JP2014125924A - Turbine stationary blade and axial flow turbine - Google Patents
Turbine stationary blade and axial flow turbine Download PDFInfo
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Abstract
Description
本発明は、タービン静翼、特に蒸気タービンに用いるタービン静翼に関する。 The present invention relates to a turbine vane, and more particularly to a turbine vane used for a steam turbine.
蒸気タービンのタービン翼に発生する二次流れ損失を低減する方法として、タービン翼が流体から最もエネルギーを受け取る位置、すなわち翼の負荷位置に着目した方法がある。この方法に関する従来技術としては、例えば特開2006−207556号公報に記載の技術が知られている。 As a method of reducing the secondary flow loss generated in the turbine blades of the steam turbine, there is a method that focuses on the position where the turbine blades receive the most energy from the fluid, that is, the load position of the blades. As a conventional technique related to this method, for example, a technique described in JP-A-2006-207556 is known.
上記特開2006−207556号公報には、翼高さ方向に異なる翼負荷分布を使用し、特に翼中央部を前半負荷型にし、翼根元部と翼先端部を後半負荷型としたタービン翼が記載されている。 In the above Japanese Patent Laid-Open No. 2006-207556, there is a turbine blade that uses different blade load distributions in the blade height direction, in particular, a blade center portion is a front half load type, and a blade root portion and a blade tip portion are latter half load types. Have been described.
翼負荷は、翼の背側圧力と腹側圧力の差で表現し、この差が大きいほど翼負荷が大きいことになる。最大負荷位置が、タービン翼の前縁に近いタービン翼は前半負荷型、タービン翼の後縁位置に近い翼は後半負荷型となる。後半負荷型のタービン翼は、翼前半部で負荷を小さくする、すなわち翼背側と翼腹側の圧力差が小さくなり、二次流れが発達しにくくなるため、前半負荷型のタービン翼よりも二次損失は小さくなると言われている。 The blade load is expressed by a difference between the back pressure and the ventral pressure of the blade, and the blade load increases as the difference increases. A turbine blade whose maximum load position is close to the leading edge of the turbine blade is a first half load type, and a blade near the trailing edge position of the turbine blade is a second half load type. In the latter half load type turbine blade, the load is reduced in the front half of the blade, that is, the pressure difference between the blade back side and the blade belly side is reduced, and the secondary flow is difficult to develop. Secondary loss is said to be small.
しかしながら、後半負荷型のタービン翼は、最大負荷位置が翼後縁に近くなるため、翼背側の最小圧力点は、翼後縁部に近い位置となる。このため、最小圧力点位置から後縁部にかけて急激な圧力回復が生じる。これは、隣り合うタービン翼の間を通過する流れが翼後縁部で減速することを意味し、同時に翼間に生じる二次流れの影響を受けやすくなることを意味する。減速した翼間流れは、隣接するタービン翼背側にぶつかり、二次損失を増大させてしまう。しかしながら、この課題について従来技術は考慮していない。 However, since the maximum load position of the latter half load type turbine blade is close to the blade trailing edge, the minimum pressure point on the blade back side is close to the blade trailing edge. For this reason, rapid pressure recovery occurs from the position of the minimum pressure point to the rear edge. This means that the flow passing between adjacent turbine blades is decelerated at the blade trailing edge, and at the same time, it is susceptible to the secondary flow generated between the blades. The reduced inter-blade flow collides with the adjacent turbine blade back side, increasing the secondary loss. However, the prior art does not consider this problem.
そこで本発明の課題は、後半負荷型のタービン翼において、二次損失を低減し、タービン段落の効率を向上させることにある。 Accordingly, an object of the present invention is to reduce the secondary loss and improve the efficiency of the turbine stage in the latter half load type turbine blade.
上記課題を解決するための、本発明は、軸流タービンのタービン静翼において、タービン静翼の翼背側面の圧力によって規定される圧力分布曲線が、翼前縁部から回転軸方向後半側まで圧力降下し、圧力降下した後圧力を一定に保持する区間を経て翼後縁部まで圧力上昇するように翼型を形成したことを特徴とする。 In order to solve the above-described problems, the present invention provides a turbine stationary blade of an axial flow turbine in which a pressure distribution curve defined by the pressure on the blade back surface of the turbine stationary blade is from the blade leading edge to the second half in the rotational axis direction. The airfoil is formed such that the pressure drops and the pressure rises to the trailing edge of the blade through a section where the pressure is kept constant after the pressure is dropped.
本発明によれば、後半負荷型のタービン翼において、二次損失を低減し、タービン段落の効率を向上させることができる。 ADVANTAGE OF THE INVENTION According to this invention, in a latter half load type turbine blade, a secondary loss can be reduced and the efficiency of a turbine stage can be improved.
以下、本発明を実施するための形態について、適宜図を参照して詳細に説明する。なお、各図面を通し、同等の構成要素には同一の符号を付してある。 Hereinafter, embodiments for carrying out the present invention will be described in detail with reference to the drawings as appropriate. In addition, the same code | symbol is attached | subjected to the equivalent component through each drawing.
本発明を蒸気タービンのタービン静翼に適用した例について説明する。図１は、本発明を適用する蒸気タービンのタービン段落の横断面を模式的に示したものである。図中、１は静翼、２は静翼１の内周側端部５を固定する環状のダイヤフラム内輪、３は静翼１の外周側端部４を固定する環状のダイヤフラム外輪、６は動翼、７は動翼６をロータに固定するロータディスク、８は動翼６の外周側先端を拘束するシュラウドカバー、９は作動流体である蒸気の流れ方向を示す。ダイヤフラム外輪３は図示しないケーシングに固定されている。図１において、Ｘ軸はタービンの回転軸方向、Ｚ軸はタービンの半径方向を示す。 An example in which the present invention is applied to a turbine stationary blade of a steam turbine will be described. FIG. 1 schematically shows a cross section of a turbine stage of a steam turbine to which the present invention is applied. In the figure, 1 is a stationary blade, 2 is an annular diaphragm inner ring that fixes the inner peripheral side end 5 of the stationary blade 1, 3 is an annular diaphragm outer ring that fixes the outer peripheral side end 4 of the stationary blade 1, and 6 is a moving ring. Reference numeral 7 denotes a rotor disk for fixing the moving blade 6 to the rotor, 8 denotes a shroud cover for restraining the outer peripheral end of the moving blade 6, and 9 denotes a flow direction of steam as a working fluid. The diaphragm outer ring 3 is fixed to a casing (not shown). In FIG. 1, the X axis indicates the rotational axis direction of the turbine, and the Z axis indicates the radial direction of the turbine.
静翼１および動翼６はそれぞれ円周方向に複数枚配置されており、隣り合う翼間に作動流体が通過する流路を形成している。動翼６は、静翼１の蒸気の流れ方向下流側に設置されており、静翼１によって加速された蒸気が動翼６に衝突することでロータが回転し、ロータの端部に据え付けられた発電機（図示せず）を回転させる。 A plurality of the stationary blades 1 and the moving blades 6 are arranged in the circumferential direction, and form a flow path through which the working fluid passes between adjacent blades. The rotor blade 6 is installed on the downstream side in the steam flow direction of the stationary blade 1, and when the steam accelerated by the stator blade 1 collides with the rotor blade 6, the rotor rotates and is installed at the end of the rotor. Rotate a generator (not shown).
図２は、円周方向に複数枚配置された静翼１の翼列の一部を模式的に示した斜視図である。図２では、説明の便宜上、円周方向に配置された静翼の一部を水平に展開して示している。静翼１ａ、ｂは、外周側端部４を環状のダイヤフラム外輪３に固定されており、内周側端部５は環状のダイヤフラム内輪２に固定されている。周方向に隣り合う静翼１ａ、ｂの壁面とダイヤフラム外輪３およびダイヤフラム内輪２の壁面に囲まれて作動流体の流路が形成されている。静翼１の前縁部１０側から流入した蒸気は、静翼１ａの翼背側の負圧面１３と静翼１ｂの腹側の圧力面１２との間を流れる。静翼間の流路は、流路入口から出口に向かって絞り流路になっており、蒸気は、静翼間の流路を通過する間に流れ方向を変えるとともに加速され、後縁部１１から下流の動翼に向かって排出される。 FIG. 2 is a perspective view schematically showing a part of the blade row of the stationary blade 1 arranged in the circumferential direction. In FIG. 2, for convenience of explanation, a part of the stationary blades arranged in the circumferential direction is shown in a horizontally developed state. The stationary blades 1 a and b have an outer peripheral side end 4 fixed to the annular diaphragm outer ring 3 and an inner peripheral side end 5 fixed to the annular diaphragm inner ring 2. A flow path for the working fluid is formed by being surrounded by the wall surfaces of the stationary blades 1 a and 1 b adjacent to each other in the circumferential direction and the wall surfaces of the diaphragm outer ring 3 and the diaphragm inner ring 2. The steam that has flowed in from the front edge 10 side of the stationary blade 1 flows between the negative pressure surface 13 on the back side of the stationary blade 1a and the pressure surface 12 on the ventral side of the stationary blade 1b. The flow path between the stationary blades is a throttle flow path from the flow path inlet to the outlet, and the steam changes its direction of flow and is accelerated while passing through the flow path between the stationary blades, and the trailing edge 11 Is discharged toward the moving blade downstream.
図３に一般的な翼間流路における流れの様子を模式的に示す。図３ではダイヤフラム外輪３およびダイヤフラム内輪２は図示を省略している。一般的に、蒸気が静翼間の流路を通過する時、翼間には圧力勾配が生じる。そのため、圧力面１２から負圧面１３に向かう二次流れ１５が生じる。一方、翼間流路の入口部では、蒸気主流が翼前縁部１０に衝突することで、圧力面１２と負圧面１３に渦が発生する。この渦は、翼間流路内で発達して流路渦１６を形成する。流路渦１６は、主に翼の内周側端部５と外周側端部４付近にそれぞれ存在する。これらの二次流れ１５、流路渦１６は、タービン翼が本来行うべき仕事の効率低下を招き、二次損失の大きな要因となる。 FIG. 3 schematically shows the flow in a general flow path between blades. In FIG. 3, the diaphragm outer ring 3 and the diaphragm inner ring 2 are not shown. Generally, when steam passes through a flow path between stationary blades, a pressure gradient is generated between the blades. Therefore, a secondary flow 15 is generated from the pressure surface 12 toward the suction surface 13. On the other hand, at the inlet portion of the inter-blade flow path, the main steam flow collides with the blade leading edge portion 10 to generate vortices on the pressure surface 12 and the negative pressure surface 13. This vortex develops in the inter-blade channel and forms a channel vortex 16. The flow path vortex 16 exists mainly in the vicinity of the inner peripheral end 5 and the outer peripheral end 4 of the blade. These secondary flow 15 and flow path vortex 16 cause a reduction in work efficiency that the turbine blades should originally perform, and cause a large secondary loss.
図４は図２に示した静翼の端部付近の断面を模式的に表した図である。図４において、Ｘ軸方向はタービンの回転軸方向を表し、Ｙ軸方向はタービンの周方向を表す。以後の説明のため、静翼１の前縁部１０から後縁部１１までのＸ方向の距離を軸コード長Ｃａｘ、前縁部１０から任意の位置までのＸ方向の距離ｘを軸コード長Ｃａｘで割った値を無次元軸コード長ｘ／Ｃａｘと定義する。 FIG. 4 is a diagram schematically showing a cross section near the end of the stationary blade shown in FIG. In FIG. 4, the X-axis direction represents the rotational axis direction of the turbine, and the Y-axis direction represents the circumferential direction of the turbine. For the following description, the distance in the X direction from the leading edge 10 to the trailing edge 11 of the stationary blade 1 is the axial code length Cax, and the distance x in the X direction from the leading edge 10 to an arbitrary position is the axial code length. The value divided by Cax is defined as dimensionless axis code length x / Cax.
静翼１は円周方向に一定の間隔を置いて配置されており、周方向に隣り合う静翼の後縁部１１同士の間隔をピッチ長ｔと呼ぶ。ピッチ長ｔは、主に周方向に配置する翼枚数によって決定される。また、後縁部１１から、翼腹側に設置された隣の静翼までの最短距離を結んだ線をスロートといい、その長さをスロート長ｓという。スロートは、翼間の流路のうち最小の流路幅となる。なお、図４において、スロート線と翼腹側に設置された隣の静翼のプロファイル線との交点Ｍをスロート点という。タービン翼は、タービン翼出口角を表す幾何学的な角度としてsin-1（ｓ／ｔ）と定義され、ピッチ長ｔとスロート長ｓの比（ｓ／ｔ）を用いた値を使用する。ｓ／ｔを調整することで、静翼を通過する蒸気を動翼へ適正に流入させることができる。一般的にｓ／ｔを小さくすることは、隣り合う静翼１の間隔が小さくなるため、スロートを通過した蒸気は二次流れの影響を受けやすくなる。また、スロートを通過した蒸気、あるいは翼間流路内で発達した流路渦１６は、翼腹側に設置された隣の翼の負圧面１３に衝突しやすくなる。負圧面に衝突することは、二次損失が大きくなることを意味する。本発明は、ｓ／ｔが小さい場合に二次損失低減の効果があり、最適なｓ／ｔは０．２０〜０．２５である。 The stationary blades 1 are arranged at a constant interval in the circumferential direction, and the interval between the trailing edge portions 11 adjacent to each other in the circumferential direction is called a pitch length t. The pitch length t is determined mainly by the number of blades arranged in the circumferential direction. A line connecting the shortest distance from the trailing edge 11 to the adjacent stationary blade installed on the blade side is called a throat, and its length is called a throat length s. The throat has the smallest channel width among the channels between the blades. In FIG. 4, the intersection M between the throat line and the profile line of the adjacent stationary blade installed on the blade side is called the throat point. The turbine blade is defined as sin −1 (s / t) as a geometric angle representing the turbine blade exit angle, and a value using a ratio (s / t) of the pitch length t and the throat length s is used. By adjusting s / t, the steam passing through the stationary blade can be appropriately introduced into the moving blade. Generally, when s / t is reduced, the interval between adjacent stationary blades 1 is reduced, so that the steam that has passed through the throat is easily affected by the secondary flow. Further, the steam passing through the throat or the flow path vortex 16 developed in the flow path between the blades easily collides with the suction surface 13 of the adjacent blade installed on the blade side. Colliding with the suction surface means that the secondary loss increases. The present invention has an effect of reducing the secondary loss when s / t is small, and the optimum s / t is 0.20 to 0.25.
次に図５を用いて本実施のタービン静翼の主な特徴である翼面の圧力分布について説明する。図５に本実施例のタービン静翼の翼面圧力分布と、従来の後半負荷型翼の圧力分布を示す。図中、上側が翼の腹側面に作用する圧力の分布、下側が翼の背側面に作用する圧力の分布である。なお、翼腹側の圧力分布は、本実施例も従来と変わらないので、本実施例の圧力分布のみ示し、従来の圧力分布は省略している。また、本実施例で述べる圧力分布形状は、インシデンス角を持つ場合（非設計点等）のような翼前縁や翼後縁近傍の圧力が急激に変化する領域は除外している。 Next, the blade surface pressure distribution, which is the main feature of the turbine stationary blade of the present embodiment, will be described with reference to FIG. FIG. 5 shows the blade surface pressure distribution of the turbine stationary blade of this embodiment and the pressure distribution of the conventional latter half load type blade. In the figure, the upper side is the distribution of pressure acting on the ventral side of the wing, and the lower side is the distribution of pressure acting on the back side of the wing. In addition, since the pressure distribution on the blade ventral side is the same as that of the conventional example, only the pressure distribution of the present example is shown, and the conventional pressure distribution is omitted. In addition, the pressure distribution shape described in the present embodiment excludes a region where the pressure in the vicinity of the blade leading edge and the blade trailing edge changes suddenly as in the case of having an incidence angle (such as a non-design point).
本実施例の圧力分布の特徴は、翼背側の圧力分布にある。従来の後半負荷型翼では、翼背側の圧力分布は、最小圧力値がｘ／Ｃａｘ＝０．８の近傍にあり、翼前縁部から最小圧力値に向かって圧力が単調に降下する。その後、圧力分布は、最小圧力値を境に上昇に転じ、後縁部に向かって圧力は上昇する。一方、本実施例の圧力分布は、前縁部から後縁部に向かって単調に降下した後、後半で後縁部に向かって上昇に転じる前に、圧力値がほぼ一定のまま保持される区間を有する。この区間では、圧力値は降下傾向が終わり、しかし上昇傾向にも転じず、圧力値が一定となる。なおここでの圧力一定とは、蒸気を加速もしくは減速させず、なるべく上流側から加速された蒸気速度を維持できる範囲内で圧力値が上下に微増減する場合も含まれる。また、圧力を一定に保持する領域は、無次元軸コード長ｘ／Ｃａｘで０．８以上１．０未満にあればよい。 The characteristic of the pressure distribution of this embodiment is the pressure distribution on the blade back side. In the conventional latter half load type blade, the pressure distribution on the blade back side is in the vicinity of the minimum pressure value x / Cax = 0.8, and the pressure monotonously decreases from the blade leading edge toward the minimum pressure value. Thereafter, the pressure distribution starts to increase at the minimum pressure value, and the pressure increases toward the trailing edge. On the other hand, in the pressure distribution of the present example, the pressure value is kept substantially constant after the monotonous drop from the leading edge toward the trailing edge and before rising toward the trailing edge in the second half. It has a section. In this section, the pressure value ends the downward trend, but does not turn upward, and the pressure value is constant. Here, the constant pressure includes a case where the pressure value slightly increases or decreases within a range in which the steam velocity accelerated from the upstream side can be maintained as much as possible without accelerating or decelerating the steam. Moreover, the area | region which hold | maintains a pressure uniformly should just be 0.8 or more and less than 1.0 in dimensionless axis code length x / Cax.
従来の後半負荷型翼は、前縁部から後半に向かって圧力値が降下する圧力降下領域と、圧力値が後縁部に向かって上昇する圧力上昇領域の二領域からなるのに対し、本実施例の圧力分布は、前縁部から後半に向かって圧力値が降下する圧力降下領域と、圧力値が一定に保持される領域と、圧力値が後縁部に向かって上昇する圧力上昇領域に区分けできる。 The conventional latter half load type wing consists of two areas: a pressure drop area where the pressure value drops from the leading edge toward the latter half and a pressure rise area where the pressure value rises toward the trailing edge. The pressure distribution of the embodiment includes a pressure drop region where the pressure value drops from the front edge toward the second half, a region where the pressure value is held constant, and a pressure rise region where the pressure value rises toward the rear edge. Can be divided into
また、本実施例の圧力分布は、圧力分布の圧力降下領域の範囲が従来よりも狭く、従来の圧力分布よりも最小圧力値が大きくなる。最小圧力値は圧力値が一定となる領域の中に位置し、翼後縁部に向かって圧力が上昇傾向に転じる位置もより後縁部側になる。 In addition, the pressure distribution of this embodiment has a narrower pressure drop range than the conventional pressure distribution, and a minimum pressure value larger than that of the conventional pressure distribution. The minimum pressure value is located in a region where the pressure value is constant, and the position where the pressure tends to increase toward the blade trailing edge is also on the trailing edge side.
図６に軸コード長方向の背側の圧力変化率を表すグラフを示した。縦軸を背側の圧力変化率、横軸を無次元軸コード長（ｘ／Ｃａｘ）とする。圧力変化率は、無次元圧力ｐの無次元軸コード長（ｘ／Ｃａｘ）の圧力変化率であり、ｄｐ／（ｘ／Ｃａｘ）で表す。ｄｐ／（ｘ／Ｃａｘ）＞０の場合、翼間の流れは加速される。一方、ｄｐ／（ｘ／Ｃａｘ）＜０の場合は、翼間の流れは減速される。ここで、無次元圧力ｐは静翼翼列入口の全圧を1として無次元化した値である。 FIG. 6 is a graph showing the pressure change rate on the back side in the axial cord length direction. The vertical axis represents the pressure change rate on the back side, and the horizontal axis represents the dimensionless axis code length (x / Cax). The pressure change rate is the pressure change rate of the dimensionless axis code length (x / Cax) of the dimensionless pressure p, and is represented by dp / (x / Cax). When dp / (x / Cax)> 0, the flow between the blades is accelerated. On the other hand, when dp / (x / Cax) <0, the flow between the blades is decelerated. Here, the dimensionless pressure p is a value obtained by dimensionlessly setting the total pressure at the stationary blade cascade inlet to 1.
本実施例の圧力分布の場合、背側の圧力変化率は、漸次増大していき、最大値は従来例よりも大きくなる。その後圧力変化率は急速に減少し、０．８〜０．９付近では加速または減速が極めて小さい領域となる。最大加速後、圧力変化率が初めて０．０２以下となる位置を第一圧力点、背側圧力変化が−０．０２を下回った位置を第二圧力点と定義し、第一圧力点と第二圧力点を図５に示した本実施例の圧力分布に当てはめたグラフを図７に示す。 In the case of the pressure distribution of this embodiment, the pressure change rate on the back side gradually increases, and the maximum value becomes larger than that of the conventional example. Thereafter, the rate of pressure change decreases rapidly, and in the vicinity of 0.8 to 0.9, the acceleration or deceleration becomes a very small region. After the maximum acceleration, the position where the pressure change rate is 0.02 or less for the first time is defined as the first pressure point, and the position where the back pressure change is below -0.02 is defined as the second pressure point. FIG. 7 shows a graph in which the two pressure points are applied to the pressure distribution of this embodiment shown in FIG.
図７より、第一圧力点から第二圧力点までの領域は、圧力値がほぼ一定となる区間と重なることが分かる。第一圧力点から第二圧力点までの圧力の変動幅は、翼入口部の全圧で無次元化した無次元圧力値で、第一圧力点から第二圧力点までの平均圧力値から２％以内が望ましい。また、前縁から第一圧力点までは降下するが途中で変曲点は持たないことが望ましい。この理由は、翼間でスムーズに加速させるためである。 From FIG. 7, it can be seen that the region from the first pressure point to the second pressure point overlaps with a section where the pressure value is substantially constant. The pressure fluctuation range from the first pressure point to the second pressure point is a dimensionless pressure value made dimensionless by the total pressure at the blade inlet, and is 2 from the average pressure value from the first pressure point to the second pressure point. % Is desirable. Moreover, it is desirable that it descends from the leading edge to the first pressure point but does not have an inflection point on the way. The reason for this is to accelerate smoothly between the blades.
次に、図4で規定したスロート点Mと翼面圧力の位置関係を示すと、スロート点Mは第一圧力点よりも上流側に位置する。つまり、スロート点Mより下流側で第一圧力点から第二圧力点までの圧力値がほぼ一定となる区間が存在する事となる。第二圧力点がスロート点の上流側である場合、または、スロート点が第一圧力点と第二圧力点に存在する場合、翼間で十分に加速されず、隣接するタービン翼背側にぶつかり、二次損失を増大させてしまう。このため一定を保持する区間を定義する第一圧力点はスロート点よりも下流側にあることが望ましい。 Next, when the positional relationship between the throat point M and the blade surface pressure defined in FIG. 4 is shown, the throat point M is located upstream of the first pressure point. That is, there is a section where the pressure value from the first pressure point to the second pressure point is substantially constant downstream from the throat point M. When the second pressure point is upstream of the throat point, or when the throat point exists at the first pressure point and the second pressure point, the blades are not accelerated sufficiently and collide with the adjacent turbine blade back side. Increase the secondary loss. For this reason, it is desirable that the first pressure point that defines the section in which the constant is maintained be downstream of the throat point.
本実施例の作用効果を図８および９に基づいて説明する。図８は、静翼の翼間流路の流速を示すグラフであり、グラフ中には従来例と本実施例をそれぞれ示す。本実施例と従来例とで翼間の流速にもっとも顕著な違いが現れる場所は、翼後縁部近傍（ｘ／Ｃａｘ＝０．８〜１．０）である。従来例では、翼間の流速の最大値はｘ／Ｃａｘ＝０．９付近をとり、この位置より後縁に向かって減速する。一方、本実施例では、従来の後半負荷型翼よりも、最大流速の位置を後縁部に近い位置に移動させ、翼間の流速を速くすることができることが図８より分かる。 The effect of a present Example is demonstrated based on FIG. 8 and 9. FIG. FIG. 8 is a graph showing the flow velocity of the flow path between the vanes of the stationary blade. The graph shows the conventional example and the present embodiment, respectively. The place where the most remarkable difference in the flow velocity between the blades in this embodiment and the conventional example appears in the vicinity of the blade trailing edge (x / Cax = 0.8 to 1.0). In the conventional example, the maximum value of the flow velocity between the blades is in the vicinity of x / Cax = 0.9, and decelerates from this position toward the trailing edge. On the other hand, in this embodiment, it can be seen from FIG. 8 that the position of the maximum flow velocity can be moved closer to the trailing edge than the conventional latter half load type blades, and the flow velocity between the blades can be increased.
最大流速の位置を後縁部に近い位置に移動させ、翼間の流速を速くすることで、翼間に生じる二次流れの影響を受けにくくなり、流れが隣に配置された翼に衝突することを抑制することができる。また、翼の負圧面に流れが衝突する場所を従来翼よりも下流側へ移動することができる。その結果、二次損失の低減につながり、タービン段落の効率を向上することができる。 By moving the position of the maximum flow velocity to a position close to the trailing edge and increasing the flow velocity between the blades, it becomes less susceptible to the secondary flow generated between the blades, and the flow collides with the adjacent blade. This can be suppressed. Further, the location where the flow collides with the suction surface of the blade can be moved downstream from the conventional blade. As a result, the secondary loss is reduced, and the efficiency of the turbine stage can be improved.
図９は従来例と本実施例の翼高さ方向の損失分布を模式的に表した図である。本実施例の損失分布を実線で示し、従来例の損失分布を点線で示す。図９の損失分布は、軸コード長Ｃａｘに対する翼高さの割合を表すアスペクト比が比較的大きい翼の例である。本例の場合、アスペクト比は約3.8である。一般的に、従来例のように翼高さ位置に応じて損失の傾向が異なり、翼の中央部付近で損失がもっとも小さくなり、外周側および内周側端部付近で損失が最大となる。これは、翼の中央部付近では翼型損失が損失の主な要因となるのに対し、翼の外周側および内周側端部付近では翼型損失の他に、前述した二次損失による損失が増加するためである。一方、本実施例の圧力分布を有する翼型を翼外周端部および内周端部に適用した場合、損失分布において翼外周端部および内周端部の損失が低減されていることが分かる。これは、本実施例による二次流れ損失の低減効果による。 FIG. 9 is a diagram schematically showing the loss distribution in the blade height direction of the conventional example and this example. The loss distribution of this example is indicated by a solid line, and the loss distribution of the conventional example is indicated by a dotted line. The loss distribution in FIG. 9 is an example of a blade having a relatively large aspect ratio that represents the ratio of the blade height to the axial cord length Cax. In this example, the aspect ratio is about 3.8. In general, the tendency of loss differs depending on the blade height position as in the conventional example, the loss is the smallest near the center of the blade, and the loss is maximized near the outer peripheral side and the inner peripheral side end. This is because airfoil loss is the main cause of loss near the center of the wing, whereas in addition to airfoil loss near the outer and inner edges of the wing, the loss due to the secondary loss described above. This is because of the increase. On the other hand, when the airfoil having the pressure distribution of the present embodiment is applied to the blade outer peripheral end and the inner peripheral end, it can be seen that the loss at the blade outer peripheral end and the inner peripheral end is reduced in the loss distribution. This is due to the effect of reducing the secondary flow loss according to this embodiment.
以上説明した本実施例の圧力分布曲線から逆解法により翼型を求めることができる。本実施例の圧力分布から求められた翼型は、二次損失による損失が支配的な静翼の翼外周端部および内周端部に適用することが望ましい。特に二次損失が支配的な翼高さ領域から翼型による損失が支配的な領域に切り替わる翼高さ位置に適用することが望ましい。例えば、本実施例の圧力分布を有する翼型は、この例で示したアスペクト比の場合、翼高さのうち翼外周端から３０％、翼内周端から３０％の領域に適用するのが望ましい。 The airfoil can be obtained by the inverse solution from the pressure distribution curve of the present embodiment described above. It is desirable that the airfoil obtained from the pressure distribution of the present embodiment is applied to the blade outer peripheral end and inner peripheral end of the stationary blade in which the loss due to the secondary loss is dominant. In particular, it is desirable to apply to the blade height position where the blade loss is switched from the blade height region where the secondary loss is dominant to the region where the airfoil loss is dominant. For example, in the case of the aspect ratio shown in this example, the airfoil having the pressure distribution of the present embodiment is applied to the region of the blade height that is 30% from the blade outer periphery and 30% from the blade inner periphery. desirable.
また、図１０には本特許の別の適用例として、タービン翼の圧力面側が凸となるよう周方向に湾曲したタービン静翼に適用した例を示す。タービン翼の圧力面側が凸とすることで、翼先端部および翼根元部に向かい翼力が生じる。この翼力により、タービン翼間流れは翼先端部および翼根元部に押し付けられ、前述した二次損失が翼中央部へ発達することを抑える。このように、二次損失低減のために三次元的な変形を実施したタービン静翼についても、本特許を適用することで、二次損失をさらに低減することが可能である。また、周方向、軸方向に傾斜または湾曲させたタービン静翼についても、同様の効果を得ることが出来る。 FIG. 10 shows an example of application of this patent to a turbine stationary blade curved in the circumferential direction so that the pressure surface side of the turbine blade is convex. Since the pressure surface side of the turbine blade is convex, blade force is generated toward the blade tip and blade root. By this blade force, the flow between the turbine blades is pressed against the blade tip portion and the blade root portion, and the above-described secondary loss is prevented from developing to the blade center portion. As described above, it is possible to further reduce the secondary loss by applying this patent to the turbine stationary blade that has been three-dimensionally modified to reduce the secondary loss. Moreover, the same effect can be acquired also about the turbine stationary blade inclined or curved to the circumferential direction and the axial direction.
１ 静翼
２ ダイヤフラム内輪
３ ダイヤフラム外輪
６ 動翼
７ ロータディスク
８ シュラウドカバー
１２ 圧力面
１３ 負圧面
DESCRIPTION OF SYMBOLS 1 Stator blade 2 Diaphragm inner ring 3 Diaphragm outer ring 6 Rotor blade 7 Rotor disk 8 Shroud cover 12 Pressure surface 13 Negative pressure surface
Claims (10)
該タービン静翼の翼背側面の圧力によって規定される圧力分布曲線が、翼前縁部から回転軸方向後半側まで圧力降下し、圧力降下した後圧力を一定に保持する区間を経て翼後縁部まで圧力上昇するように翼型を形成したことを特徴とするタービン静翼。 A turbine vane of an axial turbine,
The pressure distribution curve defined by the pressure on the blade rear side of the turbine vane drops from the blade leading edge to the latter half of the rotation axis direction, and after the pressure drop, the blade trailing edge passes through a section that keeps the pressure constant. A turbine vane characterized in that an airfoil is formed so as to increase the pressure to the part.
前記圧力分布曲線を有するタービン翼型を翼内周側端部および翼外周側端部に有することを特徴とするタービン静翼。 The turbine stationary blade according to claim 1,
A turbine vane having a turbine blade shape having the pressure distribution curve at a blade inner peripheral end and a blade outer peripheral end.
前記圧力分布曲線において前記圧力を一定に保持する区間は、周方向に隣り合う静翼との距離が最小となるスロート点よりも後縁側に位置することを特徴とするタービン静翼。 The turbine stator blade according to claim 1 or 2,
The section of the pressure distribution curve in which the pressure is kept constant is located on the trailing edge side of the throat point where the distance from the circumferentially adjacent stationary blade is minimum.
前記タービン静翼の前縁部から後縁部までのタービン回転軸方向の距離を軸コード長、前縁部から任意の位置までのタービン回転軸方向の距離を前記軸コード長で割った値を無次元軸コードと定義したとき、
前記圧力を一定に保持する区間は、無次元コード長が0.8以上1.0未満の間に位置することを特徴とするタービン静翼。 The turbine stationary blade according to any one of claims 1 to 3,
A value obtained by dividing the distance in the turbine rotating shaft direction from the front edge portion to the rear edge portion of the turbine stationary blade by the axis code length, and dividing the distance in the turbine rotating shaft direction from the front edge portion to an arbitrary position by the shaft code length. When defined as dimensionless axis code,
The section for maintaining the pressure constant is located between a dimensionless code length of 0.8 or more and less than 1.0.
タービン静翼の後縁部から、前記タービン静翼の翼腹側に隣り合う静翼までの最短距離ｓと前記タービン静翼および前記隣り合う静翼の後縁部同士の間隔長ｔとの比から定まるｓ／ｔが翼高さ方向全域で０．２０〜０．２５の範囲にあることを特徴とするタービン静翼。 The turbine stationary blade according to any one of claims 1 to 4, wherein
The ratio of the shortest distance s from the trailing edge of the turbine stationary blade to the stationary blade adjacent to the blade side of the turbine stationary blade and the distance t between the trailing edges of the turbine stationary blade and the adjacent stationary blade S / t determined from the above is in the range of 0.20 to 0.25 throughout the blade height direction.
前記静翼の下流側に配置され、ロータディスクに固定される動翼とからなるタービン段落を備えた軸流タービンであって、
翼背側面の圧力によって規定される圧力分布曲線が、翼前縁部から回転軸方向後半側まで圧力降下する領域と、圧力降下した後圧力を一定に保持する領域と、翼後縁部まで圧力上昇する領域の３つの領域からなるように翼型が形成された前記静翼を備えることを特徴とする軸流タービン。 A stationary blade, an annular diaphragm outer ring that fixes an outer peripheral side end of the stationary blade, a diaphragm inner ring that fixes an inner peripheral side end of the stationary blade,
An axial flow turbine having a turbine stage that is disposed downstream of the stationary blade and includes a moving blade fixed to a rotor disk,
The pressure distribution curve defined by the pressure on the back side of the blade shows the pressure drop from the blade leading edge to the second half in the direction of the rotation axis, the region that keeps the pressure constant after the pressure drop, and the pressure to the blade trailing edge. An axial-flow turbine comprising the stationary blade in which an airfoil is formed so as to include three regions of an ascending region.
前記圧力分布曲線を有する翼型を前記静翼の翼内周側端部および翼外周側端部に有することを特徴とする軸流タービン。 The axial turbine according to claim 6,
An axial turbine having an airfoil having the pressure distribution curve at a blade inner peripheral end and a blade outer peripheral end of the stationary blade.
前記静翼の前記圧力分布曲線において前記圧力を一定に保持する領域は、周方向に隣り合う静翼との距離が最小となるスロート点よりも後縁側に位置することを特徴とする軸流タービン。 An axial turbine according to claim 6 or 7,
In the pressure distribution curve of the stationary blade, the region where the pressure is kept constant is located on the trailing edge side from the throat point where the distance from the circumferentially adjacent stationary blade is minimum. .
前記静翼の前縁部から後縁部までのタービン回転軸方向の距離を軸コード長、前縁部から任意の位置までのタービン回転軸方向の距離を前記軸コード長で割った値を無次元軸コードと定義したとき、
圧力降下した後圧力を一定に保持する領域は、無次元コード長が0.8以上1.0未満の間に位置することを特徴とする軸流タービン。 An axial turbine according to any one of claims 6 to 8,
A value obtained by dividing the distance in the turbine rotating shaft direction from the leading edge portion to the trailing edge portion of the stationary blade by the axial cord length, and dividing the distance in the turbine rotating shaft direction from the leading edge portion to an arbitrary position by the shaft cord length. When defined as a dimension axis code,
The axial flow turbine is characterized in that the region where the pressure is kept constant after the pressure drop is located between a dimensionless code length of 0.8 or more and less than 1.0.
前記静翼の後縁部から、前記タービン静翼の翼腹側に隣り合う静翼までの最短距離と前記静翼と前記隣り合う静翼の後縁部同士の間隔長の比から定まるｓ／ｔが翼高さ方向全域で０．２０〜０．２５の範囲にあることを特徴とする軸流タービン。 An axial turbine according to any one of claims 6 to 9,
S / determined by the ratio of the shortest distance from the trailing edge of the stationary blade to the stationary blade adjacent to the blade side of the turbine stationary blade and the interval length between the stationary blade and the trailing edge of the adjacent stationary blade. t is in the range of 0.20 to 0.25 throughout the blade height direction.
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