EP2199681A1 - Gas turbine combustion chamber and gas turbine - Google Patents

Abstract

The gas turbine combustion chamber (25) comprises a rotationally symmetrical cross section which has an inner wall and a spaced external wall (37). The inner wall has an external side, where the external wall has an inner side facing the external side of the inner wall, so that a gap is provided for forming a coolant channel between the external side of the inner wall and the inner side of the external wall. The cooling ribs (43) are wound around the inner wall in a helix-shape. An independent claim is also included for a gas turbine with a rotor.

Description

The present invention relates to a gas turbine combustor having a substantially rotationally symmetrical cross section and at least one axial section, which has an inner wall with an outer side and an outer wall with an inner side facing away from the inner wall and spaced from the inner side, so that between the outer side and the Inside a at least one cooling fluid channel forming space is present. In addition, the present invention relates to a gas turbine.

A gas turbine comprises as essential components a compressor, a turbine with blades and guide vanes and at least one combustion chamber. The blades of the turbine are arranged on a shaft extending mostly through the entire gas turbine, which is coupled to a consumer, such as a generator for power generation. The shaft provided with the blades is also called turbine runner or rotor

During operation of the gas turbine, the combustion chamber is supplied with compressed air from the compressor. The compressed air is mixed with a fuel, such as oil or gas, and burned in the combustion chamber. The hot combustion gases are finally fed via a combustion chamber outlet of the turbine, where they transmit momentum to the blades under relaxation and cooling and thus do work.

In particular, the combustion chambers of so-called diffusion combustion systems, in which a fuel-rich fuel-air mixture is burned, are exposed to very high temperatures during operation of the gas turbine. The combustion chamber is in this case a mechanical container, which serves to stabilize the flame and to ensure the transfer of heated by the combustion compressor air in the turbine. Since this mechanical container is located near the flame, it is exposed to temperatures that exceed even the melting temperature of superalloys. Therefore, in order to prevent the combustion chambers from melting, they are often equipped with complex double-walled cooling systems and cooling fins between the walls. A combustion chamber for a diffusion flame, which has a double wall, is, for example, in WO 99/17057 A1 described. Likewise, a double-walled combustion chamber, but which is used in conjunction with a premix flame, ie a vortexed air-fuel mixture, is disclosed in US Pat WO 97/14875 A1 described. With increasing temperature, however, the effective flow cross-section of the cooling fluid channel decreases with such a configured cooling system. Due to the higher elongation of the inner wall when it becomes warmer during operation, the distance between the inner wall and the outer wall and thus the cross-sectional area of the cooling air passage decreases. The change in the cross-sectional area also makes it difficult to control the amount of cooling fluid flowing through the cooling fluid passage. Simple closing and opening of cooling fluid inlet openings is not possible for controlling the amount of cooling fluid, since this is mainly determined by the cross-sectional area of the cooling fluid channel.

Compared to this prior art, it is an object of the present invention to provide an advantageous gas turbine combustor available. It is another object of the present invention to provide an advantageous gas turbine.

The first object is achieved by a gas turbine combustor according to claim 1, the second object by a gas turbine according to claim 7. The appended claims contain advantageous embodiments of the invention.

A gas turbine combustor according to the invention has a substantially rotationally symmetrical cross-section and has at least one axial section, which is an inner Wall having an outer side and an outer wall having an outer side of the inner wall facing and spaced from the inner side. Thus, between the outside of the inner wall and the inside of the outer wall there is a gap forming at least one cooling fluid channel. The outside of the inner wall has cooling fins projecting toward the inside of the outer wall. In addition, the outer wall to the space leading inlet openings for a cooling fluid. In the gas turbine combustor according to the invention, the cooling fins are helically wound around the inner wall.

In comparison to axially extending cooling fins, as used in the prior art, the helically wound cooling fins have several advantages. Thus, the length of a helically wound cooling fin with the same axial length of the inner wall is longer than a linearly extending in the axial direction of the cooling fin, whereby the surface for heat transfer to the flowing along the cooling fins thermal fluid is increased compared to the axial rib. Better heat transfer and thus more efficient cooling is the result. In addition, the cooling fluid is also passed through the cooling fins, so that this flows along a helically curved flow path. It remains due to the greater length of a helical flow path longer in contact with the inner wall to be cooled, whereby the cooling fluid can absorb heat for a longer time and is thus used more efficiently.

A disadvantage of using the conventional cooling fins which are conventional in the prior art is that the inner wall of the combustion chamber experiences a significant radial thermal expansion due to the very high temperatures prevailing during operation. This thermal expansion is large enough to reduce the flow area through the gap between an inner wall with a straight fin and the outer wall. The reason for this is that the straight cooling fins do not cause rigidity of the inner wall, which would oppose a radial expansion of the inner wall. Helical ribs, on the other hand, give the inner wall a rigidity which prevents radial expansion. The thermal expansion of an inner wall provided with helical cooling ribs leads to this rather to an axial expansion and to a rotation, but hardly to a radial expansion. The radial expansion due to the prevailing high temperatures is therefore significantly reduced in comparison to combustion chambers according to the prior art, ie a combustion chamber with linearly extending cooling ribs in the axial direction. The axial thermal expansions as well as a rotation of the inner combustion chamber wall hardly influence the flow cross section of the gap between the inner wall and the outer wall, so that better control and better uniformity of the cooling fluid flow at different temperatures prevailing in the combustion chamber is possible.

Finally, helically wound cooling fins also offer an advantage in terms of production since, in contrast to straight cooling fins, they can be introduced into the inner wall in a turning process. In contrast, axial cooling ribs have to be milled in, which represents an increased outlay compared with the production of the helical cooling ribs.

In the gas turbine combustor according to the invention, the inner wall may in particular have a downstream end at which the intermediate space between the outer side of the inner wall and the inner side of the outer wall is open toward the interior of the combustion chamber. The cooling fluid can then be supplied to the combustion chamber interior, which is used in particular for diffusion flames. As a cooling fluid in this case, for example, compressor air or steam can be used. In this embodiment, the helical winding of the cooling fins causes the cooling fluid to be conducted on a helical path along the outside of the inner wall. The cooling fluid therefore occurs in a vortex that the Flame stabilized, into the combustion chamber interior. In addition, the vortex keeps the cooling fluid due to the centripetal force in the vicinity of the wall, which brings advantages in terms of cooling with it.

In an advantageous embodiment of the gas turbine combustor according to the invention, the outer wall has steps in the axial direction of the gas turbine combustor. There are a number of axially-spaced inner walls, with the inner walls being annular and increasing in diameter with axially-spaced annular inner walls. Adjacent annular inner walls are in this case partially pushed together. This embodiment makes it possible to design the cooling fluid passages between the different inner walls and the outer wall differently, be it with different flow cross sections and / or with different slopes of the helically wound ribs and / or with different rib geometries.

Furthermore, in the advantageous embodiment, each of the outer walls of the inner walls, which are partially pushed into one another, may be fastened to a fastening section of the outer wall in their section surrounding the inner of the inner walls pushed one inside the other. The inlet openings of the outer wall then adjoin these fastening sections. Each intermediate space formed between an inner wall and the outer wall can then be supplied with cooling fluid individually. In particular, in this case, each of the axially arranged in succession inner walls have a downstream end on which the existing between the outside of the respective inner wall and the inside of the outer wall gap to the combustion chamber interior is open. In this way, a further inner wall arranged in the axial direction behind an inner wall is further cooled by means of film cooling by the cooling fluid entering the combustion chamber, which flows along the inside of the following inner wall. The film cooling is done by the vortex of the entering into the combustion chamber cooling fluid also stabilized. The cooling fluid thereby remains longer on the inside of the subsequent inner wall than in the linear cooling fins used in the prior art. The efficiency of the cooling can be increased thereby.

A gas turbine according to the invention is equipped with at least one combustion chamber according to the invention. In particular, a plurality of combustion chambers according to the invention, for example six, eight or twelve combustion chambers, may be arranged around the rotor. The advantages described with reference to the gas turbine combustor according to the invention also result in the gas turbine according to the invention. Reference is therefore made to the advantages described with reference to the gas turbine combustor according to the invention.

• Figur 1 zeigt eine erfindungsgemäße Gasturbine in einem Längsteilschnitt.
• Figur 2 zeigt eine erfindungsgemäße Brennkammer in einem Längsschnitt.
• Figur 3 zeigt in einer stark schematisierten Darstellung die innere Wand einer erfindungsgemäßen Gasturbinenbrennkammer.
• Figur 4 zeigt einen vergrößerten Ausschnitt aus Figur 3.

Further features, properties and advantages of the present invention will become apparent from the following description of an embodiment with reference to the accompanying figures.

• FIG. 1 shows a gas turbine according to the invention in a longitudinal partial section.
• FIG. 2 shows a combustion chamber according to the invention in a longitudinal section.
• FIG. 3 shows in a highly schematic representation of the inner wall of a gas turbine combustor according to the invention.
• FIG. 4 shows an enlarged section FIG. 3 ,

FIG. 1 shows a gas turbine 1 in a longitudinal section. This includes a compressor section 3, a combustor section 5 and a turbine section 7. A shaft extends through all sections of the gas turbine 1. In the compressor section 3, the shaft 9 is provided with rings of compressor blades 11 and in the turbine section 7 with rings of turbine blades 13 equipped. Wreaths of compressor vanes 15 are located in the compressor section 3 between the rotor blade rings and rings of turbine vanes 17 in the turbine section 7. The vanes extend from the housing 19 of the gas turbine installation 1 essentially in the radial direction to the shaft.

During operation of the gas turbine 1, air is sucked in through an air inlet 21 of the compressor section 3 and compressed by the compressor blades 11. The compressed air is supplied in the combustion chamber section 5 arranged combustion chambers 25, in which a gaseous or liquid fuel via at least one burner 27 is injected. The resulting air-fuel mixture is ignited and burned in the combustion chambers 25. Along the flow path 29, the hot combustion exhaust gases flow from the combustor 25 into the turbine section 7, where they expand and cool, imparting momentum to the turbine blades 13. The turbine guide vanes 17 serve as nozzles for optimizing the momentum transfer to the rotor blades 13. The rotation of the shaft 9 caused by the momentum transfer is used to drive a load such as an electric generator. The expanded and cooled combustion gases are finally discharged from the gas turbine 1 through an outlet 23.

FIG. 2 shows a combustion chamber 25 of the gas turbine 1 in a schematic sectional view. The combustion chamber 25 includes a burner end 31 to which at least one burner 27 is disposed and through which both the fuel and compressor air are introduced into the combustion chamber. In addition, the combustion chamber 25 comprises a turbine-side outlet end 33, through which the hot combustion exhaust gases exit the combustion chamber 25 in the direction of the turbine section 7. The existing during operation of the gas turbine 1 in the combustion chamber 25 flame leads in a section 35 of the combustion chamber to very high temperatures, which make cooling of the combustion chamber wall necessary, especially when the flame a Diffusion flame is. In order to enable cooling, the combustion chamber wall, at least in this section 35, has a double-walled structure with an outer wall 37 and one or more inner walls 39A, 39B, 39C. Between the inner walls 39A, 39B, 39C and the outer wall 37 there are intermediate spaces 41A, 41B, 41C which form cooling fluid passages for a cooling fluid, in the present embodiment compressor air. The outer side, the substantially cylindrical inner walls 39A, 39B, 39C have helically wound ribs 43 which project towards the outer wall 37 and between which helical flow paths for the compressor air are formed. The geometry of these flow paths is in FIG. 3 which shows a schematic view of the outside of an inner wall 39. FIG. 4 shows an enlarged section FIG. 3 ,

The inner walls 39 each have a mounting portion 45, in which they are attached to a mounting portion 46 of the outer wall 37. The inner walls 39 have slightly different radii, wherein the radii in the flow direction 47 of the combustion gases increase. The fastening portions 45 remote from the ends 40 of the inner walls 39 are inserted into a part in the downstream adjacent inner wall 39. In this case, a distance between the outside of the inner inner wall (eg 39A) and the inner side of the outer Innwand (eg 39B) or the outer wall 37 remains such that on the outflow side an annular opening 42 open towards the combustion chamber interior is formed. The outer wall 37 has, in the vicinity of the fixing portions 46 to which the inner walls 39 are fixed with their fixing portions 45, through holes 49 serving as inlet openings for compressor air into the spaces 41. The compressor air then flows along the outside of the inner walls 39 to cool them. Finally, the compressor air flows through the annular opening 42 into the combustion chamber interior. Due to the helically wound flow paths between the ribs 43, the cooling air entering the combustion chamber interior forms a vortex which runs along the inside of the downstream side flows along the following inner wall 39 and thus serves as a film cooling for this inner wall 39.

In the present case, therefore, the cooling air serves as cooling air in two ways, namely firstly by cooling first the outside of an inner wall, and then the inside of the following inner wall. In the first cooling process, the helically wound cooling fins 43 on the outer sides of the inner walls 39 make the path along which the cooling air flows on the outer side of the inner walls 39 longer, as compared to straight fins, allowing a higher heat transfer to the flowing air is. The helically wound shape of the cooling ribs 43 also results in the inner walls 39 being given a higher rigidity than radial thermal expansion, so that the flow cross section between the inner side of the outer wall 37 and the outer side of the inner walls 39 remains substantially constant even at different temperatures during operation of the gas turbine remains. Instead of a radial thermal expansion, as would occur in the axial direction of the inner walls extending linear ribs occurs in the helically wound ribs 43, an axial expansion in conjunction with a twisting of the inner wall 39. The shape of the ribs 43, for example, the pitch of the helical winding, its height, its width, etc., here can, for example, be chosen so that the radial thermal expansion is minimized or so that the time required for the cooling air to to flow along the cooling fins 43, is set to a certain duration.

Furthermore, it is possible to optimize the geometry of the cooling fins 43 with regard to the vortex occurring after the entry of the cooling air into the combustion chamber interior. It should also be noted at this point that the in FIG. 4 Although shown cooling fins have a rectangular cross-section, but these may have other cross sections, such as a triangular cross-section or a trapezoidal cross-section. In addition, the ribs at different inner walls 39A, 39B, 39C be designed differently in order to optimize the cooling effect on the conditions prevailing in the respective combustion chamber section conditions.

The helical cooling ribs 43 can be produced in a simple manner by introducing helical grooves into a cylindrical inner wall by means of a turning process.

Claims (7)

A gas turbine combustor (25) of substantially rotationally symmetrical cross-section having an inner wall (39) with an outer side and an outer wall (37) with an inner side facing and spaced from the outer side of the inner wall (39) such that between the outer side the inner wall (39) and the inner side of the outer wall (37) is provided with a gap (41) forming at least one cooling fluid channel, the outer side of the inner wall (39) projecting cooling fins (37) projecting towards the inner side of the outer wall (37). 43) and wherein the outer wall (37) to the intermediate space (41) having leading inlet openings (49) for a cooling fluid,
characterized in that
the cooling fins (43) are helically wound around the inner wall (39). Gas turbine combustor (25) according to claim 1,
characterized in that
the inner wall (39) has a downstream end at which the intermediate space (41) between the outer side of the inner wall (39) and the inner side of the outer wall (37) is open towards the interior of the combustion chamber. A gas turbine combustor (25) according to claim 1 or claim 2,
characterized in that
the outer wall (37) has steps in the axial direction of the gas turbine combustor (25) and a number of axially spaced inner walls (39A, 39B, 39C) are provided, the inner walls (39A, 39B, 39C) being annular the diameters of axially-spaced annular inner walls (39A, 39B, 39C) increase, and adjacent annular inner walls (39A, 39B, 39C) are partially telescoped. Gas turbine combustor (25) according to claim 3,
characterized in that - The respective outer of the partially telescoped inner walls (39 B, 39 C) in its the inner of the telescoped inner walls (39 A, 39 B) surrounding portion is fixed to a corresponding mounting portion (46) of the outer wall (37) and - The inlet openings (49) for the cooling fluid to the mounting portions (46) of the outer wall (37) adjoin. Gas turbine combustor according to claim 2 and claim 4,
characterized in that
each of the axially spaced, inner walls (39) has a downstream end (40) at which the gap (41) between the outside of the respective inner wall (39) and the inside of the outer wall (37) opens to the interior of the combustion chamber is. Gas turbine with at least one combustion chamber (25) according to one of the preceding claims. A gas turbine according to claim 6, wherein a plurality of combustion chambers (25) according to any one of claims 1 to 5 are arranged around the rotor (9) of the gas turbine.

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