EP1270874B1 - Gas turbine with an air compressor - Google Patents

Gas turbine with an air compressor Download PDF

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Publication number
EP1270874B1
EP1270874B1 EP01114599A EP01114599A EP1270874B1 EP 1270874 B1 EP1270874 B1 EP 1270874B1 EP 01114599 A EP01114599 A EP 01114599A EP 01114599 A EP01114599 A EP 01114599A EP 1270874 B1 EP1270874 B1 EP 1270874B1
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EP
European Patent Office
Prior art keywords
combustion chambers
gas turbine
section
cross
duct
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
EP01114599A
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German (de)
French (fr)
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EP1270874A1 (en
Inventor
Robert Bland
Charles Ellis
Peter Tiemann
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Siemens AG
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Siemens AG
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Publication date
Application filed by Siemens AG filed Critical Siemens AG
Priority to DE50107283T priority Critical patent/DE50107283D1/en
Priority to EP01114599A priority patent/EP1270874B1/en
Priority to JP2002172518A priority patent/JP2003042451A/en
Priority to US10/172,016 priority patent/US6672070B2/en
Priority to CNB021233160A priority patent/CN1328492C/en
Publication of EP1270874A1 publication Critical patent/EP1270874A1/en
Application granted granted Critical
Publication of EP1270874B1 publication Critical patent/EP1270874B1/en
Anticipated expiration legal-status Critical
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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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • 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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/184—Two-dimensional patterned sinusoidal

Definitions

  • the invention relates to a gas turbine with a compressor for air flowing in a variety of fluidic parallel switched combustion chambers is heated before going over a transfer channel flows to a gas passage in a turbine.
  • FIGS to 7 The subject of said US-PS 4,719,748 is shown in FIGS to 7 and the associated description, however, also one consuming cooling device in which a combustion chamber and a from this leading to a turbine connecting channel a second wall is covered against the flow of compressed air are.
  • this second wall are a variety of Provided openings through which the compressed air targeted is directed to the wall sections to be cooled.
  • the invention is based on the object, for a gas turbine of the type mentioned above to provide an arrangement in the one unavoidable pressure loss in the stream of compressed Air is further reduced.
  • compressed Air leaves it through a ring of vanes 4 and flows in the direction of the arrows 5 first by an annular in cross section paraxial Channel section 6 of an air duct, the inside through a Wall 38 and outwardly bounded by a wall 39.
  • the webs 7 carry an annular cross-section C-shaped deflector 8 and are via webs 7 in the end anchored the channel section 6.
  • One in the end of the canal section 6 lying leg 9 of the cross section of the deflector 8 forms with its upstream edge 9 one to one to the axis 1 concentric circle curved wavy line 37.
  • the wall thickness of the deflector 8 increases, starting from the edge 9 bis to its center, strong too and is also in the circumferential direction of the Deflector 8 not constant, but waving and decreasing.
  • a radially outside Cross-section legs of the deflector 8 is substantially adapted to the contour of the combustion chambers and forms with its free end a wavy edge 35th This Outside cross-sectional leg of the deflector 8 is beyond also wavy in shape, with the thus formed Waves opposite the waves of the wavy line 37 in opposite directions are, as best seen in Fig. 3 can be seen.
  • the special shape of the deflector 8 with in its circumferential direction Waves 35 and 37 forming legs of his C-shaped Cross section forces a division in its area of the air flow in a partial flow 5a to the top of the Combustion chambers 10 and in a partial flow 5b to the bottom of the Combustion chambers 10. This is the top of the combustion chambers 10, relative to the gas turbine, radially outward and correspondingly the bottom radially inward.
  • the routes of the partial flows 5a and 5b are about the same size, so that all parts of the cooling air from the compressor 3 to the entry into the combustion chambers 10th to travel the same distance.
  • Each of the combustion chambers 10 is supported by webs 11 from the inside on an outer casing 12, which at the same time the outer wall an air passage 20 and a continuation of the air duct 6 represents the air flowing in the direction of the arrows 5.
  • the envelope 12 carries at its outer free end a cap 13 through which the air enters the interior of the combustion chamber 10 is guided.
  • the combustion chambers 10 are so close together in the circumferential direction arranged that the outer sheaths 12 at their end facing the runner 1 penetrate each other would.
  • To the combustion chambers 10 together with their outer sheaths 12 still as far as desired in the direction of the To be able to push runners 1 are at the outer claddings 12 recesses 40 (Fig. 4) provided in the area adjacent Combustion chambers 10 between them a common Air duct 20 have.
  • the interior of the combustion chambers 10 is also by a not shown nozzle fuel, such as a combustible Supplied with gas or atomised liquid fuel, by the combustion of the air in the combustion chamber 10 to a hot gas 34 is baked out.
  • nozzle fuel such as a combustible Supplied with gas or atomised liquid fuel
  • the combustion chamber 10 and the outer casing 12 holding it are mounted in a socket 14 in a housing shell 15 and one fixed to the outer sheath 12 Fixed flange 16 on the outer end of the nozzle 14.
  • One inner end 36 of the combustion chamber 10 is sealed in one
  • Kunleitkanal 17 the output of the combustion chamber 10 with a in cross-section annular gas channel 18 in one Turbine connects.
  • hot gas 34 is a Variety of, for example, 10 to 30 combustion chambers 10 evenly distributed on the circumference of the turbine plant and are their Junctions in the transfer channel 17 by a in the direction on the gas channel 18 open peripheral channel 19 with each other connected.
  • the transfer channel 17 is connected by thin webs 21 anchored a guide member 22 of the turbine.
  • the guide member 22 and a guide member 23 are in a housing shell 24 stored and are by securing blocks 25 against Rotation secured.
  • the guiding parts 22 and 23 for example, by hydraulic or pneumatic Motors 26 axially parallel over small distances, wherein a flange 27 is elastically deformed and with in it stored deformation energy for the return of the Leiters 22 or 23 is used.
  • One of the housing shells 15 and 24 enclosed volume is divided by partitions 28 into chambers.
  • the guide parts 22 and 23 have a funnel-like shape and carry on their inside in guide rings 29 fixed vanes 30, whose guide rings 29 opposite ends are firmly connected together by rings 31. Between each other adjacent wreaths of vanes 30 is ever one Wreath provided on the rotor 1 wedged blades 32, whose free tips guide rings 33 are opposite. there the guide rings 29 and 33 form an outer and the rings 31, together with feet of the blades 32, an inner boundary the gas channel 18 for the hot gas 34 in the turbine.

Description

Die Erfindung betrifft eine Gasturbine mit einem Verdichter für Luft, die in einer Vielzahl von strömungstechnisch parallel geschalteten Brennkammern aufgeheizt ist, bevor sie über einen Überleitkanal einem Gaskanal in einer Turbine zuströmt.The invention relates to a gas turbine with a compressor for air flowing in a variety of fluidic parallel switched combustion chambers is heated before going over a transfer channel flows to a gas passage in a turbine.

In Gasturbinen wird üblicherweise zur Erzielung einer wirtschaftlichen Leistungsdichte angesogene Luft zunächst verdichtet und dann in Brennkammern aufgeheizt. Das dabei erzeugte Heißgas treibt dann eine Turbine.In gas turbines is usually used to achieve an economic Power density absorbed air first compressed and then heated in combustion chambers. The generated Hot gas then drives a turbine.

Zur Erzielung eines guten Gesamtwirkungsgrads ist es u. a. erforderlich, Strömungsverluste bei der Führung der verdichteten Luft gering zu halten. Gleichzeitig sollen jedoch mit der verdichteten, noch nicht aufgeheizten Luft verschiedene Bauteile der Turbinenanlage gekühlt werden. So muß zur Vermeidung von Schäden beispielsweise ein Überleit- oder Verbindungskanal, durch den aus den Brennkammern abfließendes Heißgas zur Turbine strömt, vor Überhitzung geschützt werden.To achieve a good overall efficiency, it is u. a. required, flow losses when guiding the compacted To keep air low. At the same time, however, with the compressed, not yet heated air different Components of the turbine system to be cooled. So must to avoid damage, for example, a transfer or connection channel, through the hot gas flowing out of the combustion chambers flows to the turbine, be protected from overheating.

In der US-PS 4,719,748 ist in Fig. 1 eine hierzu verbreitet eingesetzte Anordnung angegeben. Bei dieser Anordnung liegt ein langer Verbindungskanal zwischen einer Brennkammer und einem Turbineneinlass unmittelbar in einem Luftkanal, durch den komprimierte Luft zu einem Brenner fließt. Bei dieser Anordnung ist kein Diffusor zur Luftlenkung dargestellt und die Strömungsgeschwindigkeit der Luft ist bei Erreichen des Verbindungskanals weit abgesunken. Eine korrekte Kühlung ist demzufolge allenfalls bei relativ niedrigen Temperaturen des Heißgases möglich, weil höhere Temperaturen eine spezifische Strömungsgeschwindigkeit sowohl für die komprimierte Luft als auch für das Heißgas und eine spezifische Kanalhöhe und - ausrichtung für den Luftkanal erfordern. Eine ausreichende Kühlung ist soweit erkennbar bei dieser Lösung weder für die Ober- noch für die Unterseite des Verbindungskanals erreichbar, weil einerseits das Volumen des Luftkanals in diesem Bereich sehr groß ist und weil außerdem sowohl die Länge des zu kühlenden Kanalabschnitts als auch die von der komprimierten Luft nach dem Austritt aus einem Verdichter zurückzulegende Strecke verhältnismäßig lang ist.In US-PS 4,719,748 a is spread in Fig. 1 for this purpose used arrangement indicated. In this arrangement is a long connecting channel between a combustion chamber and a turbine inlet directly in an air duct, through the compressed air flows to a burner. In this arrangement No diffuser is shown for air control and the Flow velocity of the air is when reaching the connection channel dropped far. A correct cooling is Consequently, if necessary at relatively low temperatures of the Hot gas possible because higher temperatures have a specific Flow velocity for both the compressed air and also for the hot gas and a specific channel height and - require alignment for the air duct. A sufficient Cooling is as far as can be seen in this solution neither for Still accessible to the underside of the connection channel, because on the one hand the volume of the air duct in this area is very large and because, in addition, both the length of the cooling channel section as well as those of the compressed Air after leaving a compressor zurückzulegende Range is relatively long.

Gegenstand der genannten US-PS 4,719,748 ist in den Fig. 2 bis 7 und der zugehörigen Beschreibung jedoch außerdem eine aufwendige Kühleinrichtung, bei der eine Brennkammer und ein von dieser zu einer Turbine führender Verbindungskanal durch eine zweite Wand gegen den Strom der komprimierten Luft abgedeckt sind. In dieser zweiten Wand sind eine Vielzahl von Öffnungen vorgesehen, durch die die komprimierte Luft gezielt auf die zu kühlenden Wandabschnitte gelenkt ist. Durch die für diese Lösung angegebenen Variationen für die Anzahl, die Größe und die Form dieser Öffnungen ist zwar eine gute Kühlung erreichbar, nachteilig ist aber ein hierbei nicht unerheblicher, unvermeidbarer Druckverlust der komprimierten Luft, weil diese wiederholt verzögert und wieder beschleunigt werden muß.The subject of said US-PS 4,719,748 is shown in FIGS to 7 and the associated description, however, also one consuming cooling device in which a combustion chamber and a from this leading to a turbine connecting channel a second wall is covered against the flow of compressed air are. In this second wall are a variety of Provided openings through which the compressed air targeted is directed to the wall sections to be cooled. By the Variations for the number specified for this solution Although the size and shape of these openings is a good cooling achievable, but a disadvantage is not insignificant, unavoidable pressure loss of the compressed Air, because this repeatedly delays and accelerates again must become.

Eine andere Kühlmöglichkeit mit flüssigem Kühlmittel, die auch recht aufwendig ist, wird in der US-A-4,195,474 offenbart. Gegenstand dieser Druckschrift ist ein flüssigkeitsgekühlter Überleitkanal in einer Gasturbinenanlage. Der Überleitkanal, der Heißgase vom Brenner zum Turbineneingang leitet, besteht aus einem hohlen Bauteil, welches bedingt durch die Form des Brenners ein kreisförmiges Ende aufweist. Weiterhin weist das Bauteil ein rechteckiges Ende auf, welches im eingebauten Zustand dem Turbineneingang zugewandt ist. Das Bauteil weist eine Vielzahl von Kühlröhrchen auf, die auf der Innenseite des Hohlkörpers durch Löten angebracht sind. Diese Kühlröhrchen münden in ein gemeinsames Einlassrohr, das sich entlang der äußeren radialen Mittelachse des Bauteils erstreckt. Das Einlassrohr weist eine Einlassöffnung auf, durch die Kühlflüssigkeit, z.B. Wasser, in das Bauteil zu Kühlungszwecken eingeleitet wird. Entlang der inneren radialen Mittelachse des Körpers erstreckt sich ein Sammelrohr, das Öffnungen aufweist, durch die die Kühlflüssigkeit nach der Kühlung außerhalb des Bauteils abgeleitet wird.Another cooling option with liquid coolant, the is also quite expensive, is described in US-A-4,195,474 disclosed. The subject of this document is a liquid-cooled Transfer channel in a gas turbine plant. Of the Transfer duct, the hot gases from the burner to the turbine inlet leads, consists of a hollow component, which conditionally has a circular end by the shape of the burner. Furthermore, the component has a rectangular end, which when installed, facing the turbine inlet is. The component has a plurality of cooling tubes, which is attached to the inside of the hollow body by soldering are. These cooling tubes open into a common inlet pipe, extending along the outer radial center axis of the Component extends. The inlet pipe has an inlet opening on, by the cooling liquid, e.g. Water, in the component is initiated for cooling purposes. Along the inner Radial center axis of the body extends Collection tube having openings through which the Coolant after cooling outside the component is derived.

Der Erfindung liegt nun die Aufgabe zugrunde, für eine Gasturbine der eingangs genannten Art eine Anordnung zu schaffen, in der ein unvermeidbarer Druckverlust im Strom der komprimierten Luft weiter verringert ist.The invention is based on the object, for a gas turbine of the type mentioned above to provide an arrangement in the one unavoidable pressure loss in the stream of compressed Air is further reduced.

Diese Aufgabe ist erfindungsgemäß gelöst, in dem die verdichtete Luft in einem Luftkanal vom Austritt aus dem Verdichter bis zum Eintritt in die Brennkammern auf der gesamten Strecke mit annähernd konstanter Geschwindigkeit strömt. Zweckmäßigerweise ist dabei der Überleitkanal kürzer, als eine der Brennkammern im Durchmesser misst. Diese Lösung ist überraschend vorteilhaft, weil nicht nur der Druckabfall im Luftkanal, sondern darüber hinaus auch ein Druckabfall im Überleitkanal auf einen sehr geringen Wert abgesenkt sind. Dabei wird eine konstante Geschwindigkeit der Luft im Luftkanal dadurch erreicht, daß der wirksame Querschnitt des Luftkanals über die gesamte Strecke vom Austritt aus dem Verdichter bis zum Eintritt in die Brennkammern nahezu konstant ist.This object is achieved according to the invention, in which the compacted Air in an air duct exiting the compressor until entering the combustion chambers along the entire route flows at approximately constant speed. Conveniently, is the transfer channel shorter than one of Measuring combustion chambers in diameter. This solution is surprising advantageous because not only the pressure drop in the air duct, but also a pressure drop in the transfer channel lowered to a very low value. It will a constant velocity of the air in the air duct thereby achieved that the effective cross section of the air duct over the entire distance from the exit from the compressor to the Entry into the combustion chambers is almost constant.

Weitere zweckmäßige und/oder vorteilhafte Ausgestaltungen der Erfindung sind in den Ansprüchen 3 bis 14 angegeben.Further expedient and / or advantageous embodiments of Invention are given in claims 3 to 14.

Ein Ausführungsbeispiel der Erfindung ist anhand einer Zeichnung näher erläutert. Dabei zeigen

Fig. 1
einen Ausschnitt einer Gasturbine im Längsschnitt,
Fig. 2
einen Schnitt entlang der Linie II - II in Fig. 1,
Fig. 3
einen Schnitt entlang der Linie III - III in Fig. 1 und
Fig. 4
eine Ansicht in Richtung IV aus Fig. II auf eine dort nicht dargestellte äußere Umhüllung einer Brennkammer.
An embodiment of the invention is explained in more detail with reference to a drawing. Show
Fig. 1
a section of a gas turbine in longitudinal section,
Fig. 2
a section along the line II - II in Fig. 1,
Fig. 3
a section along the line III - III in Fig. 1 and
Fig. 4
a view in the direction IV of FIG. II on an outer envelope of a combustion chamber, not shown there.

Ein ausschnittsweise dargestellter Läufer 1 einer Gasturbinenanlage rotiert um eine Achse 2. In einem Verdichter 3 komprimierte Luft verlässt diesen durch einen Kranz aus Leitschaufeln 4 und strömt in Richtung der Pfeile 5 zunächst durch einen im Querschnitt kreisringförmigen achsparallelen Kanalabschnitt 6 eines Luftkanals, der nach innen durch eine Wand 38 und nach außen durch eine Wand 39 begrenzt ist.A detail of a runner 1 a gas turbine plant rotates about an axis 2. In a compressor 3 compressed Air leaves it through a ring of vanes 4 and flows in the direction of the arrows 5 first by an annular in cross section paraxial Channel section 6 of an air duct, the inside through a Wall 38 and outwardly bounded by a wall 39.

Am Ende dieses Kanalabschnitts 6 passiert die verdichtete Luft Stege 7. Die Stege 7 tragen einen ringförmigen im Querschnitt C-förmigen Umlenker 8 und sind über Stege 7 im Ende des Kanalabschnitts 6 verankert. Ein im Ende des Kanalabschnitts 6 liegender Schenkel 9 des Querschnitts vom Umlenker 8 bildet mit seinem stromauf gewandten Rand 9 eine um einen zur Achse 1 konzentrischen Kreis geschwungene Wellenlinie 37. Die Wanddicke des Umlenkers 8 nimmt, ausgehend vom Rand 9 bis zu seiner Mitte, stark zu und ist auch in Umfangsrichtung des Umlenkers 8 nicht konstant, sondern wellenartig zu- und abnehmend. At the end of this channel section 6 passes the compressed Air webs 7. The webs 7 carry an annular cross-section C-shaped deflector 8 and are via webs 7 in the end anchored the channel section 6. One in the end of the canal section 6 lying leg 9 of the cross section of the deflector 8 forms with its upstream edge 9 one to one to the axis 1 concentric circle curved wavy line 37. The wall thickness of the deflector 8 increases, starting from the edge 9 bis to its center, strong too and is also in the circumferential direction of the Deflector 8 not constant, but waving and decreasing.

Radial über dem Umlenker 8 sind Brennkammern 10 zur Aufheizung der verdichteten Luft angeordnet. Ein radial außen liegender Querschnittsschenkel des Umlenkers 8 ist im wesentlichen an die Kontur der Brennkammern angepasst und bildet mit seinem freien Ende einen wellenförmigen Rand 35. Dieser äußere Querschnittsschenkel des Umlenkers 8 ist darüber hinaus auch in sich wellenförmig gestaltet, wobei die so gebildeten Wellen gegenüber den Wellen der Wellenlinie 37 gegenhäufig sind, wie besonders gut aus Fig. 3 ersichtlich ist.Radially above the deflector 8 are combustion chambers 10 for heating arranged the compressed air. A radially outside Cross-section legs of the deflector 8 is substantially adapted to the contour of the combustion chambers and forms with its free end a wavy edge 35th This Outside cross-sectional leg of the deflector 8 is beyond also wavy in shape, with the thus formed Waves opposite the waves of the wavy line 37 in opposite directions are, as best seen in Fig. 3 can be seen.

Die besondere Form des Umlenkers 8 mit in seiner Umfangsrichtung Wellen 35 bzw. 37 bildenden Schenkeln seines C-förmigen Querschnitts erzwingt in seinem Bereich eine Aufteilung des Luftstromes in einen Teilstrom 5a zur Oberseite der Brennkammern 10 und in einen Teilstrom 5b zur Unterseite der Brennkammern 10. Dabei liegt die Oberseite der Brennkammern 10, bezogen auf die Gasturbine, radial außen und entsprechend die Unterseite radial innen. Die Wegstrecken der Teilströme 5a und 5b sind etwa gleich groß, so daß alle Teile der Kühlluft vom Verdichter 3 bis zum Eintritt in die Brennkammern 10 gleich lange Wege zurückzulegen haben.The special shape of the deflector 8 with in its circumferential direction Waves 35 and 37 forming legs of his C-shaped Cross section forces a division in its area of the air flow in a partial flow 5a to the top of the Combustion chambers 10 and in a partial flow 5b to the bottom of the Combustion chambers 10. This is the top of the combustion chambers 10, relative to the gas turbine, radially outward and correspondingly the bottom radially inward. The routes of the partial flows 5a and 5b are about the same size, so that all parts of the cooling air from the compressor 3 to the entry into the combustion chambers 10th to travel the same distance.

Jede der Brennkammern 10 stützt sich über Stege 11 von innen an einer äußeren Umhüllung 12 ab, die gleichzeitig die Außenwand eines Luftkanals 20 ist und eine Fortsetzung des Luftkanals 6 für die in Richtung der Pfeile 5 strömende Luft darstellt. Die Umhüllung 12 trägt an ihrem äußeren freien Ende eine Kappe 13, durch die die Luft in den Innenraum der Brennkammer 10 geführt ist.Each of the combustion chambers 10 is supported by webs 11 from the inside on an outer casing 12, which at the same time the outer wall an air passage 20 and a continuation of the air duct 6 represents the air flowing in the direction of the arrows 5. The envelope 12 carries at its outer free end a cap 13 through which the air enters the interior of the combustion chamber 10 is guided.

Die Brennkammern 10 sind in Umfangsrichtung so dicht nebeneinander angeordnet, daß sich die äußeren Umhüllungen 12 an ihrem dem Läufer 1 zugekehrten Ende gegenseitig durchdringen müssten. Um die Brennkammern 10 mitsamt ihren äußeren Umhüllungen 12 trotzdem soweit wie gewünscht in Richtung auf den Läufer 1 schieben zu können, sind an den äußeren Umhüllungen 12 Ausnehmungen 40 (Fig. 4) vorgesehen, in deren Bereich benachbarte Brennkammern 10 zwischen sich einen gemeinsamen Luftkanal 20 aufweisen.The combustion chambers 10 are so close together in the circumferential direction arranged that the outer sheaths 12 at their end facing the runner 1 penetrate each other would. To the combustion chambers 10 together with their outer sheaths 12 still as far as desired in the direction of the To be able to push runners 1 are at the outer claddings 12 recesses 40 (Fig. 4) provided in the area adjacent Combustion chambers 10 between them a common Air duct 20 have.

Dem Innenraum der Brennkammern 10 ist außerdem durch eine nicht dargestellte Düse Brennstoff, beispielsweise ein brennbares Gas oder zerstäubter, flüssiger Brennstoff zugeführt, durch dessen Verbrennung die Luft in der Brennkammer 10 zu einem Heißgas 34 ausgeheizt wird.The interior of the combustion chambers 10 is also by a not shown nozzle fuel, such as a combustible Supplied with gas or atomised liquid fuel, by the combustion of the air in the combustion chamber 10 to a hot gas 34 is baked out.

Die Brennkammer 10 und die sie haltende äußere Umhüllung 12 sind in einem Stutzen 14 in einer Gehäuseschale 15 gelagert und über einen mit der äußeren Umhüllung 12 fest verbundenen Flansch 16 auf dem äußeren Ende des Stutzens 14 fixiert. Ein inneres Ende 36 der Brennkammer 10 liegt abgedichtet in einem Überleitkanal 17, der den Ausgang der Brennkammer 10 mit einem im Querschnitt kreisringförmigen Gaskanal 18 in einer Turbine verbindet. Um den Gaskanal 18 auf seinem Umfang möglichst gleichmäßig mit Heißgas 34 zu beaufschlagen, ist eine Vielzahl von beispielsweise 10 bis 30 Brennkammern 10 gleichmäßig auf den Umfang der Turbinenanlage verteilt und sind deren Einmündungen in den Überleitkanal 17 durch einen in Richtung auf den Gaskanal 18 offenen Umfangskanal 19 miteinander verbunden. Der Überleitkanal 17 ist durch dünne Stege 21 an einem Leitteil 22 der Turbine verankert.The combustion chamber 10 and the outer casing 12 holding it are mounted in a socket 14 in a housing shell 15 and one fixed to the outer sheath 12 Fixed flange 16 on the outer end of the nozzle 14. One inner end 36 of the combustion chamber 10 is sealed in one Überleitkanal 17, the output of the combustion chamber 10 with a in cross-section annular gas channel 18 in one Turbine connects. To the gas channel 18 on its circumference as possible uniformly apply hot gas 34 is a Variety of, for example, 10 to 30 combustion chambers 10 evenly distributed on the circumference of the turbine plant and are their Junctions in the transfer channel 17 by a in the direction on the gas channel 18 open peripheral channel 19 with each other connected. The transfer channel 17 is connected by thin webs 21 anchored a guide member 22 of the turbine.

Um die in Richtung der Pfeile 5 strömende komprimierte Luft möglichst verlustarm aus dem Kanalabschnitt 6 in die die Brennkammern 10 umhüllenden Kanäle 20 umzuleiten, trägt der Umlenker 8 einen in Richtung auf das freie Ende der Brennkammern 10 weisenden Querschnittsschenkel. Dessen Rand 35 folgt wellenförmig in geringem Abstand der Kontur des Überleitkanals 17 und den Konturen der in diesen einmündenden Enden 16 der Brennkammern 10. Auf diese Art und Weise ist der Luftstrom aus dem Kanalabschnitt 6 um mehr als 90° in eine Richtung parallel zu den Achsen der Brennkammern 10 umgelenkt. Dadurch sind die Brennkammern 10 mit ihren Achsen ohne besondere Nachteile stark geneigt gegen die Achse 1 positionierbar, wobei sie mit deren verdichterseitigem Ende einen spitzen Winkel einschließen, so daß sie auf einem zur Achse 2 konzentrischen Kegelmantel liegen.Around the flowing in the direction of arrows 5 compressed air as low loss from the channel section 6 in the Divert combustion chambers 10 enveloping channels 20 carries the Deflector 8 one in the direction of the free end of the combustion chambers 10 pointing cross-sectional legs. Its edge 35 follows Wavy at a small distance from the contour of the transfer channel 17 and the contours of the opening ends 16 in this the combustion chambers 10. In this way is the air flow from the channel section 6 by more than 90 ° in one direction deflected parallel to the axes of the combustion chambers 10. As a result, the combustion chambers 10 with their axes without special Disadvantages strongly tilted against the axis 1 positionable, whereby they tip off with their compressor-side end Include angle so that they are on one to the axis 2 concentric cone sheath lie.

Das Leitteil 22 und ein Leitteil 23 sind in einer Gehäuseschale 24 gelagert und sind durch Sicherungsklötze 25 gegen Rotation gesichert. Andererseits sind die Leitteile 22 und 23 jedoch durch beispielsweise hydraulische oder pneumatische Motoren 26 achsparallel über geringe Distanzen verschiebbar, wobei ein Flansch 27 elastisch verformt ist und mit in ihm gespeicherter Verformungsenergie zur Rückstellung des Leitteiles 22 bzw. 23 dient. Ein von den Gehäuseschalen 15 und 24 umfaßtes Volumen ist durch Trennwände 28 in Kammern unterteilt.The guide member 22 and a guide member 23 are in a housing shell 24 stored and are by securing blocks 25 against Rotation secured. On the other hand, the guiding parts 22 and 23 However, for example, by hydraulic or pneumatic Motors 26 axially parallel over small distances, wherein a flange 27 is elastically deformed and with in it stored deformation energy for the return of the Leitteiles 22 or 23 is used. One of the housing shells 15 and 24 enclosed volume is divided by partitions 28 into chambers.

Die Leitteile 22 und 23 haben eine trichterartige Gestalt und tragen auf ihrer Innenseite in Leitringen 29 befestigte Leitschaufeln 30, deren den Leitringen 29 gegenüberliegende Enden durch Ringe 31 fest miteinander verbunden sind. Zwischen einander benachbarten Kränzen aus Leitschaufeln 30 ist je ein Kranz aus auf dem Läufer 1 verkeilten Laufschaufeln 32 vorgesehen, deren freien Spitzen Leitringe 33 gegenüberliegen. Dabei bilden die Leitringe 29 und 33 eine äußere und die Ringe 31, zusammen mit Füssen der Laufschaufeln 32, eine innere Begrenzung des Gaskanals 18 für das Heißgas 34 in der Turbine.The guide parts 22 and 23 have a funnel-like shape and carry on their inside in guide rings 29 fixed vanes 30, whose guide rings 29 opposite ends are firmly connected together by rings 31. Between each other adjacent wreaths of vanes 30 is ever one Wreath provided on the rotor 1 wedged blades 32, whose free tips guide rings 33 are opposite. there the guide rings 29 and 33 form an outer and the rings 31, together with feet of the blades 32, an inner boundary the gas channel 18 for the hot gas 34 in the turbine.

Unmittelbar dem Heißgas 34 ausgesetzte Teile der Turbinenanlage sind über nicht dargestellte Kanäle üblicherweise durch Anzapfluft aus dem Verdichter oder aus dem Kanalabschnitt 6 gekühlt. In besonderen Einsatzfällen sind, soweit erforderlich, auch unmittelbar an den Überleitkanal 17 angrenzende, in einem toten Winkel des Luftstroms nahe dem Umlenker 8 liegende Taschen auf diese Art gekühlt. Diese Taschen sind dann zweckmäßig durch nicht dargestellte Trennwände vom Luftkanal getrennt, so daß dessen freier und wirksamer Querschnitt gerade im Bereich des Überleitkanal 17 genauer an den Querschnitt des Kanalabschnitts 6 bzw. die Summe der Einzelquerschnitte der Kanäle 20 anpassbar ist. Dieser Querschnitt ist darüber hinaus durch Variation der Wanddicke des Umlenkers 8 sowohl in dessen Umfangsrichtung als auch in dessen Querschnitt genau einstellbar.Immediately the hot gas 34 exposed parts of the turbine system are usually not shown on channels Bleed air from the compressor or from the channel section 6 cooled. In special cases, as far as necessary, also directly adjacent to the transfer channel 17, in a blind spot of the air flow near the deflector 8 lying Bags cooled in this way. These bags are then expediently by dividing walls, not shown, from the air duct separated, so that its free and effective cross-section straight in the region of the transfer channel 17 more precisely to the cross section of the channel section 6 or the sum of the individual cross sections the channels 20 is customizable. This cross section is In addition, by varying the wall thickness of the deflector. 8 both in its circumferential direction and in its cross section exactly adjustable.

Da der Querschnitt des Kanalabschnitts 6 und die Summe der Einzelquerschnitte der Kanäle 20 mindestens annähernd gleich groß sind, ist in diesen Kanalabschnitten eine konstante, gleich große Strömungsgeschwindigkeit für die verdichtete Luft gewährleistet. Diese Strömungsgeschwindigkeit ist durch die besondere Form des im Querschnitt C-förmigen Umlenkers 8 auch während der Umlenkung der verdichteten Luft um mehr als 90° beibehalten. Dadurch sind Verzögerungen und erneute Beschleunigen der verdichteten Luft vermieden und dadurch sind hierdurch bedingte Verluste stark vermindert.Since the cross section of the channel section 6 and the sum of Single cross sections of the channels 20 at least approximately equal are large, is a constant, in these channel sections, same flow rate for the compressed Air guaranteed. This flow rate is through the special shape of the cross-sectionally C-shaped deflector. 8 even during the redirection of the compressed air by more than Retain 90 °. This causes delays and renewed acceleration avoided the compressed air and thereby are This greatly reduced losses.

Claims (13)

  1. Gas turbine with a compressor for air, which is heated in a plurality of combustion chambers (10) connected in parallel with respect to flow, before it flows via a transfer duct (17) to a gas duct (18) in a turbine, characterized in that the compressed air flows with approximately constant velocity over the whole distance in an air duct (6, 20) from the outlet (4) of the compressor to the inlet into the combustion chambers (10), approximately constant velocity being achieved by that the effective cross section of the air duct (6, 20) being almost constant over the whole distance from the outlet (4) of the compressor to the inlet into the combustion chamber (10).
  2. Gas turbine according to Claim 1, characterized in that the air duct (6, 20) enforces a change in direction of more than 90° on the air flowing in the region of the transfer duct (17) and in that a deflector (8) is provided in the air duct (6, 20) in this region only.
  3. Gas turbine according to one of Claims 1 to 2, characterized in that the deflector (8) is represented by a C-shaped cross section ring.
  4. Gas turbine according to one of Claims 1 to 3, characterized in that the wall thickness of the deflector (8) is very different both in cross section and in the peripheral direction and, by this means, matches an effective cross section of the air duct in its region to the constant cross section of the air duct.
  5. Gas turbine according to one of Claims 1 to 4, characterized in that the free end (9) of one arm of the cross section of the deflector (8) is located on a cylindrical envelope concentric with the turbine centre line (2) and in that the free end (35) of the other arm follows, in wave shape and at a small distance, contours of the combustion chambers (10).
  6. Gas turbine according to one of Claims 1 to 5, characterized in that the arm of the C-shaped cross section following the contours of the combustion chambers (10) with wave-shaped edge (35) over its length respectively achieves a minimum under a combustion chamber centre line and respectively achieves a maximum under an intermediate space between adjacent combustion chambers (10).
  7. Gas turbine according to one of Claims 1 to 6, characterized in that the air duct (26) opens into more than ten and up to thirty combustion chambers (10) evenly distributed over the periphery of the turbine.
  8. Gas turbine according to one of Claims 1 to 7, characterized in that an average length of a heated gas flow (34) within the transfer duct (17) from the outlet (16) of the combustion chambers (10) to the inlet into a gas duct (18) in the turbine is approximately equal to twice the width of this gas duct (18) at the inlet into the turbine, so that the length of this gas flow (34) in the transfer duct (17) is shorter than the diameter of one of the combustion chambers (10).
  9. Gas turbine according to one of Claims 1 to 8, characterized in that centre lines of the combustion chambers (10) are located on a conical envelope and include an acute angle with the turbine centre line (2).
  10. Gas turbine according to one of Claims 1 to 9, characterized in that the air duct fans out, along the distance from the deflector (8) to the opening into the combustion chambers (10), into a number of partial air ducts (20) equal to the number of the combustion chambers (10), which partial air ducts (20) together have approximately the constant cross section of the air duct (6).
  11. Gas turbine according to one of Claims 1 to 10, characterized in that the partial air ducts (20) of adjacent combustion chambers penetrate each other at their turbine end (16), while outer walls (12) of the partial air ducts (20) are provided with a corresponding recess in this region.
  12. Gas turbine according to one of Claims 1 to 11, characterized in that the deflector (8) is supported by struts (7) by means of its cross-sectional arm (9) located upstream in the air duct (6), which struts are arranged approximately radially in the end of a circular cross section of the air duct (6).
  13. Gas turbine according to one of Claims 1 to 12, characterized in that cross-sectional arms of the C-shaped cross section deflector (8) form wavy lines opposite to one another in the peripheral direction, the wave length of which waves corresponds to the distance of the combustion chambers from one another.
EP01114599A 2001-06-18 2001-06-18 Gas turbine with an air compressor Expired - Lifetime EP1270874B1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
DE50107283T DE50107283D1 (en) 2001-06-18 2001-06-18 Gas turbine with a compressor for air
EP01114599A EP1270874B1 (en) 2001-06-18 2001-06-18 Gas turbine with an air compressor
JP2002172518A JP2003042451A (en) 2001-06-18 2002-06-13 Gas turbine having air compressor
US10/172,016 US6672070B2 (en) 2001-06-18 2002-06-17 Gas turbine with a compressor for air
CNB021233160A CN1328492C (en) 2001-06-18 2002-06-18 Gas turbine with air compressor

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP01114599A EP1270874B1 (en) 2001-06-18 2001-06-18 Gas turbine with an air compressor

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EP1270874A1 EP1270874A1 (en) 2003-01-02
EP1270874B1 true EP1270874B1 (en) 2005-08-31

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US (1) US6672070B2 (en)
EP (1) EP1270874B1 (en)
JP (1) JP2003042451A (en)
CN (1) CN1328492C (en)
DE (1) DE50107283D1 (en)

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US20030010014A1 (en) 2003-01-16
CN1392331A (en) 2003-01-22
DE50107283D1 (en) 2005-10-06
JP2003042451A (en) 2003-02-13
CN1328492C (en) 2007-07-25
US6672070B2 (en) 2004-01-06
EP1270874A1 (en) 2003-01-02

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