EP1656497B1 - Diffuser located between a compressor and a combustion chamber of a gasturbine - Google Patents

Abstract

The invention relates to a gas turbine (1), comprising an annular combustion chamber (4) and an upstream diffuser (27), with a throughflow essentially parallel to a turbine longitudinal axis (9), at a distance from said axis at least partly less than the annular combustion chamber, in which a compressed gas (K) may be divided into several partial flows (Ki, Ka) at a branching point (36), whereby at least one of the partial flows (Ki, Ka) is a cooling gas flow. A main deflection region (30) is provided in said diffuser (27), directed at an angle to the turbine longitudinal axis (9) towards the annular combustion chamber (4).

Description

The invention relates to a gas turbine with an annular combustion chamber and one of these upstream, substantially parallel to a turbine longitudinal axis and flowed from this less than the annular combustion chamber spaced diffuser, in which a compressed gas at a branch point in partial flows can be divided.

Gas turbines are used in many areas to drive generators or work machines. The energy content of a fuel is used to generate a rotational movement of a turbine shaft. For this purpose, the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor. The working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work.

In the design of such gas turbines usually a particularly high efficiency is a design target in addition to the achievable performance and in addition to a compact design. An increase in the efficiency can be achieved for thermodynamic reasons basically by increasing the outlet temperature at which the working fluid from the combustion chamber and flows into the turbine unit. Therefore, temperatures of about 1200 ° C to 1300 ° C are sought for such gas turbines and achieved.

At such high temperatures of the working medium, however, exposed to this components and components are exposed to high thermal loads. However, to ensure a comparatively long life of the affected components with high reliability, is usually a Cooling of the affected components, in particular of running and / or vanes of the turbine unit, provided. Furthermore, it can be provided to cool the combustion chamber with a coolant, in particular cooling air.

From DE 195 44 927 A1 a gas turbine is known which has a combustion chamber upstream and opening into a diffuser air compressor. A partial flow of the compressed air can be branched out of the diffuser and used for cooling structural parts, for example turbine blades of the gas turbine. However, the Kühlluftabzweigung from the diffuser is only suitable for a branch of a relatively small partial flow from the air flow leaving the air compressor. On the other hand, the main air flow conducted through the diffuser is deflected in the diffuser in the direction of the combustion chamber and supplied to it as combustion air. A cooling of the downstream of the diffuser, that is, based on the flow direction of the working medium flowing through the turbine downstream components is thus limited possible.

Furthermore, DE 196 39 623 discloses a gas turbine with a diffuser, in which the removal of the cooling air takes place by means of a tube projecting into the outlet of the diffuser. The compressed air used for combustion in an annular combustion chamber is thereby diverted by means of a C-shaped plate in the direction of the burner. Both when removing the cooling air as well as in the leadership of the burner air flow losses can occur, which should be avoided.

The invention has for its object to provide a equipped with an annular combustor compact gas turbine, which allows a favorable flow guidance of the compressor air for a particularly uniform and effective cooling of thermally loaded components.

This object is achieved by a gas turbine with the features of claim 1. In this case, the gas turbine on an annular combustion chamber and an upstream of this annular diffuser, which is at least partially disposed between the turbine longitudinal axis and the annular combustion chamber. In the diffuser, which can be flowed essentially parallel to the turbine longitudinal axis, a compressed gas can be divided into a plurality of partial flows. According to the invention, the diffuser has a main deflection region, which is directed at an acute angle from the turbine longitudinal axis in a pioneering manner onto the inner wall of the annular combustion chamber. The main deflection region is followed in the direction of the gas flowing through the diffuser, in particular air, a branching point, at which the gas flowing through the diffuser can be divided into partial flows by means of a flow dividing element. The annular and in cross-section wedge-shaped flow dividing element is arranged between the two diverging walls of the diffuser - the radially inner inner wall and the radially outer wall lying outside. Two deflecting flanks opposite the walls of the diffuser converge towards each other at an acute angle and meet at the branching point. There they include an angle bisector, which intersects the turbine longitudinal axis at an acute pitch angle greater than 15 °.

The Hauptablenkbereich is seen in the axial direction behind the compressor and in front of the annular combustion chamber, whereas the flow dividing element between the annular combustion chamber and the turbine longitudinal axis is arranged. This geometry allows for the gas turbine a compact and in particular a shortened in the axial direction design. Furthermore, the flow losses in the compressed refrigerant partial streams are reduced.

Due to the guidance of the gas flow flowing through the diffuser with a component of the flow direction which is to be directed toward the annular combustion chamber, a particularly good coolability of components spaced radially from the turbine longitudinal axis, in particular the annular combustion chamber reached. Preferably, the two partial streams divided in the diffuser are also used in connection for combustion.

In an advantageous development, the outer wall of the diffuser and the outer deflecting flank of the flow-dividing element lying opposite to it extend approximately perpendicular to the turbine longitudinal axis behind the branching point. This ensures a low-loss supply of the outer partial flow to the outer flow passage space. A short and direct supply of the partial flow is achieved accordingly.

For gas turbines with a combustion chamber not designed as an annular combustion chamber, e.g. In gas turbines with so-called Can combustion chambers, the supply of the outer combustion chamber shell is quite simple. In gas turbines with Can combustion chambers, the individual flute-shaped combustion chambers are spaced apart on a ring concentrically enclosing the turbine longitudinal axis in the circumferential direction. The supply of cooling air to the radially outer combustion chamber shells can then take place between the individual Can combustion chambers.

Furthermore, a low-loss supply of the inner partial flow to the inner flow passage space is ensured by the inner wall of the diffuser and the opposite inner deflecting edge of the flow dividing element extends approximately parallel to the turbine longitudinal axis. From the compressor outlet to the flow transfer space, a wavy guide is proposed for the inner partial flow, which achieves an improvement over a linear guide in comparison to a straight guide with regard to the pressure losses and the flow losses in the partial flow.

According to a preferred embodiment, the compressed gas, which leaves the diffuser at this point, is conducted directly into a flow transfer space at the branch point, which passes the fluidic connection to the wall cooling space produces the annular combustion chamber. The flow transfer space preferably adjoins the outside of the combustion chamber wall, so that an additional cooling of the combustion chamber wall is achieved as a result.

The ring combustion chamber is preferably formed closed coolable. Here, combustion air is preferably performed as a cooling medium in countercurrent to the flue gas through a wall space of the annular combustion chamber. The combustion air flowing through the combustion chamber wall is preferably identical here, at least with a partial flow of the compressed air, which has previously flowed through the diffuser. Preferably, the air flowing through the diffuser is supplied completely to the wall of the annular combustion chamber as cooling air and further to the annular combustion chamber as combustion air. The division of the air flow at the branch point of the diffuser serves to provide several parts of the annular combustion chamber, such as an inner shell and an outer shell, evenly with cooling air.

If the annular combustion chamber has a combustion chamber rear wall which is essentially planar at least in a partial region, the wall angle of the annular combustion chamber is understood to mean that angle which the combustion chamber rear wall encloses with the turbine longitudinal axis. A particularly uniform all-round cooling of the combustion chamber wall is preferably achieved in that the pitch angle of the flow dividing element deviates from the wall angle of the combustion chamber rear wall by not more than 20 °, in particular by not more than 15 °.

Preferably, a pipe communicating with the lower part of the channel is provided for the removal of cooling air for the turbine. This allows a further division of the compressor air flow. If the tube protrudes into the lower part of the channel and faces with its pipe opening to the flow, the extraction of turbine cooling air is particularly favorable.

The advantage of the invention lies in the fact that in a gas turbine compressed air, which serves as a cooling and then combustion air is supplied with low pressure loss of an air compressor through a compact diffuser of the annular combustion chamber, wherein a flow divider at the outlet of the diffuser uniform cooling air to the annular combustion chamber causes.

FIG 1

einen Halbschnitt durch eine Gasturbine, und

FIG 2

im Querschnitt einen Diffusor und eine Ringbrennkammer der Gasturbine nach FIG 1.

An embodiment of the invention will be explained in more detail with reference to a drawing. Herein show:

FIG. 1

a half section through a gas turbine, and

FIG. 2

in cross-section a diffuser and an annular combustion chamber of the gas turbine according to FIG. 1

Corresponding parts are provided in both figures with the same reference numerals.

The gas turbine 1 according to FIG. 1 has a compressor 2 for combustion air, an annular combustion chamber 4 and a turbine 6 for driving the compressor 2 and a generator or a working machine (not shown). For this purpose, the turbine 6 and the compressor 2 are arranged on a common, also called turbine rotor turbine shaft 8, with which the generator or the working machine is connected, and which is rotatably mounted about its central axis 9.

The annular combustion chamber 4 is equipped with a number of burners 10 for the combustion of a liquid or gaseous fuel. It is also provided on its combustion chamber wall 23 with a wall lining 24.

The turbine 6 has a number of rotatable blades 12 connected to the turbine shaft 8. The blades 12 are arranged in a ring on the turbine shaft 8 and thus form a number of blade rows. Furthermore, the turbine 6 comprises a number of fixed vanes 14, which are also secured in a ring shape with the formation of vane rows on an inner housing 16 of the turbine 6. The blades 12 serve to drive the turbine shaft 8 by momentum transfer from the turbine 6 flowing through the flue gas or working medium M. The vanes 14, however, serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings. A successive pair of a ring of vanes 14 or a row of vanes and a ring of blades 12 or a blade row is also referred to as a turbine stage.

Each vane 14 has a platform 18, also referred to as a vane foot 19, which is intended to fix the respective vane 14 in the gas turbine 1. Each blade 12 is attached to the turbine shaft 8 in an analogous manner via a blade root 19, also referred to as a platform 18, the blade root 19 each carrying a profiled blade 20 extended along a blade axis.

Between the spaced-apart platforms 18 of the guide vanes 14 of two adjacent rows of guide vanes is in each case a guide ring 21 on the inner housing 16 of the turbine 6 is arranged. The outer surface of each guide ring 21 is also exposed to the hot, the turbine 6 flowing through the working medium M and spaced in the radial direction from the outer end 22 of the blade 12 opposite him through a gap. The guide rings 21 arranged between adjacent rows of guide blades serve in particular as cover elements which protect the inner wall 16 or other housing installation parts from thermal overload by the hot working medium M flowing through the turbine 6.

To achieve a comparatively high efficiency, the gas turbine 1 is designed for a comparatively high outlet temperature of the working medium M emerging from the annular combustion chamber 4 from about 1200 ° C. to 1300 ° C.

The combustion chamber wall 23 can be cooled with cooling air compressed in the compressor 2 as coolant K. Between the combustion chamber wall 23 and the wall lining 24, cooling air K flows in a wall space or wall lining space 26 in countercurrent to the working medium M onto the burner 10. The cooling air K, which also serves as combustion air, is passed from the compressor 2 through a diffuser 27 in the direction of the annular combustion chamber 4. By the diffuser 27, the cooling and combustion air K defined split on the one hand to an outer combustion chamber shell 28 and on the other hand, an inner combustion chamber shell 29 is supplied.

In FIG. 2, the flow guidance of the cooling air K through the diffuser 27 is shown in detail. The diffuser 27 has a main deflection region 30, which adjoins the compressor 2. The compressed cooling air K flows out of the compressor 2 parallel to the central axis or turbine longitudinal axis 9 and into the main deflection region 30 of the diffuser 27. The main deflecting region 30 of the diffuser 27 arranged axially between the compressor 2 and the annular combustion chamber 4 extends radially outwards under cross-sectional expansion, ie. away from the turbine longitudinal axis 9. As a result, the flow velocity of the compressed gas used as coolant K is reduced in the main deflection region 30. If there is a flow separation on the inner wall and outer wall of the diffuser 27, such a separation occurs only at low flow velocity and correspondingly low pressure loss.

At, based on the cooling air K, the downstream end 31 of the main deflection region 30, a flow dividing element 32 is disposed adjacent to the outer combustion chamber shell 29.

The arranged between the annular combustion chamber 4 and the turbine longitudinal axis 9 flow dividing element 32 has an approximately triangular in cross-section, also referred to as a dividing fork 33 shape with an outer Ablenkflanke 34 and an inner Ablenkflanke 35. The deflection flanks 34, 35 converge toward a division tip 36 directed towards the main deflection region 30 and enclose an acute angle of less than 90 °, in particular an angle of 60 °, in the division tip 36. The dividing point or edge 36 forming a branching point divides the cooling air K flowing through the main deflecting region 30 of the diffuser 27 approximately equally into an outer cooling air flow K _a and an inner cooling air flow K _i . The outer cooling air flow K _a is fed through an outer flow transfer chamber 37 of an outer combustion chamber shell 28, while the inner cooling air flow K _{i is} fed via an inner flow transfer chamber 38 of the inner combustion chamber shell 29.

The diffuser 27 dividing the cooling air K at the flow dividing element 32 is also referred to as a split diffuser. The cooling air K flowing through the main deflecting region 30 is directed approximately C-shaped radially, relative to the turbine longitudinal axis 9, outwardly to the dividing point 36 of the flow dividing element 32. A line extending as an angle bisector 39 between the curved Ablenkflanken 34,35 through the divisional peak 36 includes with the turbine longitudinal axis 9 a pitch angle α of about 45 °. With the lower combustion chamber shell 29, the bisector 39 includes an approximately right angle. The inner cooling air flow K _i is, starting from the division tip 36, forced by the inner Ablenkflanke 35 first in a horizontal flow direction, ie parallel to the turbine longitudinal axis 9 and further through the outside of the combustion chamber wall 23 radially inward, ie towards the turbine longitudinal axis 9, directed. The inner cooling air flow K _i is thus, initially still within the undivided in the main deflection region 30 cooling air K, guided in a roughly C-shaped curved path radially outward and thereby delayed and then guided in a direction in the opposite direction, approximately C-shaped curved path radially inward. Overall, the flow through the diffuser 27 and further into the internal flow transfer space 38 describes approximately a double S-shaped path. The radii of curvature within this path are large enough to cause only small energy losses in the flow.

Further, at the downstream end 31 of the diffuser 27, guide members or fixing members 41 are disposed both in the direction of the outer flow passage space 37 and the direction of the inner flow passage space 38.

The outer cooling air flow K _a is guided by the dividing fork 33 radially, perpendicular to the turbine longitudinal axis 9, to the outside. In the course of the outer cooling air flow K _{a is guided} past the outer combustion chamber shell 28 and introduced into the wall lining room or wall cooling space 26. Again, similar to the inner cooling air flow K _i results in a flow guidance with large deflection radii, with no sudden cross-sectional enlargements occur. Due to the cooling air streams or partial streams K _a , K _i , the combustion chamber shells 28, 29 are also cooled from the outside.

The burner 10 is arranged approximately centrally in a combustion chamber rear wall 42. A straight line passing through the combustion chamber rear wall 42 encloses the turbine longitudinal axis 9 with a wall angle β of approximately 45 °. The wall angle β thus corresponds approximately to the pitch angle α. The flow splitting element 32 arranged at an angle to the turbine longitudinal axis 9 at a pitch angle α splits the main deflecting region 30 into an upper sub-channel 43 and a lower sub-channel 44, which both have approximately the same cross-section. By laterally, ie along the inner combustion chamber shell 29 offset arrangement of the flow dividing element 32 is also a targeted asymmetrical division of the cooling air flow in the diffuser 27 feasible, if, for example, the outer combustion chamber shell and the inner combustion chamber shell 29 have a different cooling air requirement.

The removal for turbine cooling air is effected by a projecting into the lower part of the duct 44 tube 45. Whose end 46 is angled in the manner of a periscope and facing with its tube opening the inner air flow K _i , so that a portion of the air flow K _{i can} flow into the tube 45 as a turbine cooling air. The turbine cooling air flows at the other end of the tube 45 into an annular channel 47 extending along the rotor, which leads the turbine cooling air to the turbine 6. There, it is used for cooling the rotor blades and guide vanes 12, 14.

Claims (11)

1. Gas turbine (1) having an annular combustion chamber (4) which is inclined relative to the turbine longitudinal axis (9) and has a combustion chamber rear wall (42) in which a wall line runs which intersects the turbine longitudinal axis (9) at an acute wall angle β of at least 30°, having a compressor (2), downstream of which, in the axial direction, a diffuser (27) is fluidically connected which is arranged radially at least partly between annular combustion chamber (4) and turbine longitudinal axis (9), characterized in that a compressed gas (K) can be divided into partial flows (K_i, K_a) in the diffuser (27) at a branching point (36) by a wedge-shaped flow-dividing element (32) formed by two deflecting flanks (34, 35), the two deflecting flanks (34, 35) enclosing an angle of less than 90° at the branching point (36), and an angle bisector (39) enclosed between them intersecting the turbine longitudinal axis (9) at an acute dividing angle α greater than 15°, and the diffuser (27) having a main deflecting region (30) which is arranged upstream of the branching point (36) and which is directed at an acute angle pointing away from the turbine longitudinal axis (9) toward an inner combustion chamber shell (29), extending transversely to the combustion chamber rear wall (42), of the annular combustion chamber (4).
2. Gas turbine (1) according to Claim 1, in which the outer deflecting flank (34), defining the radially outer partial flow (K_a), and an outer wall, opposite this deflecting flank (34), of the diffuser (27) run behind the branching point (36) approximately perpendicularly to the turbine longitudinal axis (9).
3. Gas turbine (1) according to Claim 1 or 2, in which the inner deflecting flank (35), defining the radially inner partial flow (K_i), and an inner wall, opposite this deflecting flank (35), of the diffuser (27) run behind the branching point (36) approximately parallel to the turbine longitudinal axis (9).
4. Gas turbine (1) according to Claim 3, in which the radially inner partial flow (K_i) can be directed obliquely in the direction of the turbine longitudinal axis (9) after leaving the diffuser (27).
5. Gas turbine (1) according to one of Claims 1 to 4, having a wall cooling space (26), designed as inner combustion chamber shell (29) and as outer combustion chamber shell (28), of the annular combustion chamber (4).
6. Gas turbine (1) according to Claim 5, having a flow transfer space (37, 38) which adjoins the annular combustion chamber (4) and connects the diffuser (27) to the wall cooling space (26).
7. Gas turbine (1) according to one of Claims 1 to 6, having a closed cooled annular combustion chamber (4).
8. Gas turbine (1) according to one of Claims 1 to 7, in which the annular combustion chamber (4) is cooled by the counterflow method.
9. Gas turbine (1) according to one of Claims 1 to 8, in which the dividing angle α deviates from the wall angle β by not more than 20°.
10. Gas turbine (1) according to one of Claims 1 to 9, in which a tube (45) communicating with the bottom sectional passage (44) is provided in order to bleed cooling air for the turbine.
11. Gas turbine (1) according to Claim 10, in which the tube (45) projects into the bottom sectional passage (44), and its tube opening faces the flow.

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