EP0313826A1 - Axial gas turbine - Google Patents

Axial gas turbine Download PDF

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Publication number
EP0313826A1
EP0313826A1 EP88115694A EP88115694A EP0313826A1 EP 0313826 A1 EP0313826 A1 EP 0313826A1 EP 88115694 A EP88115694 A EP 88115694A EP 88115694 A EP88115694 A EP 88115694A EP 0313826 A1 EP0313826 A1 EP 0313826A1
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EP
European Patent Office
Prior art keywords
rotor
blade
cooling air
ring
last
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Granted
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EP88115694A
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German (de)
French (fr)
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EP0313826B1 (en
Inventor
Franz Kreitmeier
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BBC Brown Boveri AG Switzerland
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BBC Brown Boveri AG Switzerland
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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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/084—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades the fluid circulating at the periphery of a multistage rotor, e.g. of drum type
    • 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • 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
    • F01D5/145—Means for influencing boundary layers or secondary circulations

Definitions

  • the present invention relates to an axially flow-through gas turbine with cooling devices for the turbine rotor and its rotor blades, the cooling air being branched off from the compressor and accelerated in a known manner by a swirl device in the circumferential direction so that it is opposite cooling air holes on the turbine rotor, through which the cooling air into the Cooling air system flows in, has zero speed in the circumferential direction.
  • a gas turbine with cooling thereof allows a higher gas inlet temperature, which increases the efficiency and the performance.
  • the cooling air duct and the cooling air flow and its distribution over the length of the turbine rotor depend on the gas temperatures prevailing in the individual stages of the turbine.
  • the heated cooling air exits into the gas flow.
  • the gas temperature has already dropped so far that the internal cooling of the rotor blades can be dispensed with. You only get cooling in the area of the blade roots through the air flowing towards the end of the rotor body, which exits into the already largely relaxed propellant gas stream before and after the foot area of the last row of blades and reaches the exhaust gas diffuser with it.
  • the cooling air is taken from the compressor after its last stage and reaches a row of axial bores distributed along the circumference of a flat annular surface of the rotor before the first turbine stage along the outer surface of the section of the shaft or drum located between the compressor and the turbine.
  • the cooling air flow passes through these bores into the cooling channels of the rotor, at the end of which it, reduced by the portion branched off for cooling the hottest rotor blades, exits into the propellant gas flow and with it into the diffuser.
  • the inflow of the cooling air to the rotor is essentially swirl-free, i.e. without a peripheral component, in the direction of rotation of the drum, it is accelerated on its way to the rotor by the friction on the circumferential surface of the drum in its circumferential direction, albeit in relation to The peripheral speed is not very strong, so that there is still a large difference in speed at the entry into the bores mentioned and into the rotor cooling channels. It must therefore be accelerated to the circumferential rotor speed. The drum and the rotor must therefore perform pumping work, which moreover increases the cooling air temperature. Like most of the flow through the cooling channels, this represents a loss factor.
  • Another loss is associated with the cooling air flow exiting the blade root of the last stage. It enters the propellant gas flow with a radially, tangentially and axially directed velocity component and forces it radially away, so that the hub boundary layer at the diffuser inlet suffers a thickening that is harmful to the recovery.
  • the present invention arose from the task of guiding the rotor and blade cooling air as well as the rotor disk cooling air in their outlet areas at the rotor end into the diffuser in such a way that their velocity vectors correspond to that of the average exhaust gas flow in the areas mentioned in terms of amount and direction essentially coincide.
  • the working capacity of the rotor cooling air should be largely used.
  • This guide is also intended to cool the rotor jacket in the area of the last stage with the same amount of rotor cooling air than is the case with the known constructions.
  • the disc cooling air quantity can thereby be reduced, which reduces the temperature differences within the rotor and thus the thermal stresses in order to achieve an extension of the service life of the turbine rotor.
  • the axially flowed through gas turbine is characterized in that for the cooling air duct in the In the area of the last stage, channels are provided which run in the area of the guide vane ring of the last stage in the rotor casing and in the area of the rotor vane ring of the last stage in its blade roots, a cooling air vane grille being present in a cooling vane ring attached to the turbine rotor, at least at the end of the last rotor vane ring Channels are oriented so that the speed vectors of the cooling air exiting into the diffuser essentially coincide with the average speed vector of the exhaust gas flow, and the limits for the outflow of the cooling air into the diffuser are designed in such a way that their separation is avoided and the propellant gas flow in the hub area of the last blade ring is homogenized.
  • Fig. 1 shows a part of a turbine rotor 1, which is composed of forged rotor disks 2, 3, 4, which along with each other on the end faces forged rings who are welded.
  • the blades of the rotor blade rings 5 to 9 are inserted in a known manner with their base of double hammer head profile into the correspondingly profiled blade fastening grooves.
  • guide vanes of guide blade rings 11 to 14 are anchored in a guide blade carrier 10 in a manner similar to the rotor blades in the rotor.
  • the guide vane attachments are only indicated schematically.
  • the last stage of the compressor (not shown) is located to the right of the first rotor blade ring 5 of the turbine -
  • the required cooling air flow is removed, whereupon it is given a tangential speed component, which is equal to the peripheral speed of the rotor cooling channels, by a swirl vane grille arranged between the compressor and the first turbine stage, which is described in the aforementioned DE-A-34 24 139.
  • the cooling air then enters the cooling duct system of the turbine at a relative speed of zero in the circumferential direction substantially axially, as indicated by the speed arrow 16, through a series of cooling air bores 15.
  • the cooling air bores 15 which are provided in large numbers distributed over an annular, flat end face 17 in front of the first rotor blade ring, the cooling air passes into an annular groove 18 which widens in cross-section to its circumference, and from this through a series of interrupted annular gaps 19 in front of the first rotor blade ring 5 and between two of the following rotor blade rings as well as through channels 20 in the area of the blade roots finally into blade root channels 21 of the last rotor blade ring 9.
  • the annular gaps 19 are delimited by the circumferential surfaces of the Rotor jacket and by asymmetrical heat accumulation segments 22, 23, which are located between two rotor blade rings and protect the rotor jacket and the rotor blade feet from overheating by the propellant gas flow.
  • the blade root passages 20, 21 can expediently be formed from two grooves in the two blades, which adjoin each other in the circumferential direction and adjoin one another in the circumferential direction, and which result in closed passages. In the case of the almost axially directed blade roots, these channels can also be provided in the blade grooves themselves, as in the blades of the last rotor blade ring 9.
  • the guide and rotor blades of the most temperature-loaded stages are designed as hollow blades with air cooling.
  • the cooling air is branched off at the blade roots from the cooling air flow described.
  • the elements of the blade cooling are not shown in FIG. 1.
  • the cooling air passes from the blade root channels 21 of the last moving blade ring 9 into a cooling air blade ring 27, which is attached to the rotor body and which has a frusto-conical rotor blade grille 28 just inside its circumference, which, evenly distributed over its circumference, has cooling air blades 31 which are preceded by a rectifier ring 29 which consists of honeycomb-shaped channels 30 distributed over the entire flow cross-section.
  • Fig. 2 shows the circled detail II of Fig. 1 on a larger scale and Fig. 3 shows the development of the section III-III shown in Fig. 2 in the form of a conical shell placed through the center of the channel.
  • the rectifier ring 29 has the task of homogenizing the cooling air jets emerging from the blade root channels 21 of the last rotor blades 9 in order to obtain a flow in the channels delimited by the blades 31 that is as free as possible from separation.
  • the cooling air vane ring 27 fulfills part of the object of the invention presented in the introduction by deflecting the flow threads of the cooling air flow in such a way that their speed vectors over the entire circumference of the diffuser hub essentially coincide with the average speed vector of the exhaust gas flow with the loss-reducing effect described at the outset, by the Low-energy boundary layer is supplied with energy at the diffuser hub and its detachment point is shifted downstream. At the same time, the energy of the rotor cooling air is partially used to deliver work to the rotor.
  • the second measure according to the invention consists in that the cooling air used for cooling the last rotor disk 4 and branched off from the compressor, such as the blade cooling air, flows out into the diffuser in a guided manner.
  • the disk cooling air passes through two disk air channels 33 provided in an outer turbine housing base 32 into one bounded by the bottom 32 and an inner turbine housing base 34 th disk-shaped cavity 35, is, as indicated by the speed arrows, deflected radially inwards against the rotor axis in this and passes through a series of inner disk air channels 36 provided near the axle in front of the rotor disk 4, where its main part is deflected upward and via an annular gap 37 and an annular space 38 is blown out through an annular slot 39 into the hub boundary layer.
  • the convexly curved inlet area 40 of the diffuser hub 41 which sucks in the outflowing disk cooling air together with the reactor cooling air through its curvature, also contributes to the intended inflow into the hub boundary layer.
  • the frustoconical lateral surface 64 of the cooling air vane ring 27 is designed to be inclined with respect to the rotor axis and its length is such that the exhaust gas flow behind the last rotor vane ring 9 is homogenized.
  • a small part of the disk cooling air flowing in through the channel 36 blocks the labyrinth 41 on the end shield.
  • FIG. 4 and 5 show a second embodiment of the rotor cooling air duct.
  • the cooling air enters via a rotor-fixed intermediate channel 44 into a blade grille 45 of a rotor-fixed blade grille ring 46 and out of this into a blade grille 47 of a guide-blade fixed blade grille ring 48, from which it is deflected into end channels 49.
  • the inlet parts of the latter consist of the front half 50 of a vane grille, the profile lugs, in a rotor-fixed vane grille ring 50 ', and the exit region from the rear half 51 of this vane grille in the cooling air vane ring 53.
  • the end channels 49 are shown in FIG.
  • FIG. 6 Another embodiment of the invention is shown in FIG. 6. After the penultimate rotor blade ring 43, the cooling air is guided axially essentially to the end of the rotor blade ring 9 and is only then blown out into the exhaust gas flow in the desired direction by a cooling air blade ring 63. After the penultimate rotor blade ring 43, it again passes, as in the embodiment according to FIG.
  • the end channels 61 extend between the two vane grids 60 and 61, preferably inclined at an angle to an axis parallel.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

Die Kühlluftführung der axial durchströmten Gasturbine verläuft im Bereich der letzten Beschaufelungsstufe (9 + 14) radial einwärts der Wärmestausegmente (23, 24) innerhalb der Aussenbegrenzung des Rotors (4) und durch Schaufelfusskanäle (21) in den Schaufelfüssen des letzten Laufschaufelkranzes (9) und schliesslich durch einen rotorfesten Kühlluftschaufelkranz (28) in den Diffusor, in den der Kühlluftstrom mit einem Geschwindigkeitsvektor eintritt, der mit dem mittleren Geschwindigkeitsvektor des in den Diffusor eintretenden Abgasstromes im wesentlichen übereinstimmt. Dadurch vermeidet man die Strömungsverluste, die mit einem Austreten des Kühlluftstromes in den Abgasstrom bereits im Bereich der letzten Stufe oder letzten Stufen auftreten. Gleichzeitig wird auf diese Weise die Temperaturdifferenz zwischen dem Rotormantel und der ebenfalls durch Verdichterzapfluft gekühlten letzten Rotorscheibe (4) verringert, wodurch auch die Wärmespannungen im Rotor reduziert werden.In the area of the last blading stage (9 + 14), the cooling air duct of the gas turbine with axial flow runs radially inwards of the heat accumulation segments (23, 24) within the outer boundary of the rotor (4) and through blade root ducts (21) in the blade roots of the last blade ring (9) and finally through a rotor-fixed cooling air scoop ring (28) into the diffuser, into which the cooling air flow enters with a speed vector which essentially corresponds to the mean speed vector of the exhaust gas flow entering the diffuser. This avoids the flow losses which already occur in the area of the last stage or stages when the cooling air flow emerges into the exhaust gas flow. At the same time, the temperature difference between the rotor jacket and the last rotor disk (4), which is also cooled by compressor bleed air, is reduced, as a result of which the thermal stresses in the rotor are also reduced.

Description

Die vorliegende Erfindung betrifft eine axial durchströmte Gasturbine mit Kühleinrichtungen für den Turbinenrotor und seine Laufschaufelkränze, wobei die Kühlluft aus dem Ver­dichter abgezweigt und auf bekannte Weise durch eine Drall­einrichtung in Umfangsrichtung so beschleunigt wird, dass sie gegenüber Kühlluftbohrungen am Turbinenrotor, durch welche die Kühlluft in das Kühlluftsystem einströmt, in der Umfangsrichtung die Geschwindigkeit Null hat.The present invention relates to an axially flow-through gas turbine with cooling devices for the turbine rotor and its rotor blades, the cooling air being branched off from the compressor and accelerated in a known manner by a swirl device in the circumferential direction so that it is opposite cooling air holes on the turbine rotor, through which the cooling air into the Cooling air system flows in, has zero speed in the circumferential direction.

Bei Gasturbinen hoher Leistungdichte kommt der Kühlung der hochtemperaturbeanspruchten Bauteile, das sind die Beschau­felung, insbesondere die Laufschaufeln, die neben hohen Temperaturen und Gaskräften auch durch Zentrifugalkräfte beansprucht sind, sowie der Rotor, besondere Bedeutung zu. Dies im Hinblick auf den Wirkungsgrad, der u.a. von der Eintrittstemperatur der Treibgase abhängt. Die höchstzulässige Eintrittstemperatur ist durch die zu erreichende Lebensdauer der wärmebeanspruchten Bauteile begrenzt.In the case of gas turbines with a high power density, the cooling of the components exposed to high temperatures, that is the blading, in particular the moving blades, which in addition to high temperatures and gas forces are also stressed by centrifugal forces, and the rotor are of particular importance. This with regard to the efficiency, which depends, among other things, on the inlet temperature of the propellant gases. The maximum permissible inlet temperature is limited by the lifespan of the heat-stressed components.

Gegenüber einer Gasturbine ohne Kühlung dieser Teile erlaubt eine Gasturbine mit Kühlung derselben eine höhere Gasein­trittstemperatur, was den Wirkungsgrad und die Leistung erhöht.Compared to a gas turbine without cooling these parts, a gas turbine with cooling thereof allows a higher gas inlet temperature, which increases the efficiency and the performance.

Stand der TechnikState of the art

Bei den bekannten Industriegasturbinen hängen die Kühlluft­führung und der Kühlluftstrom und seine Verteilung über die Länge des Turbinenrotors von den in den einzelnen Stufen der Turbine herrschenden Gastemperaturen ab. Für die ersten, am höchsten temperaturbeanspruchten Stufen kann es erforder­lich sein, die Laufschaufeln von innen her zu kühlen, indem aus der den Rotorkörper umströmenden Kühlluft ein Teil in Kühlkanäle hinein abgezweigt wird, die die betreffenden Laufschaufeln in ihrer Längserstreckung durchsetzen. Am Schaufelende tritt die erwärmte Kühlluft in den Treibgas­strom aus. In den auf die letzte gekühlte Schaufel folgenden Stufen ist die Gastemperatur bereits so weit gesunken, dass auf die innere Kühlung der Laufschaufeln verzichtet werden kann. Sie erhalten lediglich Kühlung im Bereich der Schaufel­füsse durch die am Umfang des Rotorkörpers zu seinem Ende hin strömende Luft, die dort vor und nach dem Fussbereich der letzten Laufschaufelreihe in den bereits weitgehend entspannten Treibgasstrom austritt und mit diesem in den Abgasdiffusor gelangt.In the known industrial gas turbines, the cooling air duct and the cooling air flow and its distribution over the length of the turbine rotor depend on the gas temperatures prevailing in the individual stages of the turbine. For the first stages, which are subject to the highest temperatures, it may be necessary to cool the rotor blades from the inside, by branching part of the cooling air flowing around the rotor body into cooling channels which pass through the rotor blades in question in their longitudinal extent. At the end of the blade, the heated cooling air exits into the gas flow. In the stages following the last cooled blade, the gas temperature has already dropped so far that the internal cooling of the rotor blades can be dispensed with. You only get cooling in the area of the blade roots through the air flowing towards the end of the rotor body, which exits into the already largely relaxed propellant gas stream before and after the foot area of the last row of blades and reaches the exhaust gas diffuser with it.

Die Kühlluft wird dem Verdichter nach seiner letzten Stufe entnommen und gelangt drallfrei entlang der Mantelfläche des zwischen Verdichter und Turbine befindlichen Abschnitts der Welle oder Trommel in eine Reihe axialer Bohrungen, die über den Umfang einer planen Ringfläche des Rotors ver­teilt vor der ersten Turbinenstufe vorhanden sind. Ueber diese Bohrungen gelangt der Kühlluftstrom in die Kühlkanäle des Rotors, an dessen Ende er, verringert um den zur Kühlung der heissesten Laufschaufeln abgezweigten Anteil, in den Treibgasstrom austritt und mit diesem in den Diffusor ge­langt.The cooling air is taken from the compressor after its last stage and reaches a row of axial bores distributed along the circumference of a flat annular surface of the rotor before the first turbine stage along the outer surface of the section of the shaft or drum located between the compressor and the turbine. The cooling air flow passes through these bores into the cooling channels of the rotor, at the end of which it, reduced by the portion branched off for cooling the hottest rotor blades, exits into the propellant gas flow and with it into the diffuser.

Da, wie gesagt, die Zuströmung der Kühlluft zum Rotor im wesentlichen drallfrei, also ohne Umfangskomponente, im Drehrichtungssinn der Trommel erfolgt, wird sie auf ihrem Weg zum Rotor durch die Reibung an der Mantelfläche der Trommel in deren Umfangsrichtung beschleunigt, wenn auch im Verhältnis zur Umfangsgeschwindigkeit nicht sehr stark, so dass am Eintritt in die genannten Bohrungen und in die Rotorkühlkanäle diesen gegenüber noch eine grosse Geschwin­digkeitsdifferenz besteht. Sie muss also dort auf die Rotor­umfangsgeschwindigkeit beschleunigt werden. Die Trommel und der Rotor müssen also Pumparbeit leisten, die überdies die Kühllufttemperatur erhöht. Dies stellt also wie auch grösstenteils die Durchströmung der Kühlkanäle einen Verlust­faktor dar.Since, as said, the inflow of the cooling air to the rotor is essentially swirl-free, i.e. without a peripheral component, in the direction of rotation of the drum, it is accelerated on its way to the rotor by the friction on the circumferential surface of the drum in its circumferential direction, albeit in relation to The peripheral speed is not very strong, so that there is still a large difference in speed at the entry into the bores mentioned and into the rotor cooling channels. It must therefore be accelerated to the circumferential rotor speed. The drum and the rotor must therefore perform pumping work, which moreover increases the cooling air temperature. Like most of the flow through the cooling channels, this represents a loss factor.

Ein weiterer Verlust ist mit dem am Laufschaufelfuss der letzten Stufe austretenden Kühlluftstrom verbunden. Er tritt in die Treibgasströmung mit einer radial, tangential und axial gerichteten Geschwindigkeitskomponente ein und drängt sie radial ab, so dass die Nabengrenzschicht am Diffusor­eintritt eine für den Rückgewinn schädliche Verdickung er­leidet.Another loss is associated with the cooling air flow exiting the blade root of the last stage. It enters the propellant gas flow with a radially, tangentially and axially directed velocity component and forces it radially away, so that the hub boundary layer at the diffuser inlet suffers a thickening that is harmful to the recovery.

Um die Pumpverluste zu vermeiden, wird in der DE-A-34 24 139 der Anmelderin vorgeschlagen, der Rotorkühlluft nach ihrem Austritt aus dem Verdichter durch feststehende Drallgitter mit im wesentlichen radial gerichteten Schaufeln eine im Drehsinn des Rotors gerichtete Umfangsgeschwindigkeitskom­ponente zu verleihen, in der Grösse der Umfangsgeschwindigkeit der Rotorkühlkanäle, so dass die Kühlluft nicht erst auf diese beschleunigt werden muss. Die erwähnte Pumparbeit und die damit verbundenen Verluste fallen dadurch weg.In order to avoid the pumping losses, it is proposed in DE-A-34 24 139 by the applicant to give the rotor cooling air a circumferential speed component in the direction of rotation of the rotor after it emerges from the compressor by means of fixed swirl grids with essentially radially directed blades the peripheral speed of the rotor cooling channels, so that the cooling air does not have to be accelerated to them first. The pump work mentioned and the associated losses are thereby eliminated.

Neben der Kühlung der Beschaufelung und des Rotors im Bereich der Schaufelbefestigungsnuten ist es bei Rotoren, die aus einer Reihe am Umfang miteinander verschweisster Scheiben zusammengesetzt sind, nötig, auch die letzte Rotorscheibe separat zu kühlen, um die gewünschte Lebensdauer zu erhalten. Die Kühlluft dafür wird der ersten Zapfstelle des Verdichters, also bei tiefem Druck und tiefer Temperatur, entnommen und über das Lagerschild nach der letzten Rotorscheibe in das Rotorgehäuse eingeführt, von wo ihr Hauptteil radial auswärts strömt und durch einen schmalen, von der Umfangskante der letzten Rotorscheibe und dem daran anschliessenden Innenmantel des Abgasdiffusors begrenzten Ringspalt in den Diffusor eintritt, und zwar mit einer radial auswärts gerichteten Geschwindigkeitskomponente und, wegen der Reibung der Kühlluft an der Rotorscheibe, auch einer Umfangskomponente in der Rotordrehrichtung. Ein kleiner Teil der Kühlluft sperrt das Labyrinth der Wellendurchführung am Lagerschild.In addition to cooling the blades and the rotor in the area of the blade fastening grooves, it is the case with rotors a series of discs welded together on the circumference, it is necessary to cool the last rotor disc separately in order to achieve the desired service life. The cooling air for this is taken from the first tapping point of the compressor, i.e. at low pressure and low temperature, and introduced into the rotor housing via the end shield after the last rotor disk, from where its main part flows radially outwards and through a narrow one from the peripheral edge of the last rotor disk and the adjacent inner jacket of the exhaust gas diffuser enters the diffuser, with a radially outward speed component and, due to the friction of the cooling air on the rotor disk, also a peripheral component in the direction of rotation of the rotor. A small part of the cooling air blocks the labyrinth of the shaft bushing on the end shield.

Aufgabe der ErfindungObject of the invention

Die vorliegende Erfindung entstand aus der Aufgabe, durch eine zweckentsprechende Führung sowohl der Rotor- und Schaufel­kühlluft als auch der Rotorscheibenkühlluft diese in ihren Austrittsbereichen am Rotorende so in den Diffusor hinein zu lenken, dass ihre Geschwindigkeitsvektoren mit jenem des mittleren Abgasstromes an den genannten Bereichen be­züglich Betrag und Richtung im wesentlichen übereinstimmen. Darüber hinaus soll die Arbeitsfähigkeit der Rotorkühlluft weitgehend ausgenützt werden. Durch diese Führung soll auch der Rotormantel im Bereich der letzten Stufe bei gleicher Rotorkühlluftmenge stärker gekühlt werden als dies bei den bekannten Konstruktionen der Fall ist. Dadurch kann die Scheibenkühlluftmenge reduziert werden, was die Temperatur­differenzen innerhalb des Rotors und somit die Wärmespannungen verringert, um eine Verlängerung der Lebensdauer des Turbinen­rotors zu erzielen.The present invention arose from the task of guiding the rotor and blade cooling air as well as the rotor disk cooling air in their outlet areas at the rotor end into the diffuser in such a way that their velocity vectors correspond to that of the average exhaust gas flow in the areas mentioned in terms of amount and direction essentially coincide. In addition, the working capacity of the rotor cooling air should be largely used. This guide is also intended to cool the rotor jacket in the area of the last stage with the same amount of rotor cooling air than is the case with the known constructions. The disc cooling air quantity can thereby be reduced, which reduces the temperature differences within the rotor and thus the thermal stresses in order to achieve an extension of the service life of the turbine rotor.

Die erfindungsgemässe axial durchströmte Gasturbine ist dadurch gekennzeichnet, dass für die Kühlluftführung im Bereich der letzten Stufe Kanäle vorgesehen sind, die im Bereich des Leitschaufelkranzes der letzten Stufe im Rotor­mantel und im Bereich des Laufschaufelkranzes der letzten Stufe in dessen Schaufelfüssen verlaufen, wobei mindestens am Ende des letzten Laufschaufelkranzes ein Kühlluftschaufel­gitter in einem am Turbinenrotor befestigten Kühlschaufel­kranz vorhanden ist, dessen Kanäle so orientiert sind, dass die Geschwindigkeitsvektoren der in den Diffusor austretenden Kühlluft im wesentlichen mit dem mittleren Geschwindigkeitsvek­tor der Abgasströmung übereinstimmt, und wobei die Begren­zungen für das Abströmen der Kühlluft in den Diffusor so gestaltet sind, dass deren Ablösung vermieden und die Treib­gasströmung im Nabenbereich des letzten Laufschaufelkranzes homogenisiert wird.The axially flowed through gas turbine is characterized in that for the cooling air duct in the In the area of the last stage, channels are provided which run in the area of the guide vane ring of the last stage in the rotor casing and in the area of the rotor vane ring of the last stage in its blade roots, a cooling air vane grille being present in a cooling vane ring attached to the turbine rotor, at least at the end of the last rotor vane ring Channels are oriented so that the speed vectors of the cooling air exiting into the diffuser essentially coincide with the average speed vector of the exhaust gas flow, and the limits for the outflow of the cooling air into the diffuser are designed in such a way that their separation is avoided and the propellant gas flow in the hub area of the last blade ring is homogenized.

Der Erfindungsgegenstand wird im folgenden anhand einiger in der Zeichnung dargestellter Ausführungsbeispiele näher beschrieben.The subject matter of the invention is described in more detail below with reference to some exemplary embodiments shown in the drawing.

Kurze Beschreibung der FigurenBrief description of the figures

In der Zeichnung stellen dar:

  • Fig. 1 einen Längsschnitt durch eine Hälfte eines Gasturbinen­rotors mit schematischer Darstellung der Beschaufelung,
  • Fig. 2 und 3 Details aus Fig. 1,
  • Fig. 4 ein weiteres Ausführungsbeispiel,
  • Fig. 5 Details aus diesem Ausführungsbeispiel, und die
  • Fig. 6 eine dritte Variante der Erfindung.
In the drawing:
  • 1 shows a longitudinal section through half of a gas turbine rotor with a schematic representation of the blading,
  • 2 and 3 details of Fig. 1,
  • 4 shows a further exemplary embodiment,
  • Fig. 5 details from this embodiment, and
  • Fig. 6 shows a third variant of the invention.

Fig. 1 zeigt einen Teil eines Turbinenrotors 1, der sich aus geschmiedeten Rotorscheiben 2, 3, 4 zusammensetzt, die entlang an deren Stirnseiten angeschmiedeter Ringe miteinan­ der verschweisst sind. Die Schaufeln der Laufschaufelkränze 5 bis 9 sind auf bekannte Weise mit ihrem Fuss von Doppel­hammerkopfprofil in die entsprechend profilierten Schaufel­befestigungsnuten eingesetzt. Zwischen zwei benachbarten Laufschaufelkränzen sind in einem Leitschaufelträger 10 auf ähnliche Weise wie die Laufschaufeln im Rotor Leitschaufeln von Leitschaufelkränzen 11 bis 14 verankert. Da im vorliegenden Zusammenhang unerheblich, sind die Leitschaufelbefestigungen nur schematisch angedeutet.Fig. 1 shows a part of a turbine rotor 1, which is composed of forged rotor disks 2, 3, 4, which along with each other on the end faces forged rings who are welded. The blades of the rotor blade rings 5 to 9 are inserted in a known manner with their base of double hammer head profile into the correspondingly profiled blade fastening grooves. Between two adjacent rotor blade rings, guide vanes of guide blade rings 11 to 14 are anchored in a guide blade carrier 10 in a manner similar to the rotor blades in the rotor. As irrelevant in the present context, the guide vane attachments are only indicated schematically.

Für die Kühlung des Rotormantels, worunter die äusserte Zone des Rotors mit ihren Befestigungsnuten für die Lauf­schaufeln und Wärmestausegmente zu verstehen ist, sowie der durch die Treibgastemperatur höchstbeanspruchten Lauf­schaufeln wird der letzten Stufe des nicht dargestellten Verdichters - er befindet sich rechts des ersten Laufschaufel­kranzes 5 der Turbine - der erforderliche Kühlluftstrom entnommen, worauf ihm durch ein zwischen dem Verdichter und der ersten Turbinenstufe angeordnetes Drallschaufelgitter, das in der eingangs erwähnten DE-A-34 24 139 beschrieben ist, eine tangentiale Geschwindigkeitskomponente erteilt wird, die gleich ist der Umfangsgeschwindigkeit der Rotor­kühlkanäle. Die Kühlluft tritt also dann mit der Relativge­schwindigkeit Null in Umfangsrichtung gegenüber dem Turbinen­rotor im wesentlichen axial, wie durch den Geschwindigkeits­pfeil 16 angedeutet, durch eine Reihe von Kühlluftbohrungen 15 in das Kühlkanalsystem der Turbine ein. Ueber die Kühlluft­bohrungen 15, die in grosser Zahl über eine kreisringförmige, ebene Stirnfläche 17 vor dem ersten Laufschaufelkranz verteilt vorgesehen sind, gelangt die Kühlluft in eine Ringnut 18, die sich zu ihrem Umfang hin im Querschnitt keilförmig erwei­tert, und aus dieser durch eine Reihe von unterbrochenen Ringspalten 19 vor dem ersten Laufschaufelkranz 5 und zwischen jeweils zweien der folgenden Laufschaufelkränze sowie durch Kanäle 20 im Bereich der Schaufelfüsse schliesslich in Schau­felfusskanäle 21 des letzten Laufschaufelkranzes 9. Die Ringspalte 19 sind begrenzt durch die Umfangsflächen des Rotormantels und durch unsymmetrische Wärmestausegmente 22, 23, die sich zwischen je zwei Laufschaufelkränzen befinden und den Rotormantel und die Laufschaufelfüsse vor Ueberhitzung durch den Treibgasstrom schützen. Die dem Treibgasstrom ausgesetzte zylindrische Aussenfläche der längeren der beiden unsymmetrischen Wärmestausegmente bilden zusammen mit den beiden Dichtleisten an den Deckbändern der Leitschaufeln 11 - 14 Drosselstellen, um die Verluste im Gasstrom zu mini­mieren. Für die Laufschaufeln der letzten Stufe mit ihren nahezu axial gerichteten Sägezahnfüssen ist anstatt der vor und hinter den Schaufeln angeordneten Wärmestausegmente 22, 23 ein Kranz von symmetrischen Wärmestausegmenten 24 mit einer eigenen Befestigungsnut im Rotormantel für die Aufnahme ihrer Schaufelfüsse vorgesehen. Ihre Stege 25 sind dann mit irgendwelchen Durchbrüchen 26 für die Kühlluft zu versehen.For the cooling of the rotor shell, which is to be understood as the outer zone of the rotor with its fastening grooves for the rotor blades and heat accumulation segments, and for the rotor blades which are subjected to the greatest stress by the propellant gas temperature, the last stage of the compressor (not shown) is located to the right of the first rotor blade ring 5 of the turbine - The required cooling air flow is removed, whereupon it is given a tangential speed component, which is equal to the peripheral speed of the rotor cooling channels, by a swirl vane grille arranged between the compressor and the first turbine stage, which is described in the aforementioned DE-A-34 24 139. The cooling air then enters the cooling duct system of the turbine at a relative speed of zero in the circumferential direction substantially axially, as indicated by the speed arrow 16, through a series of cooling air bores 15. Via the cooling air bores 15, which are provided in large numbers distributed over an annular, flat end face 17 in front of the first rotor blade ring, the cooling air passes into an annular groove 18 which widens in cross-section to its circumference, and from this through a series of interrupted annular gaps 19 in front of the first rotor blade ring 5 and between two of the following rotor blade rings as well as through channels 20 in the area of the blade roots finally into blade root channels 21 of the last rotor blade ring 9. The annular gaps 19 are delimited by the circumferential surfaces of the Rotor jacket and by asymmetrical heat accumulation segments 22, 23, which are located between two rotor blade rings and protect the rotor jacket and the rotor blade feet from overheating by the propellant gas flow. The cylindrical outer surface of the longer of the two asymmetrical heat accumulation segments exposed to the propellant gas flow, together with the two sealing strips on the cover strips of the guide vanes 11-14, form throttling points in order to minimize the losses in the gas flow. For the blades of the last stage with their almost axially directed sawtooth roots, instead of the heat accumulation segments 22, 23 arranged in front of and behind the blades, a ring of symmetrical heat accumulation segments 24 with its own fastening groove in the rotor casing is provided for receiving their blade roots. Your webs 25 are then to be provided with any openings 26 for the cooling air.

Die Schaufelfusskanäle 20, 21 können zweckmässig aus zwei Nuten in den jeweils beiden, in Umfangsrichtung aneinander­stossenden Seitenflanken benachbarten Laufschaufeln gebildet sein, die zusammen geschlossene Kanäle ergeben. Bei den nahezu axial gerichteten Schaufelfüssen können diese Kanäle, wie bei den Schaufeln des letzten Laufschaufelkranzes 9, aber auch in den Schaufelnuten selbst vorgesehen sein.The blade root passages 20, 21 can expediently be formed from two grooves in the two blades, which adjoin each other in the circumferential direction and adjoin one another in the circumferential direction, and which result in closed passages. In the case of the almost axially directed blade roots, these channels can also be provided in the blade grooves themselves, as in the blades of the last rotor blade ring 9.

Bei Gasturbinen hoher Leistungsdichte werden im allgemeinen die Leit- und Laufschaufeln der am stärksten temperatur­belasteten Stufen, beispielsweise die ersten zwei, als Hohl­schaufeln mit Luftkühlung ausgeführt. Für die Laufschaufeln wird die Kühlluft an den Schaufelfüssen aus dem beschrie­benen Kühlluftstrom abgezweigt. Da für die Erfindung un­wesentlich, sind die Elemente der Schaufelkühlung in Fig. 1 nicht dargestellt.In the case of gas turbines with a high power density, the guide and rotor blades of the most temperature-loaded stages, for example the first two, are designed as hollow blades with air cooling. For the moving blades, the cooling air is branched off at the blade roots from the cooling air flow described. As not essential to the invention, the elements of the blade cooling are not shown in FIG. 1.

Aus den Schaufelfusskanälen 21 des letzten Laufschaufelkranzes 9 gelangt die Kühlluft in einen Kühlluftschaufelkranz 27, der am Rotorkörper befestigt ist und der knapp innerhalb seines Umfangs ein kegelstumpfförmiges Laufschaufelgitter 28 aufweist, das, über seinen Umfang gleichmässig verteilt, Kühlluftschaufeln 31 aufweist, denen ein Gleichrichterring 29 vorgeschaltet ist, der aus über den ganzen Durchströmquer­schnitt verteilten, wabenförmigen Kanälen 30 besteht.The cooling air passes from the blade root channels 21 of the last moving blade ring 9 into a cooling air blade ring 27, which is attached to the rotor body and which has a frusto-conical rotor blade grille 28 just inside its circumference, which, evenly distributed over its circumference, has cooling air blades 31 which are preceded by a rectifier ring 29 which consists of honeycomb-shaped channels 30 distributed over the entire flow cross-section.

Die Fig. 2 zeigt das eingekreiste Detail II von Fig. 1 in grösserem Massstab und die Fig. 3 die Abwicklung des in Fig. 2 eingetragenen Schnittverlaufes III-III in Form einer durch die Kanalmitte gelegten Kegelschale. Der Gleichrichter­ring 29 hat die Aufgabe, die aus den Schaufelfusskanälen 21 der letzten Laufschaufeln 9 austretenden Kühlluftstrahlen zu homogenisieren, um eine möglichst ablösungsfreie Strömung in den von den Schaufeln 31 begrenzten Kanälen zu erhalten.Fig. 2 shows the circled detail II of Fig. 1 on a larger scale and Fig. 3 shows the development of the section III-III shown in Fig. 2 in the form of a conical shell placed through the center of the channel. The rectifier ring 29 has the task of homogenizing the cooling air jets emerging from the blade root channels 21 of the last rotor blades 9 in order to obtain a flow in the channels delimited by the blades 31 that is as free as possible from separation.

Der Kühlluftschaufelkranz 27 erfüllt einen Teil der in der Einleitung gestellten Erfindungsaufgabe, indem er die Strom­fäden des Kühlluftstromes so umlenkt, dass sich ihre Ge­schwindigkeitsvektoren auf dem ganzen Umfang der Diffusornabe im wesentlichen mit dem mittleren Geschwindigkeitsvektor des Abgasstromes decken mit der eingangs beschriebenen ver­lustmindernden Wirkung, indem der energiearmen Grenzschicht an der Diffusornabe Energie zugeführt und ihre Ablösestelle stromabwärts verschoben wird. Gleichzeitig wird die Energie der Rotorkühlluft teilweise zur Arbeitsabgabe an den Rotor ausgenützt.The cooling air vane ring 27 fulfills part of the object of the invention presented in the introduction by deflecting the flow threads of the cooling air flow in such a way that their speed vectors over the entire circumference of the diffuser hub essentially coincide with the average speed vector of the exhaust gas flow with the loss-reducing effect described at the outset, by the Low-energy boundary layer is supplied with energy at the diffuser hub and its detachment point is shifted downstream. At the same time, the energy of the rotor cooling air is partially used to deliver work to the rotor.

Diese Wirkungen des Kühlluftstromes werden unterstützt durch die zweite erfindungsgemässe Massnahme, die darin besteht, dass auch die zur Kühlung der letzten Rotorscheibe 4 benutzte, aus dem Verdichter abgezweigte Kühlluft, wie die Schaufel­kühlluft, geführt in den Diffusor ausströmt. Die Scheiben­kühlluft tritt durch zwei in einem äusseren Turbinengehäuse­boden 32 vorgesehene Scheibenluftkanäle 33 in einen vom Boden 32 und einem inneren Turbinengehäuseboden 34 begrenz­ ten scheibenförmigen Hohlraum 35 ein, wird, wie durch die Geschwindigkeitspfeile angedeutet, in diesem radial nach innen gegen die Rotorachse umgelenkt und gelangt durch eine Reihe in Achsnähe vorgesehener innerer Scheibenluftkanäle 36 vor die Rotorscheibe 4, wo ihr Hauptteil nach oben ab­gelenkt und über einen Ringspalt 37 und einen Ringraum 38 durch einen Ringschlitz 39 in die Nabengrenzschicht aus­geblasen wird. Zu der erfindungsgemäss beabsichtigten Ein­strömung in die Nabengrenzschicht trägt neben der inneren Kontur des Kühlluftschaufelringes 27 auch der konvexgekrümmte Einlaufbereich 40 der Diffusornabe 41 bei, der die ausströ­mende Scheibenkühlluft zusammen mit der Reaktorkühlluft durch seine Krümmung ansaugt. Die kegelstumpfförmige Mantel­fläche 64 des Kühlluftschaufelkranzes 27 ist gegenüber der Rotorachse so geneigt ausgeführt und in der Länge so bemessen, dass der Abgasstrom hinter dem letzten Laufschaufelkranz 9 homogenisiert wird.These effects of the cooling air flow are supported by the second measure according to the invention, which consists in that the cooling air used for cooling the last rotor disk 4 and branched off from the compressor, such as the blade cooling air, flows out into the diffuser in a guided manner. The disk cooling air passes through two disk air channels 33 provided in an outer turbine housing base 32 into one bounded by the bottom 32 and an inner turbine housing base 34 th disk-shaped cavity 35, is, as indicated by the speed arrows, deflected radially inwards against the rotor axis in this and passes through a series of inner disk air channels 36 provided near the axle in front of the rotor disk 4, where its main part is deflected upward and via an annular gap 37 and an annular space 38 is blown out through an annular slot 39 into the hub boundary layer. In addition to the inner contour of the cooling air vane ring 27, the convexly curved inlet area 40 of the diffuser hub 41, which sucks in the outflowing disk cooling air together with the reactor cooling air through its curvature, also contributes to the intended inflow into the hub boundary layer. The frustoconical lateral surface 64 of the cooling air vane ring 27 is designed to be inclined with respect to the rotor axis and its length is such that the exhaust gas flow behind the last rotor vane ring 9 is homogenized.

Ein kleiner Teil der durch den Kanal 36 einströmenden Scheiben­kühlluft sperrt das Labyrinth 41 am Lagerschild.A small part of the disk cooling air flowing in through the channel 36 blocks the labyrinth 41 on the end shield.

Die Fig. 4 und 5 zeigen eine zweite Ausführungsform der Rotorkühlluftführung. Nach dem vorletzten Laufschaufelkranz 43 tritt die Kühlluft über einen rotorfesten Zwischenkanal 44 in ein Schaufelgitter 45 eines rotorfesten Schaufelgitter­kranzes 46 ein und aus diesem in ein Schaufelgitter 47 eines leitschaufelfesten Schaufelgitterkranzes 48, aus dem sie in Endkanäle 49 umgelenkt wird. Die Eintrittspartien der­selben bestehen aus der vorderen Hälfte 50 eines Schaufel­gitters, den Profilnasen, in einem rotorfesten Schaufel­gitterkranz 50′, und der Austrittsbereich aus der hinteren Hälfte 51 dieses Schaufelgitters im Kühlluftschaufelkranz 53. Die Endkanäle 49 sind in Fig. 5 parallel zur Rotorachse verlaufend dargestellt, doch wird man sie in der Regel schräg zur Rotorachse verlaufend vorsehen, z.B. unter einem Winkel von 5 - 7°. Die am Rotorende austretende Kühlluft tritt sodann, gemeinsam mit der noch notwendigen Scheibenkühlluft, über den Ringraum 52 am Rotorende über den Einlaufbereich 40 der Diffusornabe in den Abgasstrom ein.4 and 5 show a second embodiment of the rotor cooling air duct. After the penultimate rotor blade ring 43, the cooling air enters via a rotor-fixed intermediate channel 44 into a blade grille 45 of a rotor-fixed blade grille ring 46 and out of this into a blade grille 47 of a guide-blade fixed blade grille ring 48, from which it is deflected into end channels 49. The inlet parts of the latter consist of the front half 50 of a vane grille, the profile lugs, in a rotor-fixed vane grille ring 50 ', and the exit region from the rear half 51 of this vane grille in the cooling air vane ring 53. The end channels 49 are shown in FIG. 5 running parallel to the rotor axis, however, they will usually be provided at an angle to the rotor axis, for example at an angle from 5 - 7 °. The cooling air emerging at the rotor end then, together with the disk cooling air still required, enters the exhaust gas flow via the annular space 52 at the rotor end via the inlet region 40 of the diffuser hub.

Eine weitere Ausführungsform der Erfindung zeigt die Fig. 6. Nach dem vorletzten Laufschaufelkranz 43 wird die Kühlluft im wesentlichen bis zum Ende des Laufschaufelkranzes 9 axial geführt und erst dort durch einen Kühlluftschaufelkranz 63 mit der gewünschten Richtung in den Abgasstrom ausgeblasen. Nach dem vorletzten Laufschaufelkranz 43 durchläuft sie wiederum, wie bei der Ausführung nach Fig. 4, einen Zwischen­kanal 54 und ein Schaufelgitter 55 in einem rotorfesten Schaufelgitterkranz 56, ein Schaufelgitter 57 in einem leit­schaufelfesten Schaufelgitterkranz 58, sodann einen am letzten Laufschaufelkranz 9 befestigten rotorfesten Schaufelgitter­kranz 59, dessen Schaufelgitter 60 aus den vorderen Schaufel­hälften besteht, während die hinteren Schaufelhälften das Schaufelgitter 62 im Kühlluftschaufelkranz 63 bilden. Zwischen den beiden Schaufelgittern 60 und 61 erstrecken sich wie bei der Ausführung nach Fig. 4 die Endkanäle 61, und zwar vorzugsweise unter einem Winkel gegen eine Achsparallele geneigt.Another embodiment of the invention is shown in FIG. 6. After the penultimate rotor blade ring 43, the cooling air is guided axially essentially to the end of the rotor blade ring 9 and is only then blown out into the exhaust gas flow in the desired direction by a cooling air blade ring 63. After the penultimate rotor blade ring 43, it again passes, as in the embodiment according to FIG. 4, an intermediate channel 54 and a blade grille 55 in a rotor-fixed blade grille ring 56, a blade grille 57 in a rotor blade-fixed blade grille ring 58, then a rotor-fixed blade grille 59 fastened to the last rotor blade ring 9 whose vane grille 60 consists of the front vane halves, while the rear vane halves form the vane grill 62 in the cooling air vane ring 63. As in the embodiment according to FIG. 4, the end channels 61 extend between the two vane grids 60 and 61, preferably inclined at an angle to an axis parallel.

Claims (6)

1. Axial durchströmte Gasturbine, mit Kühleinrichtungen für den Turbinenrotor (1) und seine Laufschaufelkränze (5 - 9), wobei die Kühlluft aus dem Verdichter abgezweigt und auf bekannte Weise durch eine Dralleinrichtung in Umfangsrichtung so beschleunigt wird, dass sie gegenüber Kühlluftbohrungen (15) am Turbinenrotor (1), durch welche die Kühlluft in das Kühlluftführungssystem einströmt, in der Umfangsrichtung die Geschwindigkeit Null hat, dadurch gekennzeichnet, dass für die Kühlluftführung im Bereich der letzten Stufe (9 + 14) Kanäle (26, 21, 28; 44, 45, 47, 50, 49, 51, 52, 39; 54, 55, 57, 60, 61, 62) vorgesehen sind, die im Bereich des Leitschaufelkranzes (14) der letzten Stufe im Rotormantel und im Bereich des Laufschaufelkranzes (9) der letzten Stufe in dessen Schaufelfüssen verlaufen, wobei mindestens am Ende des letzten Laufschaufelkranzes (9) ein Kühlluftschaufelgitter (28; 51; 62) in einem am Turbinenrotor (1) befestigten Kühlluftschaufelkranz (27; 53; 63) vorhanden ist, dessen Kanäle so orientiert sind, dass die Geschwindigkeits­vektoren der in den Diffusor austretenden Kühlluft im wesentlichen mit dem mittleren Geschwindigkeitsvektor der Abgasströmung übereinstimmen, und wobei die Begren­zungen für das Abströmen der Kühlluft in den Diffusor so gestaltet sind, dass deren Ablösung vermieden und die Treibgasströmung im Nabenbereich des letzten Lauf­schaufelkranzes (9) homogenisiert wird.1. Axial gas turbine, with cooling devices for the turbine rotor (1) and its rotor blades (5 - 9), the cooling air being branched off from the compressor and accelerated in a known manner by a swirl device in the circumferential direction in such a way that it is opposite cooling air bores (15) on the turbine rotor (1), through which the cooling air flows into the cooling air guide system, has zero speed in the circumferential direction, characterized in that for the cooling air guide in the area of the last stage (9 + 14), channels (26, 21, 28; 44, 45, 47, 50, 49, 51, 52, 39; 54, 55, 57, 60, 61, 62) are provided, which are in the area of the guide vane ring (14) of the last stage in the rotor casing and in the area of the moving vane ring (9) the last stage in its blade roots, at least at the end of the last rotor blade ring (9) a cooling air vane grille (28; 51; 62) is present in a cooling air vane ring (27; 53; 63) attached to the turbine rotor (1) The channels are oriented in such a way that the speed vectors of the cooling air exiting into the diffuser essentially match the mean speed vector of the exhaust gas flow, and the limits for the outflow of the cooling air into the diffuser are designed in such a way that their separation is avoided and the propellant gas flow in the hub area of the last blade ring (9) is homogenized. 2. Gasturbine nach Anspruch 1, dadurch gekennzeichnet, dass der Kühlluftkanal im Bereich des letzten Leitschaufel­kranzes (14) von einer durch symmetrische Wärmestausegmente (24) abgedeckten Ringnut im Rotorkörper und von Durch­ brüchen (26) in den Stegen (25) dieser Wärmestausegmente (24) gebildet wird, dass für die Kühlluftführung im Bereich des letzten Laufschaufelkranzes (9) Schaufelfusskanäle (21) vorgesehen sind, und dass dem Kühlluftschaufelgitter (28) im Kühlluftschaufelkranz (27), in Strömungsrich­tung gesehen, ein Gleichrichterring (29) vorgesetzt ist.2. Gas turbine according to claim 1, characterized in that the cooling air duct in the region of the last guide vane ring (14) from an annular groove covered by symmetrical heat accumulation segments (24) in the rotor body and through breaks (26) in the webs (25) of these heat accumulation segments (24) is formed such that blade root ducts (21) are provided for the cooling air guidance in the area of the last moving blade ring (9), and that the cooling air blade grille (28) in the cooling air blade ring (27), Seen in the direction of flow, a rectifier ring (29) is placed in front. 3. Gasturbine nach Anspruch 1, dadurch gekennzeichnet, dass die Kühlluftführung im Bereich des letzten Leitschaufel­kranzes (14) aus Zwischenkanälen (54) im Rotormantel, einem rotorfesten Schaufelgitter (55) am Ende dieser Zwischenkanäle und einem Schaufelgitter (57) in einem leitschaufelfesten Schaufelgitterkranz (58) besteht, und dass die Kühlluftführung im Bereich des letzten Lauf­schaufelkranzes (9) ein Schaufelgitter (60) in einem rotorfesten Schaufelgitterkranz (59) aufweist, welches Schaufelgitter (60) aus den die Schaufelnasen bildenden vorderen Schaufelhälften besteht, ferner Endkanäle (61) in den Schaufelfüssen des letzten Laufschaufelkranzes (9) sowie einen rotorfesten Kühlluftschaufelkranz (63) mit einem Kühlluftschaufelgitter (62), welches aus den hinteren Schaufelhälften besteht.3. Gas turbine according to claim 1, characterized in that the cooling air duct in the area of the last guide vane ring (14) from intermediate channels (54) in the rotor shell, a rotor-fixed blade grille (55) at the end of these intermediate channels and a blade grille (57) in a guide-blade fixed blade grille ring ( 58), and that the cooling air duct in the area of the last rotor blade ring (9) has a blade grille (60) in a rotor-fixed blade grille ring (59), which blade grille (60) consists of the front blade halves forming the blade lugs, and also end channels (61) in the blade roots of the last rotor blade ring (9) and a rotor-fixed cooling air blade ring (63) with a cooling air blade grille (62), which consists of the rear blade halves. 4. Gasturbine nach Anspruch 1, dadurch gekennzeichnet, dass die Kühlluftführung im Bereich des letzten Leitschaufel­kranzes (14) rotorfeste Zwischenkanäle (44), einen rotor­festen Schaufelgitterkranz (46) mit einem zur Rotorachse hin gerichteten, gekrümmten Schaufelgitter (45) sowie ein zur Rotorachse hin gerichtetes Schaufelgitter (47) in einem leitschaufelfesten Schaufelgitterkranz (48) auf­weist, und dass die Kühlluftführung im Bereich des letzten Laufschaufelkranzes (9) ein Schaufelgitter (50) in einem rotorfesten Schaufelgitterkranz (50′) aufweist, welches Schaufelgitter (50) aus den die Schaufelnasen bildenden vorderen Schaufelhälften besteht, ferner Endkanäle (49) im Bereich der Schaufelfüsse des letzten Laufschaufel­kranzes (9) und einen rotorfesten Kühlluftschaufelkranz (53) mit einem Kühlluftschaufelgitter (51), welches aus den hinteren Schaufelhälften besteht, ferner gekennzeich­net durch einen Ringraum (52) und einen Ringschlitz (39) zwischen dem Kühlluftschaufelkranz (53) und der Diffusor­nabe (42).4. Gas turbine according to claim 1, characterized in that the cooling air duct in the area of the last guide vane ring (14) has rotor-fixed intermediate channels (44), a rotor-fixed vane grille ring (46) with a curved vane grille (45) directed towards the rotor axis and one towards the rotor axis Directional vane grille (47) in a vane grille ring (48) fixed to the guide vanes, and that the cooling air duct in the area of the last rotor vane ring (9) has a vane grille (50) in a rotor-fixed vane grille ring (50 '), which has vane grilles (50) from which the blade lugs forming front blade halves, also end channels (49) in the area of the blade roots of the last rotating blade ring (9) and a rotor-fixed cooling air blade ring (53) with a cooling air blade grille (51), which consists of the rear blade halves, further characterized by an annular space (52) and an annular slot (39) between the cooling air blade ring (53 ) and the diffuser hub (42). 5. Gasturbine nach einem der Ansprüche 2, 3 oder 4, dadurch gekennzeichnet, dass der Einlaufbereich (40) der Diffusor­nabe (42) im Axialschnitt stromlinienförmig profiliert ist.5. Gas turbine according to one of claims 2, 3 or 4, characterized in that the inlet region (40) of the diffuser hub (42) is streamlined in axial section. 6. Gasturbine nach Anspruch 1, dadurch gekennzeichnet, dass die kegelstumpfförmige Mantelfläche (64) des Kühlluft­schaufelkranzes (27; 53; 63) gegenüber der Rotorachse so geneigt ausgeführt und so bemessen ist, dass der Abgas­strom hinter dem letzten Laufschaufelkranz (9) homogeni­siert wird.6. Gas turbine according to claim 1, characterized in that the truncated cone-shaped outer surface (64) of the cooling air vane ring (27; 53; 63) is designed to be inclined with respect to the rotor axis and is dimensioned such that the exhaust gas flow behind the last rotor blade ring (9) is homogenized.
EP88115694A 1987-10-30 1988-09-23 Axial gas turbine Expired - Lifetime EP0313826B1 (en)

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DE19873736836 DE3736836A1 (en) 1987-10-30 1987-10-30 AXIAL FLOWED GAS TURBINE
DE3736836 1987-10-30

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EP0313826A1 true EP0313826A1 (en) 1989-05-03
EP0313826B1 EP0313826B1 (en) 1992-09-02

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JP (1) JP2656576B2 (en)
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DE (2) DE3736836A1 (en)

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EP0636764A1 (en) * 1993-07-17 1995-02-01 ABB Management AG Gasturbine with cooled rotor
WO1999047798A1 (en) * 1998-03-16 1999-09-23 Siemens Westinghouse Power Corporation Pressure changing mechanism for turbine cooling air
US8277170B2 (en) 2008-05-16 2012-10-02 General Electric Company Cooling circuit for use in turbine bucket cooling
EP2520764A1 (en) * 2011-05-02 2012-11-07 MTU Aero Engines GmbH Blade with cooled root
EP2551453A1 (en) * 2011-07-26 2013-01-30 Alstom Technology Ltd Cooling device of a gas turbine compressor
EP3106613A1 (en) * 2015-06-06 2016-12-21 United Technologies Corporation Cooling system for gas turbine engines
FR3054855A1 (en) * 2016-08-08 2018-02-09 Safran Aircraft Engines ROTOR DISC OF TURBOMACHINE
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DE19852604A1 (en) * 1998-11-14 2000-05-18 Abb Research Ltd Rotor for gas turbine, with first cooling air diverting device having several radial borings running inwards through first rotor disk
DE19854907A1 (en) * 1998-11-27 2000-05-31 Rolls Royce Deutschland Cooling air conduction for high pressure axial aviation gas turbines with air flow guided through radial turbine, turbine plate, through ring gap, towards hub cob for cooling
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US6402471B1 (en) * 2000-11-03 2002-06-11 General Electric Company Turbine blade for gas turbine engine and method of cooling same
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GB0503676D0 (en) * 2005-02-23 2005-03-30 Rolls Royce Plc A lock plate arrangement
US8591184B2 (en) * 2010-08-20 2013-11-26 General Electric Company Hub flowpath contour
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US9080449B2 (en) * 2011-08-16 2015-07-14 United Technologies Corporation Gas turbine engine seal assembly having flow-through tube
CH705840A1 (en) 2011-12-06 2013-06-14 Alstom Technology Ltd High-pressure compressor, in particular in a gas turbine.
EP2725191B1 (en) 2012-10-23 2016-03-16 Alstom Technology Ltd Gas turbine and turbine blade for such a gas turbine
EP2837769B1 (en) * 2013-08-13 2016-06-29 Alstom Technology Ltd Rotor shaft for a turbomachine
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DE102022200592A1 (en) 2022-01-20 2023-07-20 Siemens Energy Global GmbH & Co. KG turbine blade and rotor
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EP0447886A1 (en) * 1990-03-23 1991-09-25 Asea Brown Boveri Ag Axial flow gas turbine
US5189874A (en) * 1990-03-23 1993-03-02 Asea Brown Boveri Ltd. Axial-flow gas turbine cooling arrangement
EP0636764A1 (en) * 1993-07-17 1995-02-01 ABB Management AG Gasturbine with cooled rotor
US6217280B1 (en) 1995-10-07 2001-04-17 Siemens Westinghouse Power Corporation Turbine inter-disk cavity cooling air compressor
WO1999047798A1 (en) * 1998-03-16 1999-09-23 Siemens Westinghouse Power Corporation Pressure changing mechanism for turbine cooling air
US8277170B2 (en) 2008-05-16 2012-10-02 General Electric Company Cooling circuit for use in turbine bucket cooling
EP2520764A1 (en) * 2011-05-02 2012-11-07 MTU Aero Engines GmbH Blade with cooled root
US9739151B2 (en) 2011-05-02 2017-08-22 Mtu Aero Engines Gmbh Blade, integrally bladed rotor base body and turbomachine
US9382802B2 (en) 2011-07-26 2016-07-05 General Electric Technology Gmbh Compressor rotor
EP2551453A1 (en) * 2011-07-26 2013-01-30 Alstom Technology Ltd Cooling device of a gas turbine compressor
US10001061B2 (en) 2014-06-06 2018-06-19 United Technologies Corporation Cooling system for gas turbine engines
EP3106613A1 (en) * 2015-06-06 2016-12-21 United Technologies Corporation Cooling system for gas turbine engines
FR3054855A1 (en) * 2016-08-08 2018-02-09 Safran Aircraft Engines ROTOR DISC OF TURBOMACHINE
WO2018029408A1 (en) * 2016-08-08 2018-02-15 Safran Aircraft Engines Turbo engine rotor disc
GB2567103A (en) * 2016-08-08 2019-04-03 Safran Aircraft Engines Turbo engine rotor disc
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Also Published As

Publication number Publication date
CA1310273C (en) 1992-11-17
JPH01151725A (en) 1989-06-14
JP2656576B2 (en) 1997-09-24
US4910958A (en) 1990-03-27
EP0313826B1 (en) 1992-09-02
DE3874283D1 (en) 1992-10-08
DE3736836A1 (en) 1989-05-11

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