TUBULAR COMBUSTION CHAMBER SYSTEM AND GAS TURBINE UNIT HAVING A TUBULAR COMBUSTION CHAMBER SYSTEM OF THIS TYPE

Abstract

A tubular combustion chamber system for a gas turbine unit includes a plurality of annularly arranged transition lines, which are designed to be connected at the upstream ends thereof to respective burners and to conduct hot gas produced by the burners to a turbine. The tubular combustion chamber system has a hot gas distributor, which is designed to be connected to the turbine and defines a ring channel, which is open to the turbine and into which the downstream ends of the transition lines lead. A gas turbine unit includes a plurality of annularly arranged burners, a turbine and a tubular combustion chamber system which connects the burners to the turbine.

Claims

1. A tubular combustion chamber system for a gas turbine unit, comprising: a plurality of annularly arranged transition ducts which are designed to be connected by their upstream ends in each case to a burner and to conduct hot gas produced by the burners to a turbine, and a hot gas manifold which is designed for connection to the turbine and which defines an annular channel, open to the turbine, into which there open the downstream ends of the transition ducts.

2. The tubular combustion chamber system as claimed in claim 1, wherein the transition ducts and the hot gas manifold are made of metal and are provided internally with a refractory lining.

3. The tubular combustion chamber system as claimed in claim 2, wherein a cross section of each transition duct tapers conically in a downstream direction, and wherein the refractory lining of the transition duct has at least one annular lining section whose outer diameter tapers conically in the downstream direction, which is held on the transition duct with radial and axial pretension.

4. The tubular combustion chamber system as claimed in claim 3, wherein the at least one annular lining section is formed by a single lining element.

5. The tubular combustion chamber system as claimed in claim 3, wherein the at least one annular lining section is formed by a plurality of ring segment-shaped lining elements which are braced against one another in the a circumferential direction.

6. The tubular combustion chamber system as claimed in claim 2, wherein the refractory lining of the hot gas manifold has a multiplicity of lining elements which are attached with radial pretension to the radially inner and outer faces of the hot gas manifold.

7. The tubular combustion chamber system as claimed in claim 2, wherein the transition ducts and the hot gas manifold are made of a high-heat-resistant metal material.

8. The tubular combustion chamber system as claimed in claim 2, wherein an outer circumferential side and/or an inner circumferential side of the hot gas manifold are/is provided with an attachment flange designed for attachment on the turbine.

9. A gas turbine unit comprising: a plurality of annularly arranged burners, a turbine, and a tubular combustion chamber system as claimed in claim 1 that connects the burners to the turbine.

10. The tubular combustion chamber system as claimed in claim 2, wherein the refractory lining comprises a ceramic lining.

11. The tubular combustion chamber system as claimed in claim 7, wherein the high-heat-resistant metal material comprises a thin-wall, high-heat-resistant metal material in the manner of a sheet.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Further features and advantages of the present invention will be apparent from the description below of a tubular combustion chamber system according to one embodiment of the present invention, with reference to the appended drawing, in which

[0017] FIG. 1 shows a perspective partial view, in partial section, of a tubular combustion chamber system according to one embodiment of the present invention, connected to a turbine of a gas turbine unit; and

[0018] FIG. 2 shows a perspective view of the arrangement represented in FIG. 1, viewed in the direction of the arrow II in FIG. 1.

DETAILED DESCRIPTION OF INVENTION

[0019] The figures show a tubular combustion chamber system 1 according to one embodiment of the present invention, connected to a turbine 2 of a gas turbine unit 3. The tubular combustion chamber system 1 comprises a plurality of annularly arranged transition ducts 4 which are designed to be connected by their upstream ends in each case to a burner 10 and to conduct hot gas produced by the burners 10 to the turbine 2; in FIG. 1, by way of example, only one individual burner 10 is shown. The tubular combustion chamber system 1 further comprises a hot gas manifold 5 which is designed for connection to the turbine 2 and which defines an annular channel 6, open to the turbine 2, into which there open the downstream ends of the transition ducts 4. The transition ducts 4 and the hot gas manifold 5 are made of metal, for example of a high-heat-resistant metal alloy. They each comprise a refractory lining 7, made advantageously of a ceramic material. The transition ducts 4 each have a cross section which tapers conically in the downstream direction. The refractory lining 7 of the transition ducts 4 comprises in each case a plurality of annular lining sections whose outer diameter tapers conically in the downstream direction, which presently are formed by annular lining elements 7a. Alternatively, however, it is also possible in principle for the annular lining sections to be formed in each case by a plurality of ring segment-shaped lining elements. The lining elements 7a of a transition duct 4 are inserted axially, starting from the upstream end of the transition duct 4, into the transition duct 4, with spring elements and/or damping elements, not shown in any more detail, being positioned along the circumference between the lining elements 7a and the inside wall of the transition duct 4, said elements being guided form-fittingly on the outer circumference of the lining elements 7a or on the inside wall of the transition duct 4. The conical configuration of the transition duct 4 and also of the lining elements 7a means that there is radial and also axial pretension of the lining elements 7a in such a way that they are held with radial and axial pretension on the transition duct 4. The tension is maintained presently by an annular pressure element 8 which is inserted into the transition duct 4 at the upstream end, is pressed against the end face of the adjacent lining element 7a, and then is attached to the transition duct 4 with generation of the desired pressing force. The attachment may be made, for example, by means of screws. The refractory lining 7 of the hot gas manifold 5 is realized by a multiplicity of lining elements 7b, which advantageously are attached likewise with radial pretension to the radially inner and outer faces of the hot gas manifold 5. To secure the tubular combustion chamber system 1 on the turbine 2, the outer circumferential side and the inner circumferential side of the hot gas manifold 5 are provided, on the free end of the hot gas manifold 5 facing the turbine 2, with attachment flanges 9 designed for attachment to the turbine 2 by means of screws.

[0020] The arrangement described above is advantageous in that, by virtue of the additional hot gas manifold 5 of the tubular combustion chamber system 1 according to the invention, the flow of hot gas impinging on the turbine 2 is very uniform, thus significantly reducing high-frequency changes in temperature and velocity. This is very beneficial for the lifetime of the turbine blades.

[0021] Further advantages are associated with the refractory lining 7 of the transition ducts 4 and of the hot gas manifold 5. This lining significantly reduces the thermal stress on the metallic components, i.e., the transition ducts 4 and the hot gas manifold 5. The smaller differences in expansion associated with this reduction, in the region of the seals to the turbine 2 and the seals between the transition ducts 4, result in less wear in this region and enable more robust assembly designs between the tubular combustion chamber system 1 and the turbine 2. Furthermore, the refractory lining 7 entails lower high-temperature requirements on the materials of the metallic components 4 and 5, thereby allowing cost savings to be made. By virtue of the lining 7, moreover, the transition ducts 4 can be implemented without an inner layer system, thereby significantly reducing the outlay for maintenance and reprocessing, since there is no need for stripping and recoating of the transition ducts 4. Furthermore, because of the use of a refractory lining 7, there is a reduction in the cooling demand of the metallic components 4 and 5 of the tubular combustion chamber system 1. In comparison to tubular combustion chamber systems without ceramic lining, the cooling air demand, according to present calculations, can be reduced by up to 50%, with a consequent increase in the performance of the gas turbine unit 3.

[0022] The invention, although having been described and illustrated in more detail through the exemplary embodiment, is nevertheless not limited by the examples disclosed, and other variations may be derived therefrom by the skilled person without departing the scope of protection of the invention.