WO2012051997A2

WO2012051997A2 - Turbine wheel arrangement for a gas turbine

Info

Publication number: WO2012051997A2
Application number: PCT/DE2011/001849
Authority: WO
Inventors: Dirk Büchler; Gerhard Buttkewitz
Original assignee: Baltico Gmbh
Priority date: 2010-10-15
Filing date: 2011-10-14
Publication date: 2012-04-26
Also published as: EP2598736A2; WO2012051997A3; DE102010048434A1; DE102010048434B4; CA2814427A1; US20130192231A1

Abstract

The invention relates to a turbine wheel arrangement for a gas turbine comprising two successive turbine wheels (18, 20) rotating in opposite directions, wherein the first turbine wheel (18) comprises flow channels (42) of a Laval cross-sectional shape, distributed over the circumference and having radially inner gas inlets (43) and radially further out gas outlets (45), in each case with a substantially tangential flow direction component, wherein the gas outlets (45) act on the second axially or radially acting turbine wheel (20) in the direction of flow. This has the effect that the thermal energy and compressive energy of the gas at the nozzle inlet is largely converted into flow energy at the outlet of the turbine stage. The rotational speeds of the two rotors coupled to the turbine wheels can be set as desired, allowing the operational states of the two systems to be optimally set without adjusting systems.

Description

Turbine wheel arrangement for a gas turbine

The invention relates to a turbine wheel arrangement for a gas turbine with two successive, opposite to each other rotating turbine wheels.

The currently known gas turbines are designed for the highest possible performance with the highest possible efficiencies, which means that the components and materials used are loaded to the maximum permissible limits. This applies in particular to the turbine blades of the first turbine stage, which are subjected to the most extreme stresses, with the result that very expensive materials are used and expensive cooling measures for blade cooling must be taken. In addition, there are loads due to high mechanical stresses due to the high flow velocities, but above all the large centrifugal forces due to the high rotational speeds. Overall, it is difficult to achieve overall efficiencies higher than 35%.

It is known in steam turbine construction, stationary stator of axial turbines with

Form Lavaldüsenkanalen, which are aligned at an angle of approximately 45 ° to the circumferential direction.

From AT 239606 a gas turbine is known, are guided in the combustion gases via two oppositely oriented radial channels to the outside and then deflected in the circumferential direction and are directed by Laval nozzle-like outlet channels on blades of an oppositely rotating paddle wheel.

The object of the invention is to provide a gas turbine, which is distinguished from conventional turbines by a lower temperature load of the turbine stage and a high temperature and Druckgefälte of the combustion chamber to the turbine blades in a large power range allows, so that high thermal efficiencies can be achieved.

The invention results from the features of the independent claims. Advantageous developments and refinements are the subject of the dependent claims.

The invention advantageously has the effect that the thermal energy and pressure energy of the gas at the nozzle inlet is largely converted into flow energy at the turbine stage outlet. The impulse force of the nozzles drives the turbine wheel. The escaping gas from the Laval nozzles has a very high flow

CONFIRMATION COPY speed is above Mach 1 and is directed to the vanes of the downstream turbine runner, creating a moment and rotation opposite the direction of rotation of the front turbine runner. The speed of the gas jet is much greater than the peripheral speed of the turbine runner. Thus, the kinetic energy remaining in the flow jet from the first energy conversion in the Laval nozzles is utilized. The rear turbine wheel is designed so that the tangential velocity at the outlet ideally almost disappears.

By suitable control of the load on each impeller e.g. By means of generators, the rotational speeds of the two rotors coupled to the turbine wheels can be set as desired, so that the operating conditions of both systems can be set optimally without adjustment systems.

By the invention, the combined loads can be thermally, mechanically and centrifugally separated. Thus, a relaxation in Laval nozzles, which reduce almost all the thermal energy and pressure energy and convert into kinetic energy. The material stresses are therefore essentially reduced to this area. The Laval nozzles themselves can be made of suitable, temperature-resistant materials such as ceramics or metal alloys and can withstand the thermal stresses. The remaining kinetic energy in the gas jet is fully utilized by the pulse in the second turbine stage. Due to the mating principle, the required speed can be halved, which considerably reduces the centrifugal forces.

Due to the concept of the invention very high compression ratios of> 25, high combustion chamber temperatures of> 2,000 ° C and correspondingly high thermal efficiencies of> 60% and overall efficiency of the gas turbine> 50% can be achieved.

Another significant advantage of the invention is that the gap between the wheels has no significant effect on the efficiency. Furthermore, the components can be made simpler and cheaper than in conventional turbines, because it is enough a single-stage turbine to achieve the entire energy conversion. It is also advantageous that cooling is only required at the Laval nozzles, while a second turbine wheel requires no cooling, as a result of which the losses occurring in conventional turbines due to the considerable bleed air for turbine cooling are avoided. to be avoided.

According to an advantageous development of the invention, a gas turbine comprises an inner rotor, on the outside of which several first rows of blades of a multi-stage axial compressor and the downstream turbine wheel are mounted, that this gas turbine further comprises a hollow shaft on the inside several second blade rows of the axial compressor are mounted, which are arranged in the axial direction alternately to the first rows of blades, that further on the hollow at least one combustion chamber and the front turbine runner are mounted. Since alternate in the axial direction of the compressor, the blade rings of the rotating shafts against each other, almost any higher number of stages in opposite constructions are possible. The speeds of the two shafts are preferably the same size. In certain applications, e.g. in a booster mode, but also a speed variance between the two shafts is possible.

Further advantages, features and details will become apparent from the following description in which, with reference to the drawings embodiments are described in detail. Described and / or illustrated features form the subject of the invention, or independently of the claims, either alone or in any meaningful Kombinatton, and may in particular also be the subject of one or more separate application / s. The same, similar and / or functionally identical parts are provided with the same reference numerals.

Show it:

1 shows an axial section through a gas turbine as a power generator,

FIG. 2 shows a cross section through the gas turbines according to FIG. 1,

FIG. 3 shows an axial section through a Pelton-type gas turbine;

Figure 4: a cross section through the gas turbine according to Figure 3, and

Figure 5: a perspective view of a gas turbine with axial turbine stage.

1 shows a gas turbine 10a in axial section, which consists of a multi-stage axial compressor 12, a single-stage centrifugal compressor 4, a combustion chamber 16, a first radial turbine runner 18 and a second radial turbine runner 20. The two radial turbine wheels 18, 20 together form a single-stage counter-rotating turbine. Each second blade ring of the axial compressor 12 and the second turbine wheel 20 are mounted together on an inner rotor 22 which is supported by bearings 24 on a fixed axis 26. On the inner rotor 22, the rotor 28a of a first generator is further attached, the outer stator is not shown.

A hollow-shaft-shaped outer rotor 30 carries on the inside the respective other blade rings of the axial compressor 12, further the blade ring of the centrifugal compressor 14, the combustion chamber 16 and the first turbine runner 18. The outer rotor 30 is supported via a first bearing 32a on the fixed axis 26 and a second bearing 32b on the inner rotor 22 from. For this purpose, the outer rotor 30 is connected via a first blade ring 34 with a rotor hub 36 which is supported on the shaft 26 via the front bearing 32a. Radially outwardly, a rotor 28b of a second generator, whose outer stator is not shown, is mounted on the outer rotor 30.

Figure 2 shows a cross section through the gas turbine 10a according to Figure 1 in the axial height of the turbine wheels 18, 20, as shown by the arrow II. FIG. 2 shows a ring of molded parts 40 of the first turbine runner 18 (FIG. 1), which have a shape that forms between these flow passages 42, each with the cross-section of a Laval nozzle. The flow channels 42 comprise substantially radially aligned gas inlets 43 and substantially tangentially or circumferentially oriented gas outlets 45. The Laval cross-sectional shape of the flow channels 42 is preferably in only one direction, namely the circumferential direction, while there is a constant cross-section in the axial direction. Alternatively, it is also possible to design the shape of the molded parts 40 such that a cross-sectional change also takes place in the axial direction. The mold parts 40 are preferably formed on an end plate 41, which closes off the flow channel for the gas flow in the axial direction.

Radially outside of the mold parts 40, outer turbine blades 44 of the second turbine runner 20 are arranged, which are acted on by the gas flow emerging through the flow passages 42 with supersonic and are set in rotation opposite to the first turbine runner 18.

In operation, in the gas turbine 10a shown in FIGS. 1 and 2, air is drawn into the inlet located at the first blade ring 34 and compressed by the axial compressor 12. The axial compressor 12 compresses the air via the two oppositely rotating shafts 22 and 30 with the blade rows arranged alternately thereon. Downstream of the axial compressor 12, the air flow is radially outward deflected and enters the blade ring of the single-stage centrifugal compressor 14, via which the pressure is further increased. Behind the outlet of the radial compressor 14, the air flow is again deflected radially inward and enters the inlet of the rotating combustion chamber 16, injected into the fuel and the air-fuel mixture is burned. Downstream of the combustion chamber 16, the gas flow is again deflected radially outward and enters the first turbine runner 18. The cross-section Laval nozzle flow channels 42 (Figure 2) of the first turbine runner 18 accelerate the gas flow to more than Mach 1, redirecting this in the circumferential direction and encounter the gas flow then predominantly in the circumferential direction. Due to the resulting angular momentum, the first turbine runner 18 is rotated in the counterclockwise direction as viewed in FIG. The supersonic gas flow then impinges on the outer turbine blades 44 of the second turbine runner 20 and is deflected there, which accordingly rotates inversely to the first turbine runner 18.

The rotation of the first turbine runner 18 is transmitted via the outer shaft 30 to the radial compressor 14 and the outer rotor-side blade rows of the axial compressor 12 and the generator rotor 28b. The opposite rotation of the second turbine runner 20 is transmitted to the generator rotor 28 a and via the inner rotor 22 to the inner rotor-side compressor blades of the axial compressor 12.

The gas turbine 10a shown in FIGS. 1 and 2 serves to generate electrical energy. Alternatively, it may also be used for other applications, e.g. be used as the engine of a motor vehicle, sea or aircraft. The generator rotors 28 may also be operated as starters for the shafts 22, 30.

FIGS. 3 and 4 show a gas turbine 10b in which, in contrast to the embodiment according to FIGS. 1 and 2, only a single-stage Pelton turbine is used. The same reference numerals as in FIGS. 1 and 2 denote the same components.

The gas turbine embodiment 10b also includes a multi-stage axial compressor 12, which comprises a hollow shaft 30, to which each second row of blades is fastened in the axial direction. In between blade rows are provided which are fixed to a main rotor 52. The main rotor 52 is supported by bearings 24 on the one hand on the axis 26 and on the other hand on a housing, not shown. On the main rotor 52, a Peltonturbinenrad 54 is further formed, which is shown in Figure 4 in cross-section and explained in more detail below. On the main rotor 52, a mounting flange 56 is further provided for discharging the generated power. The hollow shaft 30 comprises a blade ring of a centrifugal compressor 14 and is connected to a rotating combustion chamber 16 and a nozzle impeller 58. A first blade ring 34 connects the hollow shaft 30 with a hub shaft 60, which in turn is supported by means of a front bearing 32 a on the housing, not shown. At the other end, the nozzle impeller 58 is supported on the main rotor 52 by means of the bearing 32b. On the hub shaft 60, a mounting flange 62 is provided for discharging the power generated in a manner analogous to the main rotor 52.

FIG. 4 shows the nozzle impeller 58 in cross section, which has a number of nozzles 64, which have a predominantly tangential outflow direction. Opposite this, the Pelton turbine wheel 54 has a number of circumferentially distributed Pelton buckets 66, which are acted upon by the nozzles 64.

In operation, the air flow enters through the first blade ring 34 in the multi-stage axial compressor 12 and is compressed there. Subsequently, a further compression in the radial compressor 14, from where the air flow of the co-rotating combustion chamber 16 is supplied. There fuel is added and the mixture is burned. The combustion gases flow through the nozzles 64 and from there they impinge on the Pelton buckets 66. As a result, the jet impeller 58 is rotated in the counterclockwise direction as viewed in FIG. 4, while the Pelton turbine wheel 54 is rotated clockwise. The rotation of the nozzle impeller 58 is transmitted to the hollow shaft 30 and via the blade ring 34 and the hub shaft 60 on the mounting flange 62. Similarly, the rotation of the Peltonturbinenrades 54 is transmitted via the main rotor 52 to the mounting flange 56.

FIG. 5 shows a gas turbine 10c in which, in contrast to the embodiment according to FIGS. 1 and 2, the second turbine stage 20b is designed as an axial stage and not as a radial stage. The same reference numerals as in the other figures designate the same components. In this embodiment, an air supply passage 80 is provided, which supplies heated air from a combustion chamber or a heat exchanger and the first radial turbine stage 18 is applied. In the downstream region, the turbine stage 18 has a channel curvature 82 in the axial direction so that the gas flow leaves the turbine stage 18 at least substantially without a radial velocity component and acts on the second axial turbine stage 20b. The first turbine stage 18 has Laval nozzle-like flow channels 42. Downstream of the axial turbine stage 20b is a gas outlet channel 84. In one variant, the turbine stage 18 may be connected to the air supply passage 80, and in a second vanante, the turbine stage 8 is separated from the air supply passage 80 and thus the air supply passage 80 also forms the housing of the gas turbine.

Claims

claims

A gas turbine (10a) comprising a first turbine runner (18) and a second counter-rotating turbine runner (20) urged by the first turbine runner (20) further comprising an inner shaft (22) having a plurality of first outer ones Blade rows of a multi-stage axial compressor (12) and the second turbine runner (20) are mounted, further comprising a hollow shaft (30) on which the first turbine runner (18) and a plurality of second rows of blades of the axial compressor (12) are mounted, which are alternating in the axial direction are arranged to the first rows of blades,

2. Gas turbine according to claim 1, characterized in that the first turbine runner (18) distributed over the circumference flow channels (42), the gas outlets (45) having Laval cross-sectional shape.

3. Gas turbine according to claim 2, characterized in that the gas outlets (45) have a substantially tangential flow direction component.

4. Gas turbine according to claim 3, characterized in that the gas outlets (45) of the first turbine runner (18) are aligned at an angle of 10 ° to 30 ° to the tangential direction.

5. Gas turbine according to claim 1, characterized in that on the hollow shaft (30) at least one rotating combustion chamber (16) is mounted.

6. Gas turbine according to claim 1 or 2, characterized in that the flow channels (42) have a lavaldüsenartigen cross-sectional profile in a transverse flow direction and have no cross-sectional change in the other transverse flow direction.

7. Gas turbine according to claim 2, characterized in that the first turbine runner (18) comprises a disc (41) are mounted on the circumferentially distributed moldings (40) or molded, between which the flow channels (42) with Laval cross-sectional shape are formed.

8. Gas turbine according to one of the preceding claims, characterized characterized in that the downstream turbine runner (54)

Peltonturbinen blades (66).

9. Gas turbine according to claim l.dadurch in that the second turbine impeller (20b) is an axial turbine wheel.

10. A gas turbine according to claim 1, characterized in that the second turbine runner (20) is a radial turbine wheel.

11. Gas turbine according to claim 5, characterized in that each flow channel (42) is associated with Laval cross-sectional shape of a co-rotating combustion chamber (16).

12. Gas turbine according to claim 1, characterized in that a radial compressor (1) is mounted on the hollow shaft (30) downstream of the axial compressor (12).

13. Gas turbine according to one of the preceding claims, characterized in that on the inner rotor 22 of the rotor (28 a) of the first generator and on the outside of the hollow shaft (30) of the rotor (28 b) of the second generator are mounted.

14. Gas turbine according to one of the preceding claims, characterized in that by a control of the two generators, the rotational speeds of the two turbine wheels (18, 20) can be set arbitrarily, whereby the operating conditions of both systems can be optimized without adjustment.