EP0621920B1 - Cooling of the shroud of a turbine blade - Google Patents

Cooling of the shroud of a turbine blade Download PDF

Info

Publication number
EP0621920B1
EP0621920B1 EP94901787A EP94901787A EP0621920B1 EP 0621920 B1 EP0621920 B1 EP 0621920B1 EP 94901787 A EP94901787 A EP 94901787A EP 94901787 A EP94901787 A EP 94901787A EP 0621920 B1 EP0621920 B1 EP 0621920B1
Authority
EP
European Patent Office
Prior art keywords
blade
cooling
cooling air
channel
branch
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.)
Expired - Lifetime
Application number
EP94901787A
Other languages
German (de)
French (fr)
Other versions
EP0621920A1 (en
Inventor
Neil Milner Evans
Paul Hayton
Stephen Hill
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce Deutschland Ltd and Co KG
Original Assignee
BMW Rolls Royce GmbH
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by BMW Rolls Royce GmbH filed Critical BMW Rolls Royce GmbH
Publication of EP0621920A1 publication Critical patent/EP0621920A1/en
Application granted granted Critical
Publication of EP0621920B1 publication Critical patent/EP0621920B1/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • 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
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • 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
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2240/00—Components
    • F05B2240/80—Platforms for stationary or moving blades
    • F05B2240/801—Platforms for stationary or moving blades cooled platforms
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms

Definitions

  • the invention relates to a turbine blade of a gas turbine according to the preamble of claim 1.
  • a turbine blade is shown in GB-A-1 605 335.
  • a further cooled turbine blade with a likewise cooled shroud segment is known from GB-A-1 514 613, cooling air channels being provided here by a cooling branch channel which is connected to a blade cooling channel located in the middle of the blade. that lead to the surface of the shroud segment.
  • a turbine blade according to claim 1 is provided to achieve this object.
  • Advantageous further developments of the invention are the content of the subclaims.
  • a shroud segment extends over the entire blade cross-section, as is known, and can thus become relatively large, effective cooling with only a single cooling air branch duct and the cooling air bores branching therefrom cannot be sufficient.
  • at least two preferably substantially parallel cooling air branch channels are provided, each of which is supplied directly by its own blade cooling air channel and which in each case bring about effective cooling of essentially the entire shroud segment, in particular via branching film cooling holes or convection cooling holes.
  • both the cooling air branch duct and the cooling air bores which have a significantly smaller diameter, can be drilled into the shroud segment.
  • the cooling air holes should open on the surfaces of the shroud segment and thereby form film cooling holes or convection cooling holes
  • the cooling air branch channels should not open on the surface of the shroud segment, since the cooling air has a relatively large cross section -Stich channel an excessively large cooling air partial flow would escape uselessly. Therefore, the ends of each cooling air branch duct, which preferably extends over the entire shroud segment, are closed on the end sides or on the surfaces of the shroud segment. This sealing is preferably carried out by subsequent build-up welding.
  • Reference number 1 denotes a cooled turbine blade of a gas turbine, of which only the blade tip is shown in FIGS. 2 to 5. This turbine blade 1 carries a shroud segment 2.
  • the shape of the edge surfaces 22, 23 creates a positive connection between these turbine blades or their shroud segments 2, so that a circumferential blade reinforcement strip is formed.
  • This middle or rear cooling duct system consists of three meandering cooling air ducts 13.
  • the cooling air duct 3 on the blade inflow side and the further system of cooling air ducts 13 work independently of one another, i. H. the cooling air channels 3, 13 are supplied with cooling air separately from one another.
  • Two cooling air branch ducts 4, 14 and a parallel branch duct 15 are provided in the shroud segment 2.
  • the branch channels 4, 14 and the parallel branch duct 15 run essentially parallel to one another and essentially vertically to the longitudinal axis of the turbine blade 1 and, as can be seen, are essentially in the circumferential direction of the blade reinforcement strip, not shown, formed by a plurality of adjacent cover band segments 2 a conventional turbine blade arrangement.
  • the cooling air branch channel 4 is connected via a connecting channel 6 Blade cooling air duct 3 connected, d. H. the cooling air branch duct 4 is supplied with cooling air from the blade cooling air duct 3.
  • a multiplicity of cooling air bores 7 branch off from the cooling air branch duct 4, which lead to the surface of the shroud segment 2 and open at this surface, thereby forming so-called film cooling holes 8 or convection cooling bores 8. This enables convection cooling in the front area of the shroud segment 2 and, in addition, film cooling of the sealing edge 21 of this shroud segment.
  • the mutually adjacent edge surfaces 22, 23 of the individual shroud segments of mutually adjacent turbine blades are cooled, in particular, by the cooling air stream emerging via the film cooling holes 8 '.
  • the second cooling air branch duct 14 is connected to the blade cooling air duct 13 via a connecting duct 16. Cooling air bores 9 also branch off from the second cooling air branch duct 14 and also open on the surface of the shroud segment 2 as film cooling holes 10 or as convection cooling bores 10. At the same time, the parallel branch duct 15, which brings about an improved distribution of cooling air, is supplied with cooling air via these cooling air bores 9, which ensure cooling of the shroud segment 2 over a large area. In this case, only a part of the cooling air bores 9 extends from the surface of the shroud segment 2 beyond the parallel branch duct 15 to the cooling branch duct 14.
  • the number of cooling air bores 9 connecting the cooling branch duct 14 with the parallel branch duct 15 can be used to determine the cooling air flow entering the parallel branch duct 15.
  • the cooling air branch ducts 4, 14 and the parallel branch duct 15 are of relatively large cross section and are produced by drilling. These channels are closed on the end sides of the shroud segment 2, for example by welding. With the two cooling air branch ducts 4, 14 and the additional parallel branch duct 15 as well as with the film cooling holes 8 and the additional convection cooling holes or film cooling holes 10, uniform, effective cooling results not only of the shroud segment 2, but also of its edge surfaces 22 , 23 and its sealing edge 21. However, a large number of details can be designed quite differently from the exemplary embodiment shown, without leaving the content of the claims.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A cooled turbine blade (1) has a shroud segment (2) in which three cooling air branch channels (4, 14, 15) are susbstantially perpendicularly arranged with respect to the longitudinal axis of the blade and in the circumferential direction of the shrouding of the blade. Cooling air bores (7, 9) which open at the surface of the shroud segmets (7, 9), forming there film cooling holes (8, 10), branch off from the cooling air branch channels (4, 14, 15). Cooling air is supplied to a first branch channel (4) by a blade cooling channel (3) located in the blow area, whereas a second branch channel (14) is coupled to another blade cooling channel (13) and the third branch channel (15) is coupled to the second branch channel (14).

Description

Die Erfindung betrifft eine Turbinenschaufel einer Gasturbine nach dem Oberbegriff des Anspruchs 1.
Eine derartige gekühlte Gasturbinen-Schaufel ist in der GB-A-1 605 335 gezeigt. Eine weitere gekühlte Turbinenschaufel mit einem ebenfalls gekühlten Deckband-Segment ist aus der GB-A-1 514 613 bekannt, wobei hier von einem Kühl-Stichkanal, der mit einem in der Schaufelmitte liegenden Schaufel-Kühlkanal verbunden ist, Kühlluft-Kanäle vorgesehen sind, die zur Oberfläche des Deckband-Segmentes führen.The invention relates to a turbine blade of a gas turbine according to the preamble of claim 1.
Such a cooled gas turbine blade is shown in GB-A-1 605 335. A further cooled turbine blade with a likewise cooled shroud segment is known from GB-A-1 514 613, cooling air channels being provided here by a cooling branch channel which is connected to a blade cooling channel located in the middle of the blade. that lead to the surface of the shroud segment.

Durch eine Kühlung nicht nur der Schaufel, sondern auch des Deckband-Segmentes kann erreicht werden, daß die thermischen Belastungen und geometrischen Verformungen von Schaufel und Segment durch thermische Einflüsse gering gehalten werden. Mit der bekannten Kühlluftführung wird die thermische Beanspruchung einer Turbinenschaufel zwar bereits erheblich reduziert, dennoch sind weitere verbesserte Kühlungsmaßnahmen wünschenswert, die aufzuzeigen sich die vorliegende Erfindung zur Aufgabe gestellt hat.By cooling not only the blade, but also the shroud segment, it can be achieved that the thermal loads and geometric deformations of the blade and segment are kept low by thermal influences. With the known cooling air guide, the thermal stress on a turbine blade is already considerably reduced, but further improved cooling measures are desirable, which the present invention has set out to demonstrate.

Zur Lösung dieser Aufgabe ist eine Turbinenschaufel gemäß Anspruch 1 vorgesehen. Vorteilhafte Aus- und Weiterbildungen der Erfindung sind Inhalt der Unteransprüche.A turbine blade according to claim 1 is provided to achieve this object. Advantageous further developments of the invention are the content of the subclaims.

Da sich ein Deckband-Segment wie bekannt über den gesamten Schaufel-Querschnitt erstreckt und somit relativ großflächig werden kann, kann eine effektive Kühlung nur mit einem einzigen Kühlluft-Stichkanal sowie den davon abzweigenden Kühlluft-Bohrungen nicht ausreichend sein. Erfindungsgemäß sind daher zumindest zwei bevorzugt im wesentlichen parallel verlaufende Kühlluft-Stichkanäle vorgesehen, die jeweils direkt von einem eigenen Schaufel-Kühlluftkanal versorgt werden und die jeweils insbesondere über abzweigende Filmkühllöcher oder Konvektionskühlbohrungen eine wirkungsvolle Kühlung im wesentlichen des gesamten Deckband-Segmentes bewirken. Während dabei schaufel-anströmseitig oder auch im Schaufelmittenbereich in einem Schaufel-Kühlluftkanal noch eine ausreichend große Kühlluftmenge zur Versorgung eines angeschlossenen Kühlluft-Stichkanales vorhanden ist, ist schaufel-abströmseitig im Schaufel-Kühlluftkanal diese Kühlluftmenge bereits soweit reduziert, daß ein Kühlluft-Stichkanal, der von einem schaufel-abströmseitigen Schaufel-Kühlluftkanal versorgt werden würde, kaum mehr Kühlluft erhalten würde. Daher wird vorgeschlagen, im Deckband-Segment nahe des Schaufel-Abströmbereiches zusätzlich einen sog. Parallel-Kanal vorzusehen, der von einem benachbarten Kühlluft-Stichkanal mit Kühlluft versorgt wird und hierzu mit diesem bevorzugt über mehrere Kühlluft-Bohrungen verbunden ist. Diese Bohrungen können dabei die gleichen sein, die als Filmkühllöcher auf der Oberfläche des Deckband-Segmentes münden. Das Innere des Deckband-Segmentes ist damit quasi netzartig von mehreren Kühlluft-Stichkanälen sowie den hiervon abzweigenden Kühlluft-Bohrungen durchzogen, die somit den größten Bereich des Deckband-Segmentes kühlungsmäßig abzudecken in der Lage sind.Since a shroud segment extends over the entire blade cross-section, as is known, and can thus become relatively large, effective cooling with only a single cooling air branch duct and the cooling air bores branching therefrom cannot be sufficient. According to the invention, therefore, at least two preferably substantially parallel cooling air branch channels are provided, each of which is supplied directly by its own blade cooling air channel and which in each case bring about effective cooling of essentially the entire shroud segment, in particular via branching film cooling holes or convection cooling holes. While there is still a sufficiently large amount of cooling air to supply a connected cooling air branch duct on the blade inflow side or in the middle of the blade in the blade cooling air duct, this quantity of cooling air is already reduced to such an extent in the blade cooling air duct that a cooling air branch duct, the would be supplied by a blade cooling air duct downstream of the blade, hardly any more cooling air would be obtained. It is therefore proposed to additionally provide a so-called parallel duct in the shroud segment near the blade outflow region, which is supplied with cooling air from an adjacent cooling air branch duct and for this purpose is preferably connected to it via a plurality of cooling air bores. These holes can be the same be, which open as film cooling holes on the surface of the shroud segment. The interior of the shroud segment is thus quasi network-like with several cooling air branch channels and the cooling air holes branching off from them, which are thus able to cover the largest area of the shroud segment in terms of cooling.

Sowohl der Kühlluft-Stichkanal als auch die demgegenüber einen deutlich geringeren Durchmesser aufweisenden Kühlluft-Bohrungen können durch Bohren in das Deckband-Segment eingebracht werden. Während aber die Kühlluft-Bohrungen an den Oberflächen des Deckband-Segmentes münden sollen und hierbei Filmkühllöcher oder Konvektions-Kühlungsbohrungen bilden, sollten die Kühlluft-Stichkanäle nicht an der Oberfläche des Deckband-Segmentes münden, da über den jeweiligen, einen relativ großen Querschnitt aufweisenden Kühlluft-Stichkanal ein zu großer Kühlluft-Teilstrom nutzlos entweichen würde. Daher sind die Enden jedes sich bevorzugt über das gesamte Deckband-Segment erstreckenden Kühlluft-Stichkanales an den Endseiten bzw. auf den Oberflächen des Deckband-Segmentes verschlossen. Bevorzugt erfolgt dieses Verschließen durch nachträgliches Auftragsschweißen.Both the cooling air branch duct and the cooling air bores, which have a significantly smaller diameter, can be drilled into the shroud segment. However, while the cooling air holes should open on the surfaces of the shroud segment and thereby form film cooling holes or convection cooling holes, the cooling air branch channels should not open on the surface of the shroud segment, since the cooling air has a relatively large cross section -Stich channel an excessively large cooling air partial flow would escape uselessly. Therefore, the ends of each cooling air branch duct, which preferably extends over the entire shroud segment, are closed on the end sides or on the surfaces of the shroud segment. This sealing is preferably carried out by subsequent build-up welding.

Dies sowie weitere Vorteile der Erfindung werden auch aus der im folgenden erläuterten Prinzipskizze eines bevorzugten Ausführungsbeispieles ersichtlich.

Fig. 1
zeigt die Aufsicht auf ein Deckband-Segment einer erfindungsgemäßen Turbinenschaufel,
Fig. 2
den Schnitt A-A aus Fig. 1,
Fig. 3
die Ansicht X aus Fig. 1,
Fig. 4
die Ansicht Y aus Fig. 3, sowie
Fig. 5
die Ansicht Z aus Fig. 3.
This and other advantages of the invention can also be seen from the schematic diagram of a preferred exemplary embodiment explained below.
Fig. 1
shows the top view of a shroud segment of a turbine blade according to the invention,
Fig. 2
the section AA from Fig. 1,
Fig. 3
the view X from FIG. 1,
Fig. 4
the view Y from Fig. 3, and
Fig. 5
the view Z from FIG. 3.

Mit der Bezugsziffer 1 ist eine gekühlte Turbinenschaufel einer Gasturbine bezeichnet, von der in Fig. 2 bis 5 lediglich die Schaufelspitze dargestellt ist. Diese Turbinenschaufel 1 trägt ein Deckband-Segment 2.Reference number 1 denotes a cooled turbine blade of a gas turbine, of which only the blade tip is shown in FIGS. 2 to 5. This turbine blade 1 carries a shroud segment 2.

Über die einzelnen Deckband-Segmente 2 eineinander benachbarter Turbinenschaufeln wird durch die Formgebung der Randflächen 22, 23 eine formschlüssige Verbindung zwischen diesen Turbinenschaufeln bzw. deren Deckband- Segmenten 2 hergestellt, so daß ein umlaufendes Schaufel-verstärkungsband gebildet wird.Via the individual shroud segments 2 of adjacent turbine blades, the shape of the edge surfaces 22, 23 creates a positive connection between these turbine blades or their shroud segments 2, so that a circumferential blade reinforcement strip is formed.

In der Turbinenschaufel 1 verläuft, wie bekannt, ein vorderer schaufelanströmseitiger Schaufel-Kühlluftkanal 3, sowie ein weiteres System von Kühlluftkanälen 13, die dem Schaufelmittenbereich sowie der Schaufelabströmkante zugeordnet sind. Dieses mittlere bzw. hintere Kühlkanal-system besteht dabei aus drei mäanderformig aneinandergereihten Kühlluft-Kanälen 13. Der schaufelanströmseitige Kühlluftkanal 3 sowie das weitere System von Kühl-luftkanälen 13 arbeiten unabhängig voneinander, d. h. die Kühlluftkanäle 3, 13 werden voneinander getrennt mit Kühlluft versorgt.In the turbine blade 1, as is known, runs a front blade cooling air duct 3 on the blade inflow side, and a further system of cooling air channels 13, which are assigned to the blade center region and the blade trailing edge. This middle or rear cooling duct system consists of three meandering cooling air ducts 13. The cooling air duct 3 on the blade inflow side and the further system of cooling air ducts 13 work independently of one another, i. H. the cooling air channels 3, 13 are supplied with cooling air separately from one another.

Im Deckband-Segment 2 sind zwei Kühlluft-Stichkanäle 4, 14 sowie ein Parallel-Stichkanal 15 vorgesehen. Die Stichkanäle 4, 14 sowie der Parallel-Stichanal 15 verlaufen im wesentlichen parallel zueinander sowie im wesentlichen vertikal zur Längsachse der Turbinenschaufel 1 und sind wie ersichtlich im wesentlichen in Umfangsrichtung des durch eine Vielzahl von nebeneinanderliegenden Deckband-Segmenten 2 gebildeten, nicht gezeigten Schaufelverstärkungsbandes einer üblichen Turbinenschaufel-Anordnung orientiert.Two cooling air branch ducts 4, 14 and a parallel branch duct 15 are provided in the shroud segment 2. The branch channels 4, 14 and the parallel branch duct 15 run essentially parallel to one another and essentially vertically to the longitudinal axis of the turbine blade 1 and, as can be seen, are essentially in the circumferential direction of the blade reinforcement strip, not shown, formed by a plurality of adjacent cover band segments 2 a conventional turbine blade arrangement.

Nicht nur die Schaufel-Kühlluftkanäle 3, 13 werden voneinander unabhängig mit Kühlluft versorgt, sondern auch die an die Schaufel-Kühlluftkanäle angeschlossenen Kühlluft-Stichkanäle 4, 14 im Deckband-Segment 2. So ist über einen Verbindungskanal 6 der Kühlluft-Stichkanal 4 mit dem Schaufel-Kühlluftkanal 3 verbunden, d. h. der Kühlluft-Stichkanal 4 wird vom Schaufel-Kühlluftkanal 3 mit Kühlluft versorgt. Vom Kühlluft-Stichkanal 4 zweigen eine Vielzahl von Kühlluft-Bohrungen 7 ab, die zur Oberfläche des Deckband-Segmentes 2 führen und an dieser Oberfläche münden und dabei sog. Filmkühllöcher 8 oder Konvektions-Kühlungsbohrungen 8 bilden. Dies ermöglicht eine Konvek-tions-Kühlung im vorderen Bereich des Deckband-Segmentes 2 und zusätzlich eine Filmkühlung der Dichtkante 21 dieses Deckband-Segmentes. Gleichzeitig werden die einander benachbarten Randflächen 22, 23 der einzelnen Deckband-Segmente einander benachbarter Turbinenschaufeln insbesondere durch den über die Filmkühllöcher 8' austretenden Kühlluftstrom gekühlt.Not only the blade cooling air channels 3, 13 are supplied with cooling air independently of one another, but also the cooling air branch channels 4, 14 connected to the blade cooling air channels in the shroud segment 2. Thus, the cooling air branch channel 4 is connected via a connecting channel 6 Blade cooling air duct 3 connected, d. H. the cooling air branch duct 4 is supplied with cooling air from the blade cooling air duct 3. A multiplicity of cooling air bores 7 branch off from the cooling air branch duct 4, which lead to the surface of the shroud segment 2 and open at this surface, thereby forming so-called film cooling holes 8 or convection cooling bores 8. This enables convection cooling in the front area of the shroud segment 2 and, in addition, film cooling of the sealing edge 21 of this shroud segment. At the same time, the mutually adjacent edge surfaces 22, 23 of the individual shroud segments of mutually adjacent turbine blades are cooled, in particular, by the cooling air stream emerging via the film cooling holes 8 '.

Über einen Verbindungskanal 16 ist det zweite Kühlluft-Stichkanal 14 mit dem Schaufel-Kühlluftkanal 13 verbunden. Auch vom zweiten Kühlluft-Stichkanal 14 zweigen Kühlluft-Bohrungen 9 ab, die als Filmkühllöcher 10 oder als Konvektions-Kühlungsbohrungen 10 ebenfalls an der Oberfläche des Deckband-Segmentes 2 münden. Gleichzeitig wird über diese Kühlluft-Bohrungen 9, die eine großflä-chige Kühlung des Deckband-Segmentes 2 gewährleisten, der Parallel-Stichkanal 15, der eine verbesserte Verteilung von Kühlluft bewirkt, mit Kühlluft versorgt. Dabei erstreckt sich lediglich ein Teil der Kühlluft-Bohrungen 9 von der Oberfläche des Deckband-Segmentes 2 über den Parallel-Stichkanal 15 hinaus bis zum Kühl-Stichkanal 14.The second cooling air branch duct 14 is connected to the blade cooling air duct 13 via a connecting duct 16. Cooling air bores 9 also branch off from the second cooling air branch duct 14 and also open on the surface of the shroud segment 2 as film cooling holes 10 or as convection cooling bores 10. At the same time, the parallel branch duct 15, which brings about an improved distribution of cooling air, is supplied with cooling air via these cooling air bores 9, which ensure cooling of the shroud segment 2 over a large area. In this case, only a part of the cooling air bores 9 extends from the surface of the shroud segment 2 beyond the parallel branch duct 15 to the cooling branch duct 14.

Über die Anzahl dieser den Kühl-Stichkanal 14 mit dem Parallel-Stichkanal 15 verbindenden Kühlluft-Bohrungen 9 ist der in den Parallel-Stichkanal 15 gelangende Kühlluftstrom festlegbar.The number of cooling air bores 9 connecting the cooling branch duct 14 with the parallel branch duct 15 can be used to determine the cooling air flow entering the parallel branch duct 15.

Die Kühlluft-Stichkanäle 4, 14 sowie der Parallel-Stichkanal 15 sind von relativ großem Querschnitt und werden durch Bohren erzeugt. An den Endseiten des Deckband-Segmentes 2 werden diese Kanäle verschlossen, beispielsweise durch Verschweißen. Mit den beiden Kühlluft-Stichkanälen 4, 14 sowie dem zusätzlichen Parallel-Stichkanal 15 sowie mit den Filmkühllöchern 8 und den zusätzlichen Konvektions-Kühllöchern oder Filmkühllöchern 10 ergibt sich eine gleichmäßige wirkungsvolle Kühlung nicht nur des Deckband-Segmentes 2, sondern auch von dessen Randflächen 22, 23 sowie dessen Dichtkante 21. Dabei können jedoch eine Vielzahl von Details durchaus abweichend vom gezeigten Ausführungsbeispiel gestaltet sein, ohne den Inhalt der Patentansprüche zu verlassen.The cooling air branch ducts 4, 14 and the parallel branch duct 15 are of relatively large cross section and are produced by drilling. These channels are closed on the end sides of the shroud segment 2, for example by welding. With the two cooling air branch ducts 4, 14 and the additional parallel branch duct 15 as well as with the film cooling holes 8 and the additional convection cooling holes or film cooling holes 10, uniform, effective cooling results not only of the shroud segment 2, but also of its edge surfaces 22 , 23 and its sealing edge 21. However, a large number of details can be designed quite differently from the exemplary embodiment shown, without leaving the content of the claims.

Claims (4)

  1. A turbine blade of a gas turbine with at least one cooling-air channel (3) running in the blade (1) and with a shroud segment (2) arranged at the tip of the blade, which together with other segments on adjacent blades forms a blade strengthening band, with further on the pressure side of the blade (2) and essentially vertical to the axis of the blade, a cooling-air branch channel (4) joined to the blade cooling-air channel (3) in the pressure side of the blade, from which several cooling-air borings (7) lead to the surface of the shroud segment (2), and a further cooling branch channel (14) running within the shroud segment (2) which is joined to a blade cooling channel (13) in the centre of the blade or on the downstream side of the blade, from which similarly cooling-air borings (9) lead to the surface of the blade segment (2),
    characterised in that at least one further parallel branch channel (15) is provided, running essentially parallel to the cooling branch channel (14), which is joined to the cooling branch channel (14) via several cooling-air borings (9).
  2. A turbine blade in accordance with Claim 1,
    characterised in that the branch channels (4, 14, 15) are oriented essentially in the direction of the enclosing blade strengthening band and are closed at both end sides of the shroud segment (2).
  3. A turbine blade in accordance with one of the preceding claims,
    characterised in that the cooling-air borings (7, 9) run essentially vertical to the axis of the blades and extend from the surface of the shroud segment (2) to a branch channel (4, 15) or extend beyond it.
  4. A turbine blade according to claim 3,
    characterised in that the parallel branch channel (15) lies in the blade's downstream region.
EP94901787A 1992-11-19 1993-11-10 Cooling of the shroud of a turbine blade Expired - Lifetime EP0621920B1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
GB9224241 1992-11-19
GB929224241A GB9224241D0 (en) 1992-11-19 1992-11-19 A turbine blade arrangement
PCT/EP1993/003146 WO1994011616A1 (en) 1992-11-19 1993-11-10 Cooling of the shroud of a turbine blade

Publications (2)

Publication Number Publication Date
EP0621920A1 EP0621920A1 (en) 1994-11-02
EP0621920B1 true EP0621920B1 (en) 1996-03-20

Family

ID=10725339

Family Applications (1)

Application Number Title Priority Date Filing Date
EP94901787A Expired - Lifetime EP0621920B1 (en) 1992-11-19 1993-11-10 Cooling of the shroud of a turbine blade

Country Status (5)

Country Link
US (1) US5460486A (en)
EP (1) EP0621920B1 (en)
DE (1) DE59301968D1 (en)
GB (1) GB9224241D0 (en)
WO (1) WO1994011616A1 (en)

Families Citing this family (41)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2290833B (en) * 1994-07-02 1998-08-05 Rolls Royce Plc Turbine blade
US5482435A (en) * 1994-10-26 1996-01-09 Westinghouse Electric Corporation Gas turbine blade having a cooled shroud
GB2298245B (en) * 1995-02-23 1998-10-28 Bmw Rolls Royce Gmbh A turbine-blade arrangement comprising a cooled shroud band
GB2298246B (en) * 1995-02-23 1998-10-28 Bmw Rolls Royce Gmbh A turbine-blade arrangement comprising a shroud band
JP3178327B2 (en) * 1996-01-31 2001-06-18 株式会社日立製作所 Steam turbine
US5785496A (en) * 1997-02-24 1998-07-28 Mitsubishi Heavy Industries, Ltd. Gas turbine rotor
JPH1113402A (en) * 1997-06-23 1999-01-19 Mitsubishi Heavy Ind Ltd Tip shroud for gas turbine cooling blade
JP2955252B2 (en) * 1997-06-26 1999-10-04 三菱重工業株式会社 Gas turbine blade tip shroud
JP3510467B2 (en) * 1998-01-13 2004-03-29 三菱重工業株式会社 Gas turbine blades
DE69931088T2 (en) * 1998-02-04 2006-12-07 Mitsubishi Heavy Industries, Ltd. Gas turbine rotor blade
EP1013884B1 (en) 1998-12-24 2005-07-27 ALSTOM Technology Ltd Turbine blade with actively cooled head platform
DE19904229A1 (en) * 1999-02-03 2000-08-10 Asea Brown Boveri Cooled turbine blade has shroud formed by sealing rib with integrated cooling channels connected to coolant channel in blade
EP1041247B1 (en) * 1999-04-01 2012-08-01 General Electric Company Gas turbine airfoil comprising an open cooling circuit
US6761534B1 (en) 1999-04-05 2004-07-13 General Electric Company Cooling circuit for a gas turbine bucket and tip shroud
DE19963377A1 (en) * 1999-12-28 2001-07-12 Abb Alstom Power Ch Ag Turbine blade with actively cooled cover band element
DE10016081A1 (en) * 2000-03-31 2001-10-04 Alstom Power Nv Plate-shaped, projecting component section of a gas turbine
DE10064265A1 (en) 2000-12-22 2002-07-04 Alstom Switzerland Ltd Device and method for cooling a platform of a turbine blade
JP2002201913A (en) * 2001-01-09 2002-07-19 Mitsubishi Heavy Ind Ltd Split wall of gas turbine and shroud
US6506022B2 (en) * 2001-04-27 2003-01-14 General Electric Company Turbine blade having a cooled tip shroud
US6887033B1 (en) * 2003-11-10 2005-05-03 General Electric Company Cooling system for nozzle segment platform edges
EP1591625A1 (en) * 2004-04-30 2005-11-02 ALSTOM Technology Ltd Gas turbine blade shroud
EP1591626A1 (en) * 2004-04-30 2005-11-02 Alstom Technology Ltd Blade for gas turbine
EP1630354B1 (en) 2004-08-25 2014-06-18 Rolls-Royce Plc Cooled gas turbine aerofoil
US7686581B2 (en) * 2006-06-07 2010-03-30 General Electric Company Serpentine cooling circuit and method for cooling tip shroud
US7762773B2 (en) * 2006-09-22 2010-07-27 Siemens Energy, Inc. Turbine airfoil cooling system with platform edge cooling channels
US7611324B2 (en) * 2006-11-30 2009-11-03 General Electric Company Method and system to facilitate enhanced local cooling of turbine engines
CH699593A1 (en) * 2008-09-25 2010-03-31 Alstom Technology Ltd Blade for a gas turbine.
GB0901129D0 (en) * 2009-01-26 2009-03-11 Rolls Royce Plc Rotor blade
CH700686A1 (en) * 2009-03-30 2010-09-30 Alstom Technology Ltd Blade for a gas turbine.
US8356978B2 (en) * 2009-11-23 2013-01-22 United Technologies Corporation Turbine airfoil platform cooling core
US9759070B2 (en) * 2013-08-28 2017-09-12 General Electric Company Turbine bucket tip shroud
EP3329100B1 (en) * 2015-07-31 2022-12-14 General Electric Company Cooling arrangements in tip shrouded turbine rotor blades
US10156145B2 (en) 2015-10-27 2018-12-18 General Electric Company Turbine bucket having cooling passageway
US10508554B2 (en) 2015-10-27 2019-12-17 General Electric Company Turbine bucket having outlet path in shroud
US9885243B2 (en) 2015-10-27 2018-02-06 General Electric Company Turbine bucket having outlet path in shroud
US10202852B2 (en) 2015-11-16 2019-02-12 General Electric Company Rotor blade with tip shroud cooling passages and method of making same
US10301945B2 (en) * 2015-12-18 2019-05-28 General Electric Company Interior cooling configurations in turbine rotor blades
US10184342B2 (en) * 2016-04-14 2019-01-22 General Electric Company System for cooling seal rails of tip shroud of turbine blade
US10502069B2 (en) * 2017-06-07 2019-12-10 General Electric Company Turbomachine rotor blade
US10577945B2 (en) * 2017-06-30 2020-03-03 General Electric Company Turbomachine rotor blade
US20190085706A1 (en) * 2017-09-18 2019-03-21 General Electric Company Turbine engine airfoil assembly

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1245518A (en) * 1957-04-19 1960-11-10 Improvements made to hot gaseous fluid turbines
US3529902A (en) * 1968-05-22 1970-09-22 Gen Motors Corp Turbine vane
BE755567A (en) * 1969-12-01 1971-02-15 Gen Electric FIXED VANE STRUCTURE, FOR GAS TURBINE ENGINE AND ASSOCIATED TEMPERATURE ADJUSTMENT ARRANGEMENT
GB1605335A (en) * 1975-08-23 1991-12-18 Rolls Royce A rotor blade for a gas turbine engine
US4017213A (en) * 1975-10-14 1977-04-12 United Technologies Corporation Turbomachinery vane or blade with cooled platforms
GB1514613A (en) * 1976-04-08 1978-06-14 Rolls Royce Blade or vane for a gas turbine engine
US4353679A (en) * 1976-07-29 1982-10-12 General Electric Company Fluid-cooled element
JP2862536B2 (en) * 1987-09-25 1999-03-03 株式会社東芝 Gas turbine blades
GB2223276B (en) * 1988-09-30 1992-09-02 Rolls Royce Plc Turbine aerofoil blade
GB2228540B (en) * 1988-12-07 1993-03-31 Rolls Royce Plc Cooling of turbine blades
GB2250548A (en) * 1990-12-06 1992-06-10 Rolls Royce Plc Cooled turbine aerofoil blade
US5344283A (en) * 1993-01-21 1994-09-06 United Technologies Corporation Turbine vane having dedicated inner platform cooling

Also Published As

Publication number Publication date
US5460486A (en) 1995-10-24
EP0621920A1 (en) 1994-11-02
GB9224241D0 (en) 1993-01-06
WO1994011616A1 (en) 1994-05-26
DE59301968D1 (en) 1996-04-25

Similar Documents

Publication Publication Date Title
EP0621920B1 (en) Cooling of the shroud of a turbine blade
DE69815735T2 (en) Cooled gas turbine blade
EP1223308B1 (en) Turbomachine component
DE3211139C1 (en) Axial turbine blades, in particular axial turbine blades for gas turbine engines
DE69823236T2 (en) DEVICE FOR COOLING GAS TURBINE SHOVELS AND METHOD FOR THE PRODUCTION THEREOF
DE1946535C3 (en) Component for a gas turbine engine
DE10355449B4 (en) turbine blade
EP0798448B1 (en) System and device to cool a wall which is heated on one side by hot gas
DE2413292A1 (en) SHEET COOL INSERT HOLDER FOR TURBO MACHINERY
DE2241194A1 (en) FLOW MACHINE SHOVEL WITH A WING-SHAPED CROSS-SECTIONAL PROFILE AND WITH A NUMBER OF COOLING DUCTS RUNNING IN THE LENGTH DIRECTION OF THE SHOVEL
EP0985802A1 (en) Film cooling orifice and it's method of manufacture
DE69828023T2 (en) DECKBAND FOR COOLED GAS TURBINE BOOMS
DE19601818A1 (en) Turbine blade arrangement with a cover band
DE112020002736T5 (en) FLUID COMPRESSOR AND METHOD OF OPERATING A FLUID COMPRESSOR TO REDUCE OIL CARRY-OVER THROUGH A COMPRESSOR PISTON ASSEMBLY
EP0931600B1 (en) Device for descaling rolled material
DE602004006035T2 (en) Cooling device for turbine disks
DE2155344A1 (en) INTEGRAL TURBINE WHEEL WITH OPEN AXIAL BREAKTHROUGHTS ON THE OUTER WREATH AND CONTROLLED WREATH Cracks
EP1288435B1 (en) Turbine blade with at least one cooling orifice
DE1118374B (en) Electroerosion process for the production of recesses in workpieces made of conductive material
EP1138878B1 (en) Gas turbine component
DE19601819A1 (en) Turbine blade arrangement with a cooled shroud
DE2048983A1 (en) Axia) flow machine through which it flows
DE3504343A1 (en) Saw blade
DE102008024434A1 (en) Device for acting on sheet material by means of pressurized media
DE3709666C1 (en) Spray head of a spray device for release agents

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 19940727

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): DE FR GB IT SE

17Q First examination report despatched

Effective date: 19950526

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): DE FR GB IT SE

ET Fr: translation filed
REF Corresponds to:

Ref document number: 59301968

Country of ref document: DE

Date of ref document: 19960425

GBT Gb: translation of ep patent filed (gb section 77(6)(a)/1977)

Effective date: 19960328

ITF It: translation for a ep patent filed
PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed
REG Reference to a national code

Ref country code: FR

Ref legal event code: CD

REG Reference to a national code

Ref country code: FR

Ref legal event code: CJ

Ref country code: FR

Ref legal event code: CA

REG Reference to a national code

Ref country code: GB

Ref legal event code: IF02

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 20041010

Year of fee payment: 12

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 20041014

Year of fee payment: 12

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: SE

Payment date: 20041020

Year of fee payment: 12

Ref country code: DE

Payment date: 20041020

Year of fee payment: 12

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: IT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20051110

Ref country code: GB

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20051110

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20051111

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20060601

EUG Se: european patent has lapsed
GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 20051110

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: FR

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20060731

REG Reference to a national code

Ref country code: FR

Ref legal event code: ST

Effective date: 20060731