JP2003120202A - Radial turbine rotor blade - Google Patents
Radial turbine rotor bladeInfo
- Publication number
- JP2003120202A JP2003120202A JP2001318577A JP2001318577A JP2003120202A JP 2003120202 A JP2003120202 A JP 2003120202A JP 2001318577 A JP2001318577 A JP 2001318577A JP 2001318577 A JP2001318577 A JP 2001318577A JP 2003120202 A JP2003120202 A JP 2003120202A
- Authority
- JP
- Japan
- Prior art keywords
- hub
- working gas
- radial turbine
- blade
- turbine rotor
- 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
- Supercharger (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【０００１】[0001]
【発明の属する技術分野】本発明は、小型ガスタービ
ン、過給機、エキスパンダ等に適用されるラジアルター
ビン（斜流タービンも含む）に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a radial turbine (including a mixed flow turbine) applied to small gas turbines, superchargers, expanders and the like.
【０００２】[0002]
【従来の技術】図７は従来のラジアルタービン動翼の一
部を示した斜視図、図８は同ラジアルタービン動翼の翼
間中央部の軸を含む断面図（図７の二点鎖線にて切断し
た断面図）である。これらの図において、０２は作動ガ
ス入口、０３は作動ガスの流れ、０４ａ、０４ｂは翼、
０５は作動ガス出口、０７はハブ背面、０１０はハブ面
（ハブの外周側の面）、０３０はハブ、０３１はハブの
背部、０２０は回転の軸を示している。なお、ハブの背
部０３１とは、ハブ０３０におけるハブ背面０７近傍部
（ハブ背面を含む）を指している。エンジン排ガス等の
高温ガスは、図示しないスクロール等から作動ガス入口
０２へ回転中心の向きに流入し、ハブ面０１０に沿いな
がら翼０４ａ、０４ｂ間の流路へ供給される。そして、
同流路を矢印０３の向きに流れる間に膨張し、翼０４ａ
に回転仕事を与える。通常、ラジアルタービン動翼のハ
ブ背部０３１における外周縁の部分は作動ガス入口０２
へ流入した高温ガスによる熱影響を直に受けることか
ら、熱応力の集中を避けるため、図７に示すように翼０
４ａ、０４ｂ間の外縁部０９が切除された構成をしてお
り、回転中心に向かって凹の形状となっている。2. Description of the Related Art FIG. 7 is a perspective view showing a part of a conventional radial turbine rotor blade, and FIG. 8 is a cross-sectional view including a shaft of a central portion between blades of the radial turbine rotor blade (indicated by a two-dot chain line in FIG. FIG. In these figures, 02 is a working gas inlet, 03 is a working gas flow, 04a and 04b are blades,
Reference numeral 05 is a working gas outlet, 07 is a hub rear surface, 010 is a hub surface (outer peripheral surface of the hub), 030 is a hub, 031 is a back portion of the hub, and 020 is a rotation axis. The hub back portion 031 is a portion of the hub 030 near the hub back surface 07 (including the hub back surface). High-temperature gas such as engine exhaust gas flows from a scroll or the like (not shown) to the working gas inlet 02 in the direction of the center of rotation, and is supplied to the flow path between the blades 04a and 04b along the hub surface 010. And
It expands while flowing in the same flow path in the direction of arrow 03, and blades 04a
Give a spinning job to. Normally, the outer peripheral edge portion of the hub back portion 031 of the radial turbine blade is the working gas inlet 02.
In order to avoid the concentration of thermal stress, the blade 0
The outer edge portion 09 between 4a and 04b is cut off and has a concave shape toward the center of rotation.
【０００３】[0003]
【発明が解決しようとする課題】上記のように、ラジア
ルタービン動翼のハブ背部０３１は、その外縁が回転中
心に向って凹の形状となっている。従って、ラジアルタ
ービン動翼を回転軸を含む面にて切断（この場合は、回
転軸を通り図７の二点鎖線にて切断）したときの様子を
図８に示すと、ハブ背部０３１の縁には半径方向に対し
ほぼ垂直な端面０７ａ０１２が形成されることとなる。
しかしながら、このような端面が形成されると、本来、
作動ガス入口０２から作動ガス出口０５へ滑らかに流れ
るはずのガス流れが、ハブ面０１０に沿う部分で矢印０
１１のように乱れるため、作動効率が低下することとな
る。As described above, the hub back portion 031 of the radial turbine blade has an outer edge which is concave toward the center of rotation. Therefore, when the radial turbine blade is cut along the plane including the rotation axis (in this case, cut along the two-dot chain line of FIG. 7 that passes through the rotation axis), the edge of the hub back 031 is shown in FIG. Thus, an end face 07a012 that is substantially perpendicular to the radial direction is formed on the.
However, when such an end face is formed, originally,
The gas flow that should flow smoothly from the working gas inlet 02 to the working gas outlet 05 is indicated by the arrow 0 at the portion along the hub surface 010.
Since it is disturbed as shown by 11, the operating efficiency is reduced.
【０００４】また、図９に作動ガス入口０２から作動ガ
ス出口０５までの流路における子午面上の流路長さに対
する流路面積の分布を示すが、これによるとハブ背部０
３１の端面０７ａ０１２が形成された部分の近傍で作動
ガス入口０２から作動ガス出口０５までの流路面積が不
連続となるため（図８のＩＩ、ＩＩＩ部分）、翼０４
ａ、０４ｂ間を流れる高温ガスの膨張が不連続となって
内部損失の増加をもたらすこととなる。FIG. 9 shows the distribution of the flow passage area with respect to the flow passage length on the meridian plane in the flow passage from the working gas inlet 02 to the working gas outlet 05.
Since the flow passage area from the working gas inlet 02 to the working gas outlet 05 is discontinuous in the vicinity of the portion where the end face 07a012 of 31 is formed (II and III portions in FIG. 8), the blade 04
The expansion of the high temperature gas flowing between a and 04b becomes discontinuous and causes an increase in internal loss.
【０００５】[0005]
【課題を解決するための手段】上述した課題を解決する
ため、本願は以下のような手段を提供するものである。
（１）ハブ背部の外周縁が各翼間において回転中心に向
かい凹となるように形成されているラジアルタービン動
翼であって、当該外周縁を形成するハブ背部の外周側の
端部が回転軸を含む断面で見たときに鋭角状となるよう
に形成されていることを特徴背板の外周側の縁部が回転
中心に向かって凹となるように形成されているラジアル
タービン動翼において、同縁部の軸を含む断面でのハブ
面と背板の半径線との成す形状が鋭角状に形成されてい
ることを特徴とする。
（２）ハブと隣接する翼とで作動ガスの流路が形成され
たラジアルタービン動翼であって、作動ガス入口から作
動ガス出口までの流路面積が、連続的に増加するよう形
成されたことを特徴作動ガスの流路がハブと背板と隣接
する翼とで囲まれた範囲にて形成されるラジアルタービ
ン動翼であって、作動ガス入口から作動ガス出口までの
流路面積が常に一定の増加割合となるように形成された
ことを特徴とする。
（３）ハブと背板と隣接する翼とで作動ガスの流路が形
成されたラジアルタービン動翼において、背板、ハブ及
び翼で囲まれた隅に、作動ガス流路に対し滑らかな面を
形成する覆い部が形成されていることを特徴作動ガスの
流路がハブと背板と隣接する翼とで囲まれた範囲にて形
成されるラジアルタービン動翼において、背板、ハブ及
び翼で囲まれた隅に、作動ガス流路に対し滑らかな面を
形成するように肉盛部が形成されていることを特徴とす
る。なお、ここで、ラジアルタービン動翼とは、いわゆ
る斜流タービン動翼も含むものである。In order to solve the above-mentioned problems, the present application provides the following means. (1) A radial turbine rotor blade is formed such that the outer peripheral edge of the hub back portion is concave between the blades toward the center of rotation, and the end portion on the outer peripheral side of the hub back portion that forms the outer peripheral edge rotates. It is characterized in that it is formed to have an acute angle when viewed in a cross section including the shaft.In a radial turbine blade in which the outer peripheral edge of the back plate is concave toward the center of rotation The shape formed by the hub surface and the radial line of the back plate in a cross section including the axis of the same edge portion is formed into an acute angle. (2) A radial turbine moving blade in which a working gas passage is formed between the hub and the adjacent blade, and the passage area from the working gas inlet to the working gas outlet is formed to continuously increase. It is a radial turbine moving blade in which the flow path of the working gas is formed in a range surrounded by the hub and the blade adjacent to the back plate, and the flow path area from the working gas inlet to the working gas outlet is always It is characterized in that it is formed to have a constant increase rate. (3) In a radial turbine rotor blade in which a working gas passage is formed by a hub, a blade adjacent to a back plate, and a smooth surface for the working gas passage at a corner surrounded by the back plate, the hub and the blade. In a radial turbine rotor blade in which a flow path of working gas is formed in a range surrounded by a hub, a back plate and an adjacent blade, a back plate, a hub and a blade are provided. It is characterized in that a built-up portion is formed in a corner surrounded by so as to form a smooth surface with respect to the working gas passage. Here, the radial turbine moving blade also includes a so-called mixed flow turbine moving blade.
【０００６】そして、上記手段は以下のような作用を奏
する。
（１）作動ガス入口からラジアルタービン動翼に流入し
た作動ガスは、背板ハブ面に沿いながら翼間の流路へ供
給されるが、その際、ハブ背部背板の縁が鋭角状に形成
されていことから、ガス流れに乱れが生じない。
（２）作動ガス入口から作動ガス出口までの流路面積が
連続であることから、翼間を流れる高温ガスの膨張が連
続となる。
（３）作動ガス入口からラジアルタービン動翼に流入し
た作動ガスは、肉盛覆い部の表面に沿ってスムーズに流
れることとなる。The above-mentioned means have the following effects. (1) The working gas flowing into the radial turbine rotor blade from the working gas inlet is supplied to the flow path between the blades along the back plate hub surface. At that time, the edge of the hub back plate is formed into an acute angle. Therefore, there is no turbulence in the gas flow. (2) Since the flow passage area from the working gas inlet to the working gas outlet is continuous, the expansion of the hot gas flowing between the blades is continuous. (3) The working gas that has flowed into the radial turbine blade from the working gas inlet flows smoothly along the surface of the overlay covering portion.
【０００７】[0007]
【発明の実施の形態】図１は本発明の第１の実施形態に
おけるラジアルタービン動翼の一部を示した斜視図、図
２は同ラジアルタービン動翼の翼近傍をラジアルタービ
ン動翼の翼中央部の軸を含む断面で切断した断面図（図
１の一点鎖線にて切断した断面図）である。これらの図
において、２は作動ガス入口、３は作動ガスの流れ、４
ａ、４ｂは翼、５は作動ガス出口、１０はハブ面、３０
はハブ、３１はハブの背部、３２は背板、２０は回転の
軸を示している。なお、背板３２とは、図１に示すよう
に、ハブ背部３１から翼４ａ、４ｂに向けて延設された
板状の部材を指している。本実施の形態においては、図
２に示すようにラジアルタービン動翼のハブ背部３１に
おける縁の部分は、ハブ３０の半径方向に対して垂直な
面を形成することなく角度θの鋭角状に形成されてい
る。これにより、図２に示すように作動ガスの流れ１１
に生ずる乱れが少なくなり、作動効率の低下が抑制され
る。また、図２において、図６と同様の位置にＩＩ−Ｉ
Ｉ線、ＩＩＩ−ＩＩＩ線を施すと、対応する図３（ｂ）
に示すように、作動ガス入口２から作動ガス出口５まで
の流路面積が不連続となる状態も従来の場合と比べ緩和
され、翼４ａ、４ｂ間における作動ガスの内部損失が軽
減される。なお、ここで、鋭角状とは、背板７の縁部分
が尖った形成になっている場合だけでなく、縁部分の先
端が丸みを帯びていて全体として鋭角に形成されている
場合も含む。DESCRIPTION OF THE PREFERRED EMBODIMENTS FIG. 1 is a perspective view showing a part of a radial turbine moving blade according to a first embodiment of the present invention, and FIG. 2 is a view showing the vicinity of the radial turbine moving blade of the radial turbine moving blade. It is sectional drawing cut | disconnected by the cross section containing the axis | shaft of a center part (sectional view cut | disconnected by the dashed-dotted line of FIG. 1). In these figures, 2 is a working gas inlet, 3 is a working gas flow, 4
a, 4b are blades, 5 is a working gas outlet, 10 is a hub surface, 30
Is a hub, 31 is the back of the hub, 32 is the back plate, and 20 is the axis of rotation. The back plate 32 is a plate-shaped member extending from the hub back portion 31 toward the wings 4a and 4b, as shown in FIG. In the present embodiment, as shown in FIG. 2, the edge portion of the hub back portion 31 of the radial turbine blade is formed into an acute angle with an angle θ without forming a surface perpendicular to the radial direction of the hub 30. Has been done. As a result, as shown in FIG.
The turbulence that occurs in 1) is reduced, and the decrease in operating efficiency is suppressed. Further, in FIG. 2, II-I is located at the same position as in FIG.
When the I line and the III-III line are applied, the corresponding FIG.
As shown in, the state in which the flow passage area from the working gas inlet 2 to the working gas outlet 5 is discontinuous is also relaxed as compared with the conventional case, and the internal loss of the working gas between the blades 4a and 4b is reduced. Here, the acute shape includes not only the case where the edge portion of the back plate 7 is formed to be sharp, but also the case where the tip of the edge portion is rounded and formed as an overall acute angle. .
【０００８】図４３は本発明の第２の実施形態における
ラジアルタービン動翼の一部を示した斜視図、図５４は
同ラジアルタービン動翼の翼近傍をラジアルタービン動
翼の翼中央部の軸を含む断面図である。本実施の形態に
おいては、ラジアルタービン動翼の背板３２、ハブ面１
０及び翼４ａ（又は４ｂ）で囲まれた隅を肉盛りして覆
い部１０８を形成しに、作動ガス入口２から作動ガス出
口５へ向かうガス流路に対し滑らかな面を構成するよう
にしている。備えた覆い部１０８が形成されている。こ
れにより、図４に示すように作動ガスは肉盛部覆い部１
０８の表面に沿ってスムーズに流れるようになり、作動
効率が大きく向上する。また、作動ガス入口２から作動
ガス出口５までの流路面積を子午面上の流路長さに対し
て連続な状態にすることができるため（図３（ｃ））、
翼４ａ、４ｂ間における作動ガスの内部損失が大幅に軽
減される。FIG. 43 is a perspective view showing a part of the radial turbine moving blade according to the second embodiment of the present invention, and FIG. 54 shows the vicinity of the radial turbine moving blade and the shaft of the central portion of the radial turbine moving blade. It is sectional drawing containing. In the present embodiment, the back plate 32 of the radial turbine blade and the hub surface 1
0 and the corners surrounded by the blades 4a (or 4b) are built up to form a covering portion 108, so that a smooth surface is formed with respect to the gas flow path from the working gas inlet 2 to the working gas outlet 5. ing. The provided cover portion 108 is formed. As a result, as shown in FIG.
It smoothly flows along the surface of 08, greatly improving the operating efficiency. Further, since the flow passage area from the working gas inlet 2 to the working gas outlet 5 can be made continuous with respect to the flow passage length on the meridian plane (FIG. 3 (c)),
The internal loss of the working gas between the blades 4a and 4b is significantly reduced.
【０００９】図５におけるＩＩ−ＩＩ断面図（ハブ背部
３１７の外周縁における半径方向の最も低い位置を通り
平均流路方向に垂直な面で切断した断面図）を６図に示
す。覆い部１０８のＩＩ−ＩＩ断面における周方向の脚
長１２１と翼４ａ、４ｂの高さ方向の脚長１２２が、隣
り合う翼４ａ、４ｂ間のピッチ長１２０の１／５乃至１
／２程度となるように形成すると、作動ガス入口２から
作動ガス出口５までの流路面積の連続性を妨げることな
く作動ガス流れ１１の乱れの低減を実現でき好ましいラ
ジアルタービン動翼を形成することができる。なお、肉
盛部覆い部１０８の表面形状は滑らかな曲面で形成する
ことが望ましいが、必ずしも曲面形状のみに限定される
ものではない。作動ガス入口２から作動ガス出口５まで
の流路面積の連続性を妨げず、作動ガスがスムーズに流
れる形状であれば平面に形成しても構わない。また、図
６に１２３で示された線、即ち、図２のＩＩ−ＩＩ線で
切断したときに隣り合う翼４ａ、４ｂ間に形成される覆
い部１０８の表面切断線が、円弧や楕円の一部、放物線
となるように構成してもよい。FIG. 6 is a sectional view taken along line II-II in FIG. 5 (a sectional view taken along a plane which passes through the lowest position in the radial direction on the outer peripheral edge of the hub back portion 317 and is perpendicular to the average flow path direction). The leg length 121 in the circumferential direction and the leg length 122 in the height direction of the blades 4a and 4b in the II-II cross section of the cover portion 108 are ⅕ to 1 of the pitch length 120 between the adjacent blades 4a and 4b.
When it is formed so as to be about 1/2, the disturbance of the working gas flow 11 can be reduced without disturbing the continuity of the flow passage area from the working gas inlet 2 to the working gas outlet 5, and a preferable radial turbine blade is formed. be able to. The surface shape of the built-up portion covering portion 108 is preferably formed by a smooth curved surface, but is not necessarily limited to the curved surface shape. It may be formed on a flat surface as long as it has a shape in which the flow area of the working gas from the working gas inlet 2 to the working gas outlet 5 is not hindered and the working gas smoothly flows. Further, the line indicated by 123 in FIG. 6, that is, the surface cutting line of the covering portion 108 formed between the adjacent blades 4a and 4b when cut along the line II-II in FIG. 2 is an arc or an ellipse. A part may be configured to be a parabola.
【００１０】[0010]
【発明の効果】本願の請求項１に記載された発明による
と、背板ハブ背部の外周側におけるの縁部近傍のハブ面
に沿う部分にて生ずる作動ガスの乱れが少なくなるた
め、作動効率の低下を抑制することができる。また、作
動ガス入口から作動ガス出口までの流路面積が不連続と
なる状態を緩和できるため、翼間における作動ガスの内
部損失を軽減することが可能となる。また、本願の請求
項２に記載された発明によると、作動ガス入口から作動
ガス出口までの流路面積を子午面上の流路長さ対して連
続な状態にすることができるため、翼間における作動ガ
スの内部損失を大幅に軽減することが可能となる。ま
た、本願の請求項３に記載された発明によると、作動ガ
スが肉盛部覆い部の表面に沿ってスムーズに流れるよう
になるため、作動効率を大きく向上させることが可能と
なる。また、作動ガス入口から作動ガス出口までの流路
面積を子午面上の流路長さに対して連続な状態にするこ
とができるため、翼間における作動ガスの内部損失を大
幅に軽減することが可能となる。According to the invention described in claim 1 of the present application, the working gas is less disturbed in the portion along the hub surface in the vicinity of the edge on the outer peripheral side of the back plate hub, so that the working efficiency is reduced. Can be suppressed. Further, since it is possible to mitigate the state in which the flow path area from the working gas inlet to the working gas outlet is discontinuous, it is possible to reduce the internal loss of the working gas between the blades. Further, according to the invention described in claim 2 of the present application, the flow passage area from the working gas inlet to the working gas outlet can be made continuous with respect to the length of the flow passage on the meridian plane, so It is possible to significantly reduce the internal loss of the working gas in. Further, according to the invention described in claim 3 of the present application, since the working gas flows smoothly along the surface of the overlaying portion covering portion, the operating efficiency can be greatly improved. Also, since the flow passage area from the working gas inlet to the working gas outlet can be made continuous with respect to the flow passage length on the meridian surface, the internal loss of working gas between the blades can be greatly reduced. Is possible.
【図１】本発明の第１の実施の形態におけるラジアルタ
ービン動翼の一部を示した斜視図。FIG. 1 is a perspective view showing a part of a radial turbine rotor blade according to a first embodiment of the present invention.
【図２】本発明の第１の実施の形態におけるラジアルタ
ービン動翼の翼間中央部を、軸を含む面で切断したとき
の断面図。FIG. 2 is a cross-sectional view of the radial turbine rotor blade in the first embodiment of the present invention taken along a plane including an axis at a center portion between blades.
【図３】本発明及び従来技術におけるラジアルタービン
動翼の流路断面積の分布を示したグラフ。FIG. 3 is a graph showing the distribution of the flow passage cross-sectional areas of the radial turbine blade in the present invention and the prior art.
【図４】本発明の第２の実施の形態におけるラジアルタ
ービン動翼の一部を示した斜視図。FIG. 4 is a perspective view showing a part of a radial turbine rotor blade according to a second embodiment of the present invention.
【図５】本発明の第２の実施の形態におけるラジアルタ
ービン動翼の翼間中央部を、軸を含む面で切断したとき
の断面図。FIG. 5 is a cross-sectional view of a radial turbine rotor blade in a second embodiment of the present invention, in which a center portion between blades is cut along a plane including a shaft.
【図６】図５におけるＩＩ−ＩＩ断面図6 is a sectional view taken along line II-II in FIG.
【図７】従来のラジアルタービン動翼の一部を示した斜
視図。FIG. 7 is a perspective view showing a part of a conventional radial turbine moving blade.
【図８】従来のラジアルタービン動翼の翼間中央部を、
軸を含む面で切断したときの断面図。FIG. 8 shows a center portion between blades of a conventional radial turbine rotor blade,
Sectional drawing when it cut | disconnects by the surface containing an axis | shaft.
【図９】従来のラジアルタービン動翼の流路断面積の分
布を示したグラフ。FIG. 9 is a graph showing a distribution of flow path cross-sectional areas of a conventional radial turbine blade.
０２、２ 作動ガス入口 ０３、３ 作動ガスの流れ ０４ａ、０４ｂ 翼 ４ａ、４ｂ 翼 ０５、５ 作動ガス出口 ０７、７ ハブ背部背面 ０７ａ０１０ ハブ背部の端面 ０２０、２０ 回転の軸 ０３０、３０ ハブ ０３１、３１ ハブの背部 ３２ 背板 １０８ 覆い肉盛部 02, 2 Working gas inlet 03,3 Working gas flow 04a, 04b wings 4a, 4b wings 05, 5 Working gas outlet 07, 7 Hub back side 07a010 End surface of hub back 020, 20 rotation axis 030, 30 hub 031, 31 Hub back 32 backboard 108 Overlay part
───────────────────────────────────────────────────── フロントページの続き (72)発明者 荻田 浩司 東京都港区芝五丁目33番８号 三菱自動車 工業株式会社内 Ｆターム(参考） 3G002 AA01 AB00 BA01 BB01 3G005 EA04 EA16 FA00 GB79 GB81 ─────────────────────────────────────────────────── ─── Continued front page (72) Inventor Koji Ogita Mitsubishi Motors, 5-3-8 Shiba, Minato-ku, Tokyo Industry Co., Ltd. F term (reference) 3G002 AA01 AB00 BA01 BB01 3G005 EA04 EA16 FA00 GB79 GB81
Claims (3)
中心に向かい凹となるように形成されているラジアルタ
ービン動翼であって、当該外周縁を形成するハブ背部の
外周側の端部が回転軸を含む断面で見たときに鋭角状と
なるように形成されていることを特徴とするラジアルタ
ービン動翼。1. A radial turbine moving blade, wherein the outer peripheral edge of the hub back portion is formed so as to be concave between the blades toward the center of rotation, and the outer peripheral end portion of the hub back portion forming the outer peripheral edge. Is formed so as to have an acute angle when viewed in a cross section including a rotation axis.
形成されたラジアルタービン動翼であって、作動ガス入
口から作動ガス出口までの流路面積が、連続的に増加す
るよう形成されたことを特徴とするラジアルタービン動
翼。2. A radial turbine rotor blade having a flow passage for working gas formed between a hub and a blade adjacent to the hub, wherein the flow passage area from the working gas inlet to the working gas outlet is continuously increased. Radial turbine rotor blades characterized by being made.
流路が形成されたラジアルタービン動翼において、背
板、ハブ及び翼で囲まれた隅に、作動ガス流路に対し滑
らかな面を形成する覆い部が形成されていることを特徴
とするラジアルタービン動翼。3. A radial turbine rotor blade having a flow passage for working gas formed by a hub, a blade adjacent to the back plate and a blade adjacent to the back plate, the corner being surrounded by the back plate, the hub and the blade being smooth with respect to the working gas flow passage. A radial turbine rotor blade having a cover portion that forms a flat surface.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP2001318577A JP2003120202A (en) | 2001-10-16 | 2001-10-16 | Radial turbine rotor blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP2001318577A JP2003120202A (en) | 2001-10-16 | 2001-10-16 | Radial turbine rotor blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JP2003120202A true JP2003120202A (en) | 2003-04-23 |
Family
ID=19136256
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP2001318577A Pending JP2003120202A (en) | 2001-10-16 | 2001-10-16 | Radial turbine rotor blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP2003120202A (en) |
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2009191639A (en) * | 2008-02-12 | 2009-08-27 | Toyota Central R&D Labs Inc | Variable capacity turbine and variable capacity turbo charger |
| CN101915126A (en) * | 2010-06-04 | 2010-12-15 | 清华大学 | Tandem blade type mixed-flow or radial-flow turbine |
| JP2014001712A (en) * | 2012-06-20 | 2014-01-09 | Toyota Central R&D Labs Inc | Radial turbine rotor, and variable geometry turbocharger including the same |
| WO2014008117A1 (en) * | 2012-07-02 | 2014-01-09 | Borgwarner Inc. | Method for turbine wheel balance stock removal |
| JP2019100205A (en) * | 2017-11-29 | 2019-06-24 | 三菱重工業株式会社 | Turbine wheel, turbocharger and manufacturing method of turbine wheel |
-
2001
- 2001-10-16 JP JP2001318577A patent/JP2003120202A/en active Pending
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2009191639A (en) * | 2008-02-12 | 2009-08-27 | Toyota Central R&D Labs Inc | Variable capacity turbine and variable capacity turbo charger |
| CN101915126A (en) * | 2010-06-04 | 2010-12-15 | 清华大学 | Tandem blade type mixed-flow or radial-flow turbine |
| JP2014001712A (en) * | 2012-06-20 | 2014-01-09 | Toyota Central R&D Labs Inc | Radial turbine rotor, and variable geometry turbocharger including the same |
| WO2014008117A1 (en) * | 2012-07-02 | 2014-01-09 | Borgwarner Inc. | Method for turbine wheel balance stock removal |
| JP2019100205A (en) * | 2017-11-29 | 2019-06-24 | 三菱重工業株式会社 | Turbine wheel, turbocharger and manufacturing method of turbine wheel |
| JP7002306B2 (en) | 2017-11-29 | 2022-01-20 | 三菱重工業株式会社 | How to manufacture turbine wheels, turbochargers and turbine wheels |
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