIG. 9 and the flow 19 in FIG. 4 are on the blade rear surface 4). In the opposite direction), and the wing back 4
Deterioration of airfoil flow caused by interference with the upper boundary layer,
The large pressure loss resulting from this is suppressed. That is, some of the secondary loss is reduced. This action is similar in some respects to that of a vortex generator.
【００１９】[0019]
【発明の効果】本発明によれば、ハブコーナ渦の発生が
抑制されたことにより、これに起因した二次損失が低減
され、効率向上を図ることができる。According to the present invention, since the generation of hub corner vortices is suppressed, the secondary loss resulting from this is reduced, and the efficiency can be improved.
【図１】本発明の実施例に係る動翼の斜視図である。FIG. 1 is a perspective view of a moving blade according to an embodiment of the present invention.
【図２】図１の動翼の平面図である。2 is a plan view of the moving blade of FIG. 1. FIG.
【図３】図１の動翼の側面図である。3 is a side view of the rotor blade of FIG. 1. FIG.
【図４】本発明の原理説明図である。FIG. 4 is a diagram illustrating the principle of the present invention.
【図５】従来の動翼の斜視図である。FIG. 5 is a perspective view of a conventional moving blade.
【図６】図５の動翼の側面図である。6 is a side view of the rotor blade of FIG.
【図７】圧縮機における段効率−流量係数を示す図であ
る。FIG. 7 is a diagram showing a stage efficiency-flow coefficient in a compressor.
【図８】圧縮機における二次的な流れを示す説明図であ
る。FIG. 8 is an explanatory diagram showing a secondary flow in the compressor.
【図９】ハブコーナ渦の発生を説明するための図であ
る。FIG. 9 is a diagram for explaining generation of a hub corner vortex.
１ 翼 ３ プラットフォーム ４ 翼背面 １０ 背面前部付根フィレット １１ 背面後部付根フィレット １６ 急変更点 1 wing 3 platform 4 wing back surface 10 back front root fillet 11 back rear root fillet 16 sudden changes
Claims (1)
との接合部にてこれらを翼背面で滑らかに接続するフィ
レットを、翼の前縁から背面最大反り点を越したある点
までは曲率半径が大きく、そのある点から翼の後縁まで
は曲率半径が小さく、かつそのある点における曲率半径
は急激に変化させてなる軸流ターボ機械翼。1. A fillet that smoothly connects a blade and a platform or a casing wall surface at a joint between the blade and the back surface of the blade has a large radius of curvature from a front edge of the blade to a point beyond the maximum back warpage point, An axial-flow turbomachine blade with a small radius of curvature from that point to the trailing edge of the blade, and with a sudden change in the radius of curvature at that point.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP22225091A JPH0544691A (en) | 1991-08-07 | 1991-08-07 | Axial flow turbomachinery blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP22225091A JPH0544691A (en) | 1991-08-07 | 1991-08-07 | Axial flow turbomachinery blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH0544691A true JPH0544691A (en) | 1993-02-23 |
Family
ID=16779452
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP22225091A Withdrawn JPH0544691A (en) | 1991-08-07 | 1991-08-07 | Axial flow turbomachinery blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH0544691A (en) |
Cited By (14)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH10122192A (en) * | 1996-10-14 | 1998-05-12 | Ishikawajima Harima Heavy Ind Co Ltd | Axial flow compressor moving blade |
| US5947683A (en) * | 1995-07-11 | 1999-09-07 | Mitsubishi Heavy Industries, Ltd. | Axial compresssor stationary blade |
| JP2001090691A (en) * | 1999-09-23 | 2001-04-03 | General Electric Co <Ge> | Flow passage for pre-scroll of stress-reduced compressor |
| WO2005042925A1 (en) * | 2003-10-31 | 2005-05-12 | Kabushiki Kaisha Toshiba | Turbine cascade structure |
| JP2010196625A (en) * | 2009-02-26 | 2010-09-09 | Mitsubishi Heavy Ind Ltd | Turbine blade and gas turbine |
| WO2011085721A3 (en) * | 2010-01-16 | 2011-10-20 | Mtu Aero Engines Gmbh | Rotor blade for a turbomachine and turbomachine |
| JP2013144986A (en) * | 2013-03-13 | 2013-07-25 | Mitsubishi Heavy Ind Ltd | Turbine blade, and gas turbine |
| WO2014160215A1 (en) * | 2013-03-13 | 2014-10-02 | United Technologies Corporation | Rotor blade with a conic spline fillet at an intersection between a platform and a neck |
| JP2016527431A (en) * | 2013-07-15 | 2016-09-08 | ユナイテッド テクノロジーズ コーポレイションＵｎｉｔｅｄ Ｔｅｃｈｎｏｌｏｇｉｅｓ Ｃｏｒｐｏｒａｔｉｏｎ | Turbine vane with variable fillet |
| US9915206B2 (en) | 2013-03-15 | 2018-03-13 | United Technologies Corporation | Compact aero-thermo model real time linearization based state estimator |
| WO2018147162A1 (en) * | 2017-02-07 | 2018-08-16 | 株式会社Ihi | Blade of axial flow machine |
| EP3711945A1 (en) * | 2017-10-23 | 2020-09-23 | MTU Aero Engines GmbH | Blade and rotor for a turbomachine and turbomachine |
| JP2020153320A (en) * | 2019-03-20 | 2020-09-24 | 三菱日立パワーシステムズ株式会社 | Turbine blade and gas turbine |
| JP2021116739A (en) * | 2020-01-27 | 2021-08-10 | 三菱パワー株式会社 | Turbine blade |
-
1991
- 1991-08-07 JP JP22225091A patent/JPH0544691A/en not_active Withdrawn
Cited By (33)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5947683A (en) * | 1995-07-11 | 1999-09-07 | Mitsubishi Heavy Industries, Ltd. | Axial compresssor stationary blade |
| JPH10122192A (en) * | 1996-10-14 | 1998-05-12 | Ishikawajima Harima Heavy Ind Co Ltd | Axial flow compressor moving blade |
| JP2001090691A (en) * | 1999-09-23 | 2001-04-03 | General Electric Co <Ge> | Flow passage for pre-scroll of stress-reduced compressor |
| WO2005042925A1 (en) * | 2003-10-31 | 2005-05-12 | Kabushiki Kaisha Toshiba | Turbine cascade structure |
| US7625181B2 (en) | 2003-10-31 | 2009-12-01 | Kabushiki Kaisha Toshiba | Turbine cascade structure |
| JP2010196625A (en) * | 2009-02-26 | 2010-09-09 | Mitsubishi Heavy Ind Ltd | Turbine blade and gas turbine |
| WO2011085721A3 (en) * | 2010-01-16 | 2011-10-20 | Mtu Aero Engines Gmbh | Rotor blade for a turbomachine and turbomachine |
| US9482099B2 (en) | 2010-01-16 | 2016-11-01 | Mtu Aero Engines Gmbh | Rotor blade for a turbomachine and turbomachine |
| US9932834B2 (en) | 2013-03-13 | 2018-04-03 | United Technologies Corporation | Rotor blade with a conic spline fillet at an intersection between a platform and a neck |
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