EP2551453A1 - Cooling device of a gas turbine compressor - Google Patents

Abstract

The compressor rotor has a ring (26) in area of compressor rotor ends. The ring forms a gap (29) over a rotor disk (25). The ring is fastened on the rotor disk. The ring has grooves (20) for accommodating rotor blades (21) in the region of the compressor rotor end. A unit is provided for directing an axial flow of cooling medium from the compressor rotor end through the ring. Another unit is provided for deflecting the cooling medium such that the cooling medium flows back in the axial direction through an annular gap (29) between the ring and the rotor disk. Independent claims are included for the following: (1) a gas turbine with a compressor, combustion chamber, turbine and rotor; and (2) a method for cooling a compressor rotor of a gas turbine.

Description

TECHNICAL AREA

The present invention relates to the field of turbomachinery. It relates to a compressor rotor according to the preamble of claim 1 and a gas turbine comprising such a rotor and a method for cooling a gas turbine with such a rotor.

STATE OF THE ART

Fig. 1 shows the basic scheme of a gas turbine, as used for example as a stationary industrial turbine for the production of energy. The gas turbine 10 of Fig. 1 comprises a compressor 12 which sucks and compresses combustion air through an air inlet 11. The compressed air is introduced into a combustion chamber 13 and used there for the combustion of a fuel 14. The resulting hot gases are in a subsequent Relaxed turbine 15 under work and discharged as exhaust 16 to the outside or used in a heat recovery steam generator.

The blades required for the compressor 12 and the turbine 15 are usually mounted on a rotor 17 having corresponding rotor disks. In the compressor 12, compression air temperatures at the compressor end of several 100 ° C arise during the compression of the combustion air. Cooling of the rotor in this area on the one hand reduces the thermal load on the materials used, but on the other hand can also contribute to improving the overall efficiency of the gas turbine. For cooling, a part of the compressed air can be branched off, in a cooling device 18 (in Fig. 1 dashed) cooled down and then fed to the cooling in the end region of the compressor 12.

In the publication EP 0 799 971 B1 It has been proposed to provide a thermal insulation in the end region of the compressor to protect the rotor from thermal overload in the end region of the compressor, which reduces the heat input from the compressor passage in the rotor body. However, this is a purely passive measure that does not allow dissipation of heat

From the GB 2 350 408 A It is also known, in the rotor of a turbomachine, to arrange a concentric annular thermal shield at a distance around the rotor, which carries the blades and reduces the heat input into the rotor. In addition, a cooling medium, which dissipates heat, can flow through a gap between the ring and the rotor. The disadvantage here is that the shielding makes special demands on the material for lack of own cooling. In addition, cooling medium flowing through the shielding elements can only be returned to the compressor, thereby increasing the compressor discharge temperature. In addition, this cooling medium is no longer available for the combustion chamber or turbine cooling.

Today's design at the end of the compressor, beyond which the invention goes, is according to Fig. 2 from compressor blades 21 which are fixed in circumferential grooves 20 'on the rotor 17 and the rotor disk 25'. In most industrial gas turbines, a portion of the compressed compressor air is tapped and fed instead of the combustion, used as cooling air of hot parts (rotor, hot gas parts). To improve cooling efficiency, some of the compressor air is sent through a cooler to lower the temperature of the cooling medium (see above). In the gas turbine of Fig. 2 a portion of this pre-cooled cooling air 24 via (stationary) structural parts 23 'of a downstream to the compressor 12 subsequent central part 19 is guided back to the end of the compressor 12. The cooling air is used here for flushing the cavity 22 between the compressor rotor end and the middle section 19 and as cooling air for the rotor disk 25 'in the region of the compressor rotor end. The aim is to use the cooling air to lower the rotor temperature in this area.

Although the goal is to lower the rotor temperature in this range with the cooling air, this type of cooling is not efficient enough for the rotor disk 25 '. To increase the power and thus the efficiency of a gas turbine, the combustion temperature and / or the mass flow can be increased. An increase in performance can be achieved by an improved compressor. This results in a higher mass flow, so that the pressure and thus the air temperature at the end of the compressor increases and thereby the rotor temperature also increases. With a higher rotor temperature in the region of the compressor rotor end, however, the life of the rotor is adversely affected.

PRESENTATION OF THE INVENTION

It is therefore an object of the invention to provide a rotor which avoids the disadvantages of prior rotors described and in particular characterized in that in a comparatively simple construction, the thermal load of the rotor is significantly reduced at the compressor rotor end.

These and other objects are achieved by the entirety of the features of claim 1.

The rotor according to the invention, which is intended in particular for use in a gas turbine, comprises a rotor, which has at least one groove into which a plurality of rotor blades can be inserted and held on the rotor, and furthermore a device for cooling the rotor in the area of the compressor rotor end. The invention is characterized in that the rotor in the region of the compressor rotor end has a ring which is pushed concentrically and at a distance with formation of a gap over a rotor disk of the rotor and attached to the rotor disk, that the blades in the region of the compressor rotor end in corresponding grooves on Ring are used and held there, that first means for axially flowing through the ring are provided with a cooling medium from Kompressorrotorende forth, and in that second means for deflecting the emerging from the ring cooling medium are provided, such that the cooling medium through the gap between the Ring and the rotor disc enclosed by the ring flows back in the axial direction. The gap between the ring and the rotor disc enclosed by the ring, for example, has the shape of an annular gap, which may be interrupted by fastening elements which connect the ring with the rotor disk.

An embodiment of the compressor according to the invention is characterized in that the first means comprises a plurality of distributed over the circumference of the ring arranged axial bores through which the cooling medium flows.

Another embodiment of the compressor according to the invention is characterized in that the second means comprise an annular deflection region formed in the rotor disk, which communicates with the first means or axial bores and the gap between the ring and the rotor disk, and a reversal of the flow direction of the cooling medium causes.

According to another embodiment, the ring is fastened by a positive connection between the inner circumferential surface of the ring and the outer circumferential surface of the rotor disk on the rotor disk.

Typically, the positive engagement is in the nature of circumferentially distributed, radially aligned hammerhead connections or fir tree root connections.

A further embodiment of the invention is characterized in that the ring with the upstream end face abuts against an annular abutment surface of the rotor disk, and that the ring and the rotor disk are connected to one another in this region.

The connection between ring and rotor disk can be effected by a positive connection.

But it is also conceivable that the connection between the ring and the rotor disk by a material bond, in particular by welding, is effected.

In addition to the compressor rotor is a gas turbine comprising a compressor, a combustion chamber, a turbine and a rotor subject of the invention, wherein the rotor (34) comprises a compressor rotor (17) according to one of the embodiments described above.

An embodiment of the gas turbine according to the invention is characterized in that the ring is arranged in the installed state at the downstream end side next to stationary structural parts, and that the cooling medium for cooling the compressor rotor end is brought over the structural parts.

Preferably, deflecting elements are arranged at the transition between the structural parts and the ring, which impart a twist in the direction of rotation of the compressor to the cooling medium emerging from the structural parts.

In particular, the deflecting elements may be formed as baffles.

But it is also just as conceivable that the deflection elements are designed as swirl nozzles.

According to a further embodiment of the invention, at least one seal is arranged between the structural parts and the ring.

In particular, the seal may be formed as a labyrinth seal or brush seal.

According to one embodiment, such a seal is mounted on a radius which is smaller than the distance from the center of the rotor to the first means for axially flowing through the ring with a cooling medium. This seal prevents a bypass of the cooling medium around the ring.

According to one embodiment, such a seal is mounted on a radius which is greater than the distance from the center of the rotor to the first means for axially flowing through the ring with a cooling medium. This seal prevents backflow of the cooling medium into the main flow of the compressor.

In addition to the compressor rotor and the gas turbine, a method for cooling a compressor rotor of a gas turbine according to the invention described above is the subject of the invention. The gas turbine comprises a compressor, a combustion chamber and a turbine. The compressor itself has a plurality of blades, which are inserted into corresponding grooves on a compressor rotor and held there. Furthermore, the compressor rotor in the region of the compressor rotor end has a ring which is pushed concentrically and with the formation of a gap over a rotor disk of the compressor rotor and attached to the rotor disk, wherein the rotor blades are inserted into corresponding grooves on the ring in the region of the compressor rotor end and held there , Further, in the ring first means are provided for axially flowing through the ring with a cooling medium from the Kompressorotorende forth and second means for deflecting the emerging from the ring cooling medium is provided. The method is characterized in that a cooling medium from the compressor end is passed through the first means of the ring, the cooling medium is subsequently deflected by second means, and the cooling medium is finally returned through the gap between the ring and the ring surrounded by the rotor disc in the axial direction ,

According to one embodiment of the method, cooling medium is subjected to a twist before it is introduced into the first means of the ring.

All explained advantages can be used not only in the respectively specified combinations, but also in other combinations or alone, without departing from the scope of the invention. For example, the compressor rotor described by way of example of a gas turbine having a compressor, a combustor, and a turbine may be used for gas turbines with sequential combustion, ie, gas turbines containing one or more compressors, a first combustor, a high pressure turbine, a second combustor (sequential combustor ) and a low pressure turbine. Corresponding is also a gas turbine with sequential combustion and the rotor according to the invention and a method for cooling a compressor rotor for a gas turbine with sequential combustion included in the invention.

BRIEF EXPLANATION OF THE FIGURES

Fig. 1

das grundsätzliche Schema einer Gasturbine, wie sie zur Verwirklichung der Erfindung geeignet ist;

Fig. 2

einen Längsschnitt durch eine Gasturbine im Bereich des Kompressorrotorendes mit einer Kühlung, wie sie bisher verwendet worden ist;

Fig. 3

in einer zu Fig. 2 vergleichbaren Darstellung ein Kompressorende mit einer verbesserten Kühlung gemäss einem Ausführungsbeispiel der Erfindung; und

Fig. 4

den Querschnitt in der Ebene A-A durch den Kompressor gemäss Fig. 3.

The invention will be explained in more detail with reference to embodiments in conjunction with the drawings. Show it

Fig. 1

the basic scheme of a gas turbine, as it is suitable for the realization of the invention;

Fig. 2

a longitudinal section through a gas turbine in the region of the compressor rotor end with a cooling, as has been used so far;

Fig. 3

in one too Fig. 2 comparable illustration of a compressor end with improved cooling according to an embodiment of the invention; and

Fig. 4

the cross section in the plane AA through the compressor according to Fig. 3 ,

WAYS FOR CARRYING OUT THE INVENTION

In order to improve the cooling in the region of the compressor rotor end, according to the present invention, a cooling circuit is produced below the high-pressure compressor or compressor rotor end with the aid of a separate ring.

Like from the Fig. 3 becomes clear, the ring 26 is pushed onto the rotor disk 25 during manufacture. The connection between ring 26 and rotor disk 25 can be made in different ways. On the one hand, according to Fig. 4 a positive fit 30 between the opposite lateral surfaces of the ring 26 and the rotor disk 25 are used, which in particular has the form of a distributed over the circumference, radially oriented hammer head connection.

On the other hand, there is the possibility of also connecting the upstream end face of the ring 26 with the adjacent abutment surface 31 of the rotor disk 25 via positive locking, or else via a material connection (in particular by welding). A self-assurance during operation is also present, because with the rotor running 17, the thrust of the ring 26 is directed against the stop surface 31.

The cooling air 24 is guided through the structural parts 23 of the middle part to the cavity at the end of the compressor 17. From the cavity, the cooling air passes into distributed over the circumference of the ring 26 arranged axial bores 27 in the ring 46. At the upstream end of the ring 26, the exiting the ring 26 cooling air is deflected in a deflection (180 °) and passes through the gap 29 between the rotor disk 25 and the ring 26 again towards the turbine.

Between the ring 26 and the structural parts 23 of the middle part 19, a seal 32 is preferably provided in order to minimize the slight leakage. This can e.g. a conventional labyrinth seal or brush seal.

In order to minimize the entry losses of the cooling air into the ring 26, deflecting elements 33, in particular in the form of a swirl nozzle or baffles, can be provided at the outlet of the cooling air from the structural parts 23, which impart a twist in the direction of rotation of the compressor to the cooling air emerging from the structural parts 23 ,

Through the gap 29 between the ring 26 and the rotor 17, the contact surface between the ring 26 and rotor 17 is further reduced. This reduces the heat conduction from the hot ring 26 into the rotor 17.

• Das System verwendet gekühlte Kompressorluft die über das Sekundärsystem zugeführt wird.
• Die Schaufelzahl im Kompressor 12 ist frei wählbar, weil im separaten Ring 26 umlaufende Nuten wie zuvor vorhanden sind.
• Der separate Ring 26 kann je nach Fixierung über Formschluss oder Stoffschluss aus einem anderen Material als die Rotorscheibe 25 bestehen.
• Axialbohrungen 27 durch den Ring 26 ermöglichen der Kühlluft das Durchströmen und Kühlen des Rings 26.
• Eine Rückführung der Kühlluft bei der Formschlussverbindung ist gegeben.
• Ein doppeltes Dichtsystem sorgt dafür, dass die unterschiedlichen Kühlluftströme gegeneinander getrennt werden.
• Die Kühlluft kann nach dem Durchströmen des Rings 26 weiter für die Kühlung der Brennkammer bzw. Turbine verwendet werden.

Overall, the main features of the invention are:

• The system uses cooled compressor air supplied via the secondary system.
• The number of blades in the compressor 12 is arbitrary, because in the separate ring 26 circumferential grooves as before exist.
• The separate ring 26 may be made of a different material than the rotor disk 25 depending on the fixation via positive engagement or material connection.
• Axial bores 27 through the ring 26 allow the cooling air to flow through and cool the ring 26th
• A return of the cooling air in the positive connection is given.
• A double sealing system ensures that the different cooling air flows are separated from each other.
• The cooling air can continue to be used for the cooling of the combustion chamber or turbine after flowing through the ring 26.

LIST OF REFERENCE NUMBERS

10

gas turbine

11

air intake

12

compressor

13

combustion chamber

14

fuel

15

turbine

16

exhaust

17

compressor rotor

18

Cooling device (external)

19

midsection

20.20 '

Groove (encircling)

21

blade

22

cavity

23.23 '

structure part

24

cooling air

25.25 '

rotor disc

26

ring

27

axial bore

28

deflection

29

gap

30

form-fit

31

stop surface

32

poetry

33

Deflector (e.g., baffle, swirl nozzle)

34

rotor

Claims (15)

Compressor rotor (17), comprising at least one groove (20) for receiving moving blades (21), and a device (26, 27, 28, 29) for cooling the compressor rotor (17) in the region of the compressor rotor end, characterized in that the compressor rotor (17) in the region of the compressor rotor end has a ring (26) which concentrically and at a distance by forming a gap (29) over a rotor disk (25) of the compressor rotor (17) pushed and fixed to the rotor disk (25) that the Ring (26) has grooves (20) for receiving blades (21) in the region of the compressor rotor end, that first means (27) for axially flowing through the ring (26) are provided with a cooling medium from the compressor rotor end, and that second means (28 ) are provided for deflecting the emerging from the ring (26) cooling medium, such that the cooling medium through the annular gap (29) between the ring (26) and the ring (26) enclosed rotor disc (25) in the axial Direction flows back. Compressor rotor according to claim 1, characterized in that the first means comprises a plurality of over the circumference of the ring (26) distributed axial bores (27) through which the cooling medium flows. Compressor rotor (17) according to claim 1 or 2, characterized in that the second means comprise an annular deflection region (28) formed in the rotor disk (25), which is connected to the first means or axial bores (27) and the annular gap (29) between the ring (26) and the rotor disc (25) is in communication and causes a reversal of the flow direction of the cooling medium. Compressor rotor (17) according to one of claims 1-3, characterized in that the ring (26) by a positive connection (30) between the inner lateral surface of the ring (26) and the outer circumferential surface of the rotor disc (25) on the rotor disc (25) is fixed. Compressor rotor (17) according to any one of claims 1-3, characterized in that the ring (26) abuts with the upstream end face on an annular abutment surface (31) of the rotor disc (25), and that the ring (26) and the rotor disc (25) are interconnected in this area. Compressor rotor (17) according to claim 5, characterized in that the connection between the ring (26) and rotor disc (25) is effected by a positive connection. Compressor rotor (17) according to claim 5, characterized in that the connection between the ring (26) and rotor disc (25) by a material bond, in particular by welding, is effected. Gas turbine (10) comprising a compressor (12), a combustion chamber (13), a turbine (15) and a rotor (34), characterized in that the rotor (34) comprises a compressor rotor (17) according to one of claims 1 to 8 includes. Gas turbine (10) according to claim 8, characterized in that the ring (26) on the downstream end side directly adjacent stationary structural parts (23) is arranged, and that the cooling medium for cooling the compressor rotor end via the structural parts (23) is introduced. Gas turbine (10) according to claim 9, characterized in that at the transition between the structural parts (23) and the ring (26) deflecting elements (33) are arranged, which impose a spin in the direction of rotation of the compressor from the cooling medium emerging from the structural parts (23) , Gas turbine (10) according to claim 10, characterized in that the deflecting elements (33) are designed as baffles. Gas turbine (10) according to claim 10, characterized in that the deflecting elements (33) are designed as swirl nozzles. Gas turbine (10) according to claim 8, characterized in that between the structural parts (23) and the ring (26) a seal (32) is arranged. Method for cooling a compressor rotor (17) of a gas turbine (10), wherein the gas turbine (10) comprises a compressor (12), a combustion chamber (13) and a turbine (15),
the compressor (12) having a plurality of blades (21) inserted and held in respective grooves (20) on a compressor rotor (17),
the compressor rotor (17) has in the region of the compressor rotor end a ring (26) which is pushed concentrically and forming a gap (29) over a rotor disk (25) of the compressor rotor (17) and attached to the rotor disk (25) and the rotor blades (21) in the region of the compressor rotor end are inserted into corresponding grooves (20) on the ring (26) and held there,
and first means (27) are provided for axially flowing through the ring (26) with a cooling medium from the compressor rotor end, and second means (28) are provided for diverting the cooling medium emerging from the ring (26),
characterized in that a cooling medium from the compressor end through the first means (27) of the ring (26) is passed, the cooling medium is then deflected by second means (28), and the cooling medium finally through the gap (29) between the ring (26 ) and the annular disc (25) enclosed by the ring (26) is returned in the axial direction. A method according to claim 14, characterized in that the cooling medium is subjected to a twist before it is introduced into the first means (27) of the ring (26).

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