EP0661413B1 - Axial blade cascade with blades of arrowed leading edge - Google Patents

Axial blade cascade with blades of arrowed leading edge Download PDF

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Publication number
EP0661413B1
EP0661413B1 EP19940119705 EP94119705A EP0661413B1 EP 0661413 B1 EP0661413 B1 EP 0661413B1 EP 19940119705 EP19940119705 EP 19940119705 EP 94119705 A EP94119705 A EP 94119705A EP 0661413 B1 EP0661413 B1 EP 0661413B1
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EP
European Patent Office
Prior art keywords
blade
cascade
blades
gravity
rotor
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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.)
Expired - Lifetime
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EP19940119705
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German (de)
French (fr)
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EP0661413A1 (en
Inventor
Reinhard Dr. Niehus
Norbert Hübner
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MTU Aero Engines GmbH
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MTU Motoren und Turbinen Union Muenchen GmbH
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Publication of EP0661413A1 publication Critical patent/EP0661413A1/en
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    • 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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • 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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form

Definitions

  • the invention relates to an axial vane grille according to the preamble of claim 1.
  • the object is characterized by the characteristic features of claim 1 solved.
  • the design according to the invention has the advantage that the straight, swept course in addition to that from the prior art Technology known effect, influencing the secondary flow Blade grilles, the radial pressure gradient over the blades in the limit area can be influenced, so that the unwanted The formation of horseshoe vertebrae on the side wall is at least reduced becomes. Since the improvement in step efficiency that can be achieved in this way not solely on influencing the rotor blades of The centrifugal secondary flow is based on how this is done the state of the art is known, the invention in running and Guide vane grilles for compressor and turbine blades Be made use of.
  • Flow-favorable positive arrow angles ⁇ G and ⁇ N between the Blade leading edge and a solder on the radially inner or radial The outer boundary of the ring channel is between 5 ° and 45 °. Allow negative arrow angles ⁇ G and ⁇ N between -10 ° and 0 ° even with structurally difficult conditions with the sweep related benefits.
  • the transition area is designed with a constant curvature and low stress will.
  • the curved section following the straight section up to starting from a relative distance of 25% of the associated bucket height from the respective limit into the interior of the blade.
  • the rotor blades In the case of a design of the rotor blade as a hollow blade to avoid a high bending moment stress under the influence of centrifugal force standing rotor blading of rotor blade grids the rotor blades have cavities that at least overlap extend part of the blade length, the extension of the Cavities are so distributed over the blade profile depth that the The focus of the profile cuts is on a common level. In turbine blades, the cavities can be used as cooling channels be trained.
  • the axial offset allows the individual profile cuts have a favorable influence on the bending moment stress of the airfoil are exerted, the offset can be chosen such that the blade center of gravity in the The line of gravity of the disc comes to rest.
  • the center of gravity is also on the common level. Strength-reducing tensions in the blade root area are thus avoided.
  • the focal points of the blades are such Blade grid and the center of gravity of the shroud so axially spaced from the center of gravity of the vane grille that the center of gravity the disk receiving the rotor blades on the center of gravity of the Blade grid is located. This in turn results in a low bending stress Design in the area of the blade feet.
  • Blade trailing edges result in constant or even Increase or decrease of the airfoil depth over the airfoil length.
  • FIG. 1a to 4b show representations in the z-r plane.
  • the one in figure 1a schematically shown upper half of a two-stage axial turbine has guide and rotor blade grilles arranged axially one after the other in pairs 2 or 3 with concave against the direction of flow S curved guide or blades 4,5 are equipped.
  • the Blades 6 of the guide and moving blades 4.5 extend radially concentric in a to the machine longitudinal axis A of the axial turbine 1 arranged ring channel 7.
  • the axially concentric hub and housing side walls form the radially inner and radially outer channel boundary 8 or 9 of the ring channel 7 and give this one with respect to the flow direction S divergent course.
  • the rotor blade grids 3 are of disc construction, i.e.
  • the blades 5 are each grid-wise on a disk 10 appropriate.
  • 1b shows an axial turbine designed according to the prior art 1 'whose guide and rotor blade grille 2' or 3 'with non-curved Guide and rotor blades 4 'or 5' is equipped.
  • Threading means the positioning of individual profile cuts Pn of an airfoil 6 with respect to a reference line perpendicular to the machine longitudinal axis A, called threading axis F. which, in the case of rotor blades 5, generally runs through the blade center of gravity SP L, is to be understood for the profiling of an airfoil 6.
  • the profile cuts P n coincide with lines of the same relative blade height h in the zr plane.
  • the associated blade height h in turn results from the distance to be measured perpendicular to the longitudinal axis A of the machine between the inner and outer channel boundaries 8 and 9.
  • the blade leading edge 11 has a rectilinear section B G and B N in the rz plane.
  • the arrow angle . ⁇ .G or ⁇ .N to be measured relative to a perpendicular L to the respective channel boundary 8.9 is 25 ° within the sections B G and B N on the housing side and 45 ° on the hub side.
  • the blade leading edge 11 each has a curved extending section Ü G or Ü N , which corresponds to a second-order or higher-order polynomial.
  • the blade edge 11 is again designed to be straight in the rz plane.
  • the transitions from curved to rectilinear course in the blade leading edge 11 are formed continuously.
  • the shape of the trailing edge 12 of the blade results from the specification of the blade depth t (h), which decreases linearly here with increasing duct height h.
  • the rotor blade grille shown in FIG. 3 in the r-z plane 3 is of disk construction, the rotor blades 5 About their molded blade feet 13 positively in evenly in the circumferential direction ⁇ . spaced-apart disk grooves 14 the disc 10 are attached.
  • the centers of gravity SP G and SP S of the rotor blade grille 3 and the disk 10 lying on the machine longitudinal axis A coincide.
  • the focal points SP L of the moving blades 5 lie on a common plane E by corresponding threading of the profile cuts P, which is perpendicular to the machine axis A and runs through the common focal point SPS and SPG of the disk 10 or of the moving blade grid 3.
  • the rotor blade grille 3 is to avoid pressure losses and to improve the flow quality with a in the circumferential direction.
  • segmented shroud 15 provided which includes the blades at the radially outer end. By balancing the shroud segments in the z direction, the centers of gravity SP D of the shroud segments 15 are also on the plane E, whereby bending stresses in the blades 5 are avoided or reduced.
  • Fig. 4 shows an alternative embodiment of a blade 5 to avoid bending stresses in the blade 5 due to unbalanced centers of gravity SP P of the profile cuts P n .
  • the interior of the airfoil 6 has a cavity 16 which extends over the channel height h and whose extension over the airfoil depth t (h) is designed such that the centers of gravity SP P of the profile cuts P n lie in an r-common plane.
  • FIG. 5 shows an additional concave with respect to the blade suction side 18, Blade 6 curved in the circumferential direction.
  • This additional Curvature can advantageously influence the radial Pressure gradients in the outflow plane of a guide or rotor blade 4.5 to take. Due to the circumferential bend, the profile cuts close to the limit aerodynamically relieved. With higher loads at the same time of the center area of the blade 4.5, which makes a total of Blade 4.5 a more efficient efficiency can be achieved.

Description

Die Erfindung betrifft ein Axial-Schaufelgitter nach dem Oberbegriff des Patentanspruches 1.The invention relates to an axial vane grille according to the preamble of claim 1.

Zur Verbesserung der Strömungsverhältnisse in Schaufelgittern von Axial-Strömungsmaschinen ist es aus der EP 0425 889 A1 bekannt, die Schaufelvorderkante der Laufschaufeln im Bereich der Blattspitze gegen die Strömungsrichtung gegenüber einem Vorderkantenverlauf im mittleren Schaufelblattbereich zu neigen und dieser Neigung eine Neigung der Blattspitze in Rotationsrichtung des Laufschaufelgitters zu überlagern. Dieser Schaufelvorderkantenverlauf soll zu einer Verbesserung des Wirkungsgrades des Laufschaufelgitters führen, wobei folgender Effekt zu Nutze gemacht werden soll:To improve the flow conditions in vane grids axial flow machines it is known from EP 0425 889 A1, the blade leading edge of the blades in the area of the blade tip against the flow direction compared to a leading edge course to incline in the middle airfoil area and this inclination an inclination of the blade tip in the direction of rotation of the rotor blade grille to overlay. This blade leading edge course is said to lead to an improvement in the efficiency of the rotor blade grille, the following effect should be used:

Die Neigung der Schaufelvorderkanten gegen die Strömungsrichtung führt zu einem ebenfalls gegen die Strömungsrichtung geneigten Verlauf der Isobaren. Hieraus ergibt sich ein Anstieg des statischen Druckes nach radial außen, wodurch die unter dem Einfluß der Zentrifugalkraft stehende Grenzschicht der Rotorschaufel stromabwärts abgelenkt wird. Dadurch kann ein blattspitzennahes Ablösen der Grenzschicht vermieden werden. The inclination of the blade leading edges against the direction of flow leads to a course that is also inclined against the direction of flow the isobars. This results in an increase in the static Pressure to the radially outward, resulting in the influence of centrifugal force standing boundary layer of the rotor blade is deflected downstream becomes. This can cause the boundary layer to come off near the tip of the blade be avoided.

Hiervon ausgehend ist es Aufgabe der Erfindung, eine für Leit- und Laufschaufelgitter geeignete Schaufelgestaltung zur Verbesserung des Stufenwirkungsgrades anzugeben.Proceeding from this, it is an object of the invention, one for guidance and Blade grid suitable blade design for improvement of the stage efficiency.

Erfindungsgemäß wird die Aufgabe durch die kennzeichnenden Merkmale des Patentanspruches 1 gelöst.According to the invention, the object is characterized by the characteristic features of claim 1 solved.

Die erfindungsgemäße Gestaltung hat den Vorteil, daß durch den geradlinigen, gepfeilten Verlauf zusätzlich zu dem aus dem Stand der Technik bekannten Effekt, der Beeinflussung der Sekundärströmung bei Laufschaufelgittern, der radiale Druckgradient über den Schaufeln im begrenzungsnahen Bereich beeinflußbar ist, so daß die unerwünschte Ausbildung von Hufeisenwirbeln an der Seitenwand zumindest vermindert wird. Da die hierdurch erzielbare Verbesserung des Stufenwirkungsgrades nicht alleine auf der Beeinflussung der bei Rotorschaufeln von der Zentrifugalkraft geprägten Sekundärströmung beruht, wie dies aus dem Stand der Technik bekannt ist, kann die Erfindung bei Lauf- und Leitschaufelgittern von Verdichter- und Turbinenbeschaufelungen zu Nutze gemacht werden. Als Seitenwand wird sowohl die nabenseitige, also radial innere, als auch die gehäuseseitige, also radial äußere Begrenzung des Ringkanals verstanden, wobei diese als in Umfangs- und axiale Richtung ausgebildete Schaufelplattform bzw. als Deckband oder Maschinengehäuse ausgeführt sein kann. Vorzugsweise wird die erfindungsgemäße Gestaltung der Schaufelgitter an den Schaufeln sowohl nabenseitig an den Schaufelfüßen als auch gehäuseseitig an den Schaufelblattspitzen ausgeführt sein. Vorteilhafte Ausführungsformen der Erfindung ergeben sich durch die Merkmale der Patentansprüche 2 bis 13.The design according to the invention has the advantage that the straight, swept course in addition to that from the prior art Technology known effect, influencing the secondary flow Blade grilles, the radial pressure gradient over the blades in the limit area can be influenced, so that the unwanted The formation of horseshoe vertebrae on the side wall is at least reduced becomes. Since the improvement in step efficiency that can be achieved in this way not solely on influencing the rotor blades of The centrifugal secondary flow is based on how this is done the state of the art is known, the invention in running and Guide vane grilles for compressor and turbine blades Be made use of. Both the hub-side, So radially inner, as well as the housing side, ie radially outer Understanding the limitation of the ring channel, these as in circumferential and Axial direction trained blade platform or as a shroud or Machine housing can be executed. Preferably, the invention Design of the blade grille on the blades both on the hub side on the blade feet and on the housing side on the blade tip be executed. Advantageous embodiments of the Invention result from the features of claims 2 to 13.

Eine optimale Beeinflussung des begrenzungsnahen Druckfeldes und der dortigen Sekundärströmung ergibt sich bei geradlinigen Schaufelvorderkantenverlauf innerhalb der naben- bzw. gehäuseseitigen Grenzschicht, wobei je nach Grenzschichtdicke der geradlinige Verlauf innerhalb der Bereiche der Blattspitzen bzw. Schaufelfüße sich bis zu einem Abstand von der jeweiligen Seitenwand zur Schaufelmitte von ca. 10% der zugehörigen Schaufelhöhe erstreckt. Die zu einem Schaufelpunkt zugehörige Schaufelhöhe ergibt sich durch den zur Maschinenlängsachse senkrechten, durch den Schaufelpunkt gehenden Abstand zwischen der radial inneren Begrenzung und der radial äußeren Seitenwand.An optimal influencing of the pressure field close to the limit and the secondary flow there arises in the case of a straight blade leading edge profile within the hub or housing-side boundary layer, the straight line depending on the boundary layer thickness within the areas of the blade tips or blade feet up to a distance from the respective side wall to the blade center of approx. 10% of the associated bucket height extends. That to a shovel point associated bucket height results from the to the machine longitudinal axis vertical distance going through the blade point between the radially inner boundary and the radially outer side wall.

Strömungsgünstige positive Pfeilungswinkel δ G und δ N zwischen der Schaufelvorderkante und einem Lot auf die radial innere bzw. radial äußere Begrenzung des Ringkanals betragen zwischen 5° und 45°. Negative Pfeilungswinkel δ G und δ N zwischen -10° und 0° erlauben auch unter konstruktiv schwierigen Bedingungen die mit der Pfeilung verbundenen Vorteile.Flow-favorable positive arrow angles δ G and δ N between the Blade leading edge and a solder on the radially inner or radial The outer boundary of the ring channel is between 5 ° and 45 °. Allow negative arrow angles δ G and δ N between -10 ° and 0 ° even with structurally difficult conditions with the sweep related benefits.

Im Anschluß an den geradlinigen Verlauf der Schaufelvorderkanten weisen diese in einem Übergangbereich zur Schaufelmitte hin einen gekrümmten Verlauf nach einem Polynom zweiter oder höherer Ordnung auf. Durch Variation der Krümmung kann wiederum der radiale Druckgradient strömungsgünstig manipuliert werden. Darüberhinaus kann bei Rotorschaufeln, die einer hohen Fliehkraftbeanspruchung unterliegen, der Übergangsbereich mit einer stetigen Krümmung spannungsarm gestaltet werden. In einer bevorzugten Ausführung erstreckt sich der gekrümmte Abschnitt im Anschluß an den geradlinigen Abschnitt bis zu einem relativen Abstand von 25% der zugehörigen Schaufelhöhe ausgehend von der jeweiligen Begrenzung ins Schaufelinnere.Following the straight line of the blade leading edges indicate this in a transition area to the center of the blade curved course after a polynomial of second or higher order on. The radial pressure gradient can in turn be varied by varying the curvature can be manipulated in terms of flow. In addition, at Rotor blades, which are subject to high centrifugal force, the transition area is designed with a constant curvature and low stress will. In a preferred embodiment, the curved section following the straight section up to starting from a relative distance of 25% of the associated bucket height from the respective limit into the interior of the blade.

Im Falleeiner Ausführung der Laufschaufel als Hohl schaufel kann zur Vermeidung einer hohen Biegemomentbeanspruchung der unter Fliehkrafteinfluß stehenden Rotorbeschaufelung von Laufschaufelgittern weisen die Rotorschaufeln Hohlräume auf, die sich zumindest über einen Teil der Schaufellänge erstrecken, wobei die Ausdehnung der Hohlräume derart über die Schaufelprofiltiefe verteilt sind, daß die Schwerpunkte der Profilschnitte auf einer gemeinsamen Ebene liegen. Bei Laufschaufelgittern von Turbinen können die Hohlräume als Kühlkanäle ausgebildet sein.In the case of a design of the rotor blade as a hollow blade to avoid a high bending moment stress under the influence of centrifugal force standing rotor blading of rotor blade grids the rotor blades have cavities that at least overlap extend part of the blade length, the extension of the Cavities are so distributed over the blade profile depth that the The focus of the profile cuts is on a common level. In turbine blades, the cavities can be used as cooling channels be trained.

Für massiv ausgeführte Laufschaufeln kann durch axialen Versatz der einzelnen Profilschnitte günstiger Einfluß auf die Biegemomentbeanspruchung des Schaufelblattes ausgeübt werden, wobei der Versatz derart gewählt werden kann, daß der Schaufelschwerpunkt in der Schwerelinie der Scheibe zum Liegen kommt.For solid blades, the axial offset allows the individual profile cuts have a favorable influence on the bending moment stress of the airfoil are exerted, the offset can be chosen such that the blade center of gravity in the The line of gravity of the disc comes to rest.

Bei der Ausführung des Laufschaufelgitters mit einer die Rotorschaufeln aufnehmenden Scheibe liegt der Scheibenschwerpunkt ebenfalls auf der gemeinsamen Ebene. Festigkeitsmindernde Spannungen im Schaufelfußbereich werden somit vermieden. Das gleiche gilt für Laufschaufelgitter mit einem Deckband, welches das Schaufelgitter kanalkonzentrisch umgibt und mit den Schaufel spitzen verbunden ist, bzw. dort angebracht ist. Dabei sind die Schwerpunkte der Schaufeln eines solchen Schaufelgitters und der Schwerpunkt des Deckbandes derart axial vom Schwerpunkt des Schaufelgitters beabstandet, daß der Schwerpunkt der die Rotorschaufeln aufnehmenden Scheibe auf dem Schwerpunkt des Schaufelgitters liegt. Hierdurch ergibt sich wiederum eine biegespannungsarme Gestaltung im Bereich der Schaufelfüße.When designing the rotor blade grille with one, the rotor blades receiving disc, the center of gravity is also on the common level. Strength-reducing tensions in the blade root area are thus avoided. The same applies to blade grilles with a shroud that concentrates the vane grille surrounds and is connected to the blade tip, or there is appropriate. The focal points of the blades are such Blade grid and the center of gravity of the shroud so axially spaced from the center of gravity of the vane grille that the center of gravity the disk receiving the rotor blades on the center of gravity of the Blade grid is located. This in turn results in a low bending stress Design in the area of the blade feet.

Einen dem Verlauf der Schaufelvorderkanten ähnlichen Verlauf der Schaufelhinterkanten ergibt sich bei konstanter oder bei gleichmäßiger Zu- bzw. Abnahme der Schaufelblattiefe über die Schaufelblattlänge.A course similar to the course of the blade leading edges Blade trailing edges result in constant or even Increase or decrease of the airfoil depth over the airfoil length.

Bevorzugte Ausführungsformen der Erfindung werden nachfolgend unter Bezugnahme auf die beigefügte Zeichnung erläutert. Es zeigt:

Fig. 1 a
einen Längsschnitt durch die Niederdruckturbine eines Strahltriebwerkes mit gekrümmten Turbinenschaufeln,
Fig. 1 b
einen Längsschnitt durch die Niederdruckturbine eines Strahltriebwerkes mit geradlinig verlaufenden Turbinenschaufeln,
Fig. 2
einen vergrößerten Ausschnitt eines gekrümmten Schaufelblattes gemäß Fig. 1a,
Fig. 3
einen teilweisen Längsschnitt eines Laufschaufelgitters mit Scheibe und Deckband,
Fig. 4
einen Schnitt durch die Skelettfläche einer Laufschaufel mit hohlem Schaufelblatt und
Fig. 5
eine Ansicht eines Laufschaufelblattes mit Krümmung in Umfangsrichtung
Preferred embodiments of the invention are explained below with reference to the accompanying drawings. It shows:
Fig. 1 a
2 shows a longitudinal section through the low-pressure turbine of a jet engine with curved turbine blades,
Fig. 1 b
2 shows a longitudinal section through the low-pressure turbine of a jet engine with straight-line turbine blades,
Fig. 2
2 shows an enlarged section of a curved airfoil according to FIG. 1a,
Fig. 3
a partial longitudinal section of a rotor blade grille with washer and shroud,
Fig. 4
a section through the skeletal surface of a blade with a hollow blade and
Fig. 5
a view of a blade with curvature in the circumferential direction

Für Richtungs- und Bezugsangaben wird ein in der Strömungsmechanik übliches Axial-Umfangs-Radial-Koordinatensystem z-ϕ.-r verwendet. Die Figuren 1a bis 4b zeigen Darstellungen in der z-r-Ebene. Die in Figur 1a schematisch dargestellte obere Hälfte einer zweistufigen Axialturbine weist paarweise axial hintereinander angeordnete Leit- und Laufschaufelgitter 2 bzw. 3 auf, die mit konkav entgegen der Strömungsrichtung S gekrümmten Leit- bzw. Laufschaufeln 4,5 bestückt sind. Die Schaufelblätter 6 der Leit- und Laufschaufeln 4,5 erstrecken sich radial in einem zur Maschinenlängsachse A der Axialturbine 1 konzentrisch angeordneten Ringkanal 7. Die achskonzentrisch verlaufenden naben- und gehäuseseitigen Seitenwände bilden die radial innere und radial äußere Kanalbegrenzung 8 bzw. 9 des Ringkanals 7 und geben diesem einen bezüglich der Strömungsrichtung S divergenten Verlauf.For directional and reference information, use a in fluid mechanics Usual axial circumference-radial coordinate system z-ϕ.-r used. The Figures 1a to 4b show representations in the z-r plane. The one in figure 1a schematically shown upper half of a two-stage axial turbine has guide and rotor blade grilles arranged axially one after the other in pairs 2 or 3 with concave against the direction of flow S curved guide or blades 4,5 are equipped. The Blades 6 of the guide and moving blades 4.5 extend radially concentric in a to the machine longitudinal axis A of the axial turbine 1 arranged ring channel 7. The axially concentric hub and housing side walls form the radially inner and radially outer channel boundary 8 or 9 of the ring channel 7 and give this one with respect to the flow direction S divergent course.

Die Laufschaufelgitter 3 sind in Scheibenbauweise ausgeführt, d.h., die Laufschaufeln 5 sind jeweils gitterweise an einer Scheibe 10 angebracht.The rotor blade grids 3 are of disc construction, i.e. The blades 5 are each grid-wise on a disk 10 appropriate.

Fig. 1b zeigt eine nach dem Stand der Technik ausgebildete Axialturbine 1' deren Leit- und Laufschaufelgitter 2' bzw. 3' mit ungekrümmten Leit- und Laufschaufeln 4' bzw. 5' bestückt ist.1b shows an axial turbine designed according to the prior art 1 'whose guide and rotor blade grille 2' or 3 'with non-curved Guide and rotor blades 4 'or 5' is equipped.

Fig. 2 zeigt die Auffädelung einzelner Profilschnitte P1,P2, P3 und P4 eines Schaufelblattes 6 der Axialturbine 1. Unter Auffädelung ist die Positionierung einzelner Profilschnitte Pn eines Schaufelblattes 6 bezüglich einer senkrecht auf der Maschinenlängsachse A stehenden Referenzlinie, Fädelachse F genannt, die bei Laufschaufeln 5 im allgemeinen durch den Schaufelschwerpunkt SPL läuft, zur Profilgebung eines Schaufelblattes 6 zu verstehen. Die Profilschnitte Pn fallen in der z-r-Ebene definitionsgemäß mit Linien gleicher relativer Schaufelhöhe h zusammen. Die zugehörige Schaufelhöhe h wiederum ergibt sich aus dem senkrecht zur Maschinenlängsachse A zu messenden Abstand zwischen der inneren und äußeren Kanal begrenzung 8 bzw. 9. Die in Fig. 2 erkenntlichen Profilschnitte P1, P2, P3 und P4 sind bei 5-,25-,75- bzw. 95%iger relativer Schaufel höhe gezogen und trennen Bereiche des Schaufelblattes 6 mit unterschiedlicher Formgebung der Schaufelvorderkante 11 ab. In den Bereichen der Randschnitte PG und PN bis zum Profilschnitt P1 bzw. P4 weist die Schaufelvorderkante 11 einen geradlinig verlaufenden Abschnitt BG bzw. BN in der r-z-Ebene auf. Der gegenüber einer Lotsrechten L auf die jeweilige Kanalbegrenzung 8,9 zu messende Pfeilungswinkel .δ.G bzw. δ.N beträgt innerhalb der Abschnitte BG und BN gehäuseseitig 25° und nabenseitig 45°. Im Anschluß an die geradlinig verlaufenden Abschnitte festgelegten Übergangsbereiche zwischen den Profilschnitten P3 und P4 sowie P1 und P2 weist die Schaufelvorderkante 11 jeweils einen gekrümmten verlaufenden Abschnitt ÜG bzw. ÜN auf, der einem Polynom zweiter oder höherer Ordnung entspricht. Im Mittenbereich zwischen den Profilschnitten P3 und P2 ist die Schaufelkante 11 in der r-z-Ebene wiederum geradlinig ausgeführt. Zur Vermeidung unerwünschter aerodynamischer Effekte und von Spannungskonzentrationen sind die Übergänge von gekrümmten zum geradlinigen Verlauf in der Schaufelvorderkante 11 stetig ausgebildet. Der Verlauf der Schaufelhinterkante 12 ergibt sich durch Vorgabe der Schaufelblattiefe t(h), die hier mit zunehmender Kanalhöhe h linear abnimmt. 2 shows the threading of individual profile cuts P 1 , P 2 , P 3 and P 4 of an airfoil 6 of the axial turbine 1. Threading means the positioning of individual profile cuts Pn of an airfoil 6 with respect to a reference line perpendicular to the machine longitudinal axis A, called threading axis F. which, in the case of rotor blades 5, generally runs through the blade center of gravity SP L, is to be understood for the profiling of an airfoil 6. By definition, the profile cuts P n coincide with lines of the same relative blade height h in the zr plane. The associated blade height h in turn results from the distance to be measured perpendicular to the longitudinal axis A of the machine between the inner and outer channel boundaries 8 and 9. The profile cuts P 1 , P 2 , P 3 and P 4 shown in FIG. , 25%, 75% or 95% relative blade height and separate areas of the blade 6 with different shapes of the blade leading edge 11. In the areas of the edge cuts P G and P N up to the profile cut P 1 and P 4 , the blade leading edge 11 has a rectilinear section B G and B N in the rz plane. The arrow angle .δ.G or δ.N to be measured relative to a perpendicular L to the respective channel boundary 8.9 is 25 ° within the sections B G and B N on the housing side and 45 ° on the hub side. Following the straight-line sections defined transition areas between the profile sections P 3 and P 4 and P 1 and P 2 , the blade leading edge 11 each has a curved extending section Ü G or Ü N , which corresponds to a second-order or higher-order polynomial. In the middle region between the profile cuts P 3 and P 2 , the blade edge 11 is again designed to be straight in the rz plane. In order to avoid undesired aerodynamic effects and stress concentrations, the transitions from curved to rectilinear course in the blade leading edge 11 are formed continuously. The shape of the trailing edge 12 of the blade results from the specification of the blade depth t (h), which decreases linearly here with increasing duct height h.

Das in der Fig. 3 in der r-z-Ebene dargestellte Laufschaufelgitter 3 ist in Scheibenbauweise ausgeführt, wobei die Laufschaufeln 5 über ihre angeformten Schaufelfüße 13 formschlüssig in gleichmäßig in Umfangsrichtung ϕ. voneinander beabstandeten Scheibennuten 14 der Scheibe 10 angebracht sind.The rotor blade grille shown in FIG. 3 in the r-z plane 3 is of disk construction, the rotor blades 5 About their molded blade feet 13 positively in evenly in the circumferential direction ϕ. spaced-apart disk grooves 14 the disc 10 are attached.

Zur Vermeidung unnötiger Biegespannungen während des Betriebes in der Scheibe 10 und in den Laufschaufeln 5 des Laufschaufelgitters 3 fallen die auf der Maschinenlängsachse A liegenden Schwerpunkte SPG und SPS des Laufschaufelgitters 3 bzw. der Scheibe 10 zusammen. In diesem Sinne liegen die Schwerpunkte SPL der Laufschaufeln 5 durch entsprechende Auffädelung der Profilschnitte P auf einer gemeinsamen Ebene E, die senkrecht zur Maschinenachse A steht und durch den gemeinsamen Schwerpunkt SPS und SPG der Scheibe 10 bzw. des Laufschaufelgitters 3 verläuft. Das Laufschaufelgitter 3 ist zur Vermeidung von Druckverlusten und zur Verbesserung der Strömungsqualität mit einem in Umfangsrichtung .ϕ. segmentierten Deckband 15 versehen, welches die Laufschaufeln am radial äußeren Ende umfaßt. Durch balancieren der Deckbandsegmente in z-Richtung liegen die Schwerpunkte SPD der Deckbandsegmente 15 ebenfalls auf der Ebene E, wodurch Biegespannungen in den Laufschaufeln 5 vermieden oder reduziert werden.In order to avoid unnecessary bending stresses during operation in the disk 10 and in the rotor blades 5 of the rotor blade grille 3, the centers of gravity SP G and SP S of the rotor blade grille 3 and the disk 10 lying on the machine longitudinal axis A coincide. In this sense, the focal points SP L of the moving blades 5 lie on a common plane E by corresponding threading of the profile cuts P, which is perpendicular to the machine axis A and runs through the common focal point SPS and SPG of the disk 10 or of the moving blade grid 3. The rotor blade grille 3 is to avoid pressure losses and to improve the flow quality with a in the circumferential direction. segmented shroud 15 provided which includes the blades at the radially outer end. By balancing the shroud segments in the z direction, the centers of gravity SP D of the shroud segments 15 are also on the plane E, whereby bending stresses in the blades 5 are avoided or reduced.

Fig. 4 zeigt eine alternative Ausführung einer Laufschaufel 5 zur Vermeidung von Biegespannungen in der Laufschaufel 5 aufgrund unbalancierter Schwerpunktslagen SPP der Profilschnitte Pn. Hierzu weist das Innere des Schaufelblattes 6 einen sich über die Kanalhöhe h hinweg erstreckenden Hohlraum 16 auf, dessen Erstreckung über die Schaufelblattiefe t(h) derart gestaltet ist, daß die Schwerpunkte SPP der Profilschnitte Pn in einer r-gemeinsamen-Ebene liegen.Fig. 4 shows an alternative embodiment of a blade 5 to avoid bending stresses in the blade 5 due to unbalanced centers of gravity SP P of the profile cuts P n . For this purpose, the interior of the airfoil 6 has a cavity 16 which extends over the channel height h and whose extension over the airfoil depth t (h) is designed such that the centers of gravity SP P of the profile cuts P n lie in an r-common plane.

Fig. 5 zeigt ein zusätzlich bezüglich der Schaufelsaugseite 18 konkav, in Umfangsrichtung gekrümmtes Schaufelblatt 6. Durch diese zusätzliche Krümmung läßt sich vorteilhaft Einfluß auf den radialen Druckgradienten in der Abströmebene einer Leit- oder Laufschaufel 4,5 nehmen. Aufgrund der Umfangsbiegung werden die begrenzugsnahen Profilschnitte aerodynamisch entlastet. Bei gleichzeitiger höheren Belastung des Mittenbereiches der Schaufel 4,5, wodurch insgesamt für die Schaufel 4,5 ein günstigerer Wirkungsgrad erzielt werden kann.5 shows an additional concave with respect to the blade suction side 18, Blade 6 curved in the circumferential direction. This additional Curvature can advantageously influence the radial Pressure gradients in the outflow plane of a guide or rotor blade 4.5 to take. Due to the circumferential bend, the profile cuts close to the limit aerodynamically relieved. With higher loads at the same time of the center area of the blade 4.5, which makes a total of Blade 4.5 a more efficient efficiency can be achieved.

Claims (14)

  1. Cascade (2, 3) for axial flow machines having an annular channel (7), the leading edges (11) of the blades (4, 5) of the cascade (2, 3) arranged in the annular channel (7) being swept back in the axial direction in the zones of the radially inner and/or radially outer channel boundary (8, 9) of the annular channel (7) near the edges, characterised in that in the zones near the edges the blade leading edges (11) have a section (BG or BN) extending in a straight line, and in transition zones at the junction with this section have a section [ÜG or ÜN] extending in a curve.
  2. Cascade according to claim 1, characterised in that the sections (BG or BN) extending in a straight line extend from the blade tip (17) or from the hub section PN of the blade (6) by a distance of up to 30% of the associated blade height (h).
  3. Cascade according to claim 1 or 2, characterised in that positive sweepback angles (δG or δN), measured respectively in a longitudinal section between a section (BG or BN), extending in a straight line, of a blade leading edge (11) and a perpendicular line (L) to a radially outer or radially inner boundary (9, 8) of the annular channel (7) at the intersection with the leading edge of the blade (11), are between 5° and 45°.
  4. Cascade according to one of the preceding claims, characterised in that the curved shape of the blade leading edges (11) corresponds in the transition zones to polynomials of the second or higher order.
  5. Cascade according to one of the preceding claims, characterised in that at the junction with the transition zones in the centre region of the blades (6) the leading blade edges (11) have a section (M) extending in a straight line.
  6. Cascade according to claim 5, characterised in that the sections (ÜG or ÜN) extending in a curve occupy up to 50% of the associated blade height.
  7. Cascade according to one of the preceding claims, characterised in that the blades (6) are curved in the peripheral direction in addition to the sweepback.
  8. Cascade according to one of the preceding claims, characterised in that the blade is a rotor blade cascade (3) with rotor blades (5).
  9. Cascade according to claim 7, characterised in that the rotor blades (5) have hollow chambers (16) which extend at least over part of the length of the blades, the extent of the hollow chambers (16) being distributed in such a way over the blade profile depth (t) that the centres of gravity of the profile sections (Pn) are in a common plane (E) perpendicular to the longitudinal axis (A) of the machine.
  10. Cascade according to claim 9, characterised in that the hollow chambers (16) are cooling channels.
  11. Cascade according to one of claims 8 to 10, characterised in that the rotor blade cascade (3) has a disc (10) receiving the rotor blades (5), the centre of gravity (SPs) of which disc lies in the common plane (E).
  12. Cascade according to one of claims 8 to 11, characterised in that the rotor blade cascade (3) has a cover strip (15) and a disc (10) receiving the rotor blades (5), the centres of gravity (SPL) of the rotor blades (5) and the centre of gravity (SPD) of the cover strip (15) being spaced in such a way axially from the centre of gravity (SPG) of the rotor blade cascade (3) that the centre of gravity (SPs) of the disc (10) lies on the centre of gravity (SPG) of the rotor blade cascade (3).
  13. Cascade according to one of the preceding claims, characterised in that the blade depth t(h) between the leading edge (11) and the trailing edge (12) of the rotor blades (5) of a cascade is constant over the blade height (h) or extends linearly.
  14. Cascade according to one of the preceding claims, characterised in that a rounding-off radius (RN or RG) is superimposed on the shape of the leading edge (11) and/or the trailing edge (12) of the blades (5, 6) in the region of the profile sections (PN or PG) near the edge.
EP19940119705 1993-12-23 1994-12-14 Axial blade cascade with blades of arrowed leading edge Expired - Lifetime EP0661413B1 (en)

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DE19934344189 DE4344189C1 (en) 1993-12-23 1993-12-23 Axial vane grille with swept front edges
DE4344189 1993-12-23

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JPH10103002A (en) * 1996-09-30 1998-04-21 Toshiba Corp Blade for axial flow fluid machine
JP4217000B2 (en) * 1997-09-08 2009-01-28 シーメンス アクチエンゲゼルシヤフト Blades for fluid machinery and steam turbines
DE59709447D1 (en) * 1997-11-17 2003-04-10 Alstom Switzerland Ltd Power stage for turbine with axial flow
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EP0957236A1 (en) * 1998-05-15 1999-11-17 Asea Brown Boveri AG Turbine rotor blade
US6491493B1 (en) 1998-06-12 2002-12-10 Ebara Corporation Turbine nozzle vane
WO2000061918A2 (en) * 1999-03-22 2000-10-19 Siemens Westinghouse Power Corporation Airfoil leading edge vortex elimination device
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Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB719061A (en) * 1950-06-21 1954-11-24 United Aircraft Corp Blade arrangement for improving the performance of a gas turbine plant
DE2034890A1 (en) * 1969-07-21 1971-02-04 Rolls Royce Ltd Derby, Derbyshire (Großbritannien) Blade for axial flow machines
US4012172A (en) * 1975-09-10 1977-03-15 Avco Corporation Low noise blades for axial flow compressors
JPS5447907A (en) * 1977-09-26 1979-04-16 Hitachi Ltd Blading structure for axial-flow fluid machine
FR2556409B1 (en) * 1983-12-12 1991-07-12 Gen Electric IMPROVED BLADE FOR A GAS TURBINE ENGINE AND MANUFACTURING METHOD
GB2164098B (en) * 1984-09-07 1988-12-07 Rolls Royce Improvements in or relating to aerofoil section members for turbine engines
JP2665005B2 (en) * 1989-10-24 1997-10-22 三菱重工業株式会社 Blades of axial flow machines
WO1993005275A1 (en) * 1991-08-30 1993-03-18 Airflow Research And Manufacturing Corporation Forward skew fan with rake and chordwise camber corrections

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