CN111829013B - Transition duct for a gas turbine assembly and gas turbine assembly comprising such a transition duct - Google Patents

Abstract

The invention relates to a transition duct for a gas turbine assembly and a gas turbine assembly comprising the transition duct. Specifically, a transition duct for a gas turbine assembly and configured for channeling hot gases from a combustor to a turbine, the transition duct comprising: a tubular wall structure having an inner surface in contact with hot gas and an outer surface in contact with cooling air; a connection structure extending between the inner surface and the outer surface and configured to define a plurality of cooling channels extending through the tubular wall structure; wherein the connection structure is a flexible connection structure configured to allow circumferential movement of the outer surface relative to the inner surface.

Description

Transition duct for a gas turbine assembly and gas turbine assembly comprising such a transition duct

Cross Reference to Related Applications

The present patent application claims priority from european patent application No. 19170045.9 filed on date 2019, month 4 and 18, the entire disclosure of which is incorporated herein by reference.

Technical Field

The present invention relates to the technical field of gas turbine assemblies for power plants. In particular, the present invention relates to a particular component of a gas turbine assembly, namely a tube (referred to in the art as a "transition" tube) configured for directing hot gases exiting a combustor toward a turbine inlet. In more detail, the present invention relates to the tubular wall structure of the above-mentioned transition duct (known in the art as "liner").

Furthermore, the invention relates to a gas turbine assembly comprising the above-mentioned transition duct. In more detail, the gas turbine assembly of the present invention preferably relates to the technical field of so-called can-combustor gas turbines. Specifically, each can combustor includes: an upstream first combustor (referred to in the art as a "premix" combustor) configured to receive compressed air and mix the air with fuel; a downstream second combustor (referred to in the art as a "reheat" combustor) configured to receive the hot gas exiting the first combustor and to add fuel to the hot gas for spontaneous combustion/ignition, and a transition duct (as described above) for directing the hot gas exiting the reheat combustor to the turbine inlet. In this case, the bushing of the present invention may be defined as a "sequential" bushing. Of course, the present invention is not limited to the above-described burner having a reheat structure. Indeed, the present invention may also be applied without any modification to the non-reheat combustor.

Background

As is well known, a gas turbine assembly (hereinafter just gas turbine) for a power plant comprises a rotor having an axis (i.e. a gas turbine axis), a compressor, a combustor unit and at least one turbine. The compressor is configured to compress air supplied at the compressor inlet. Compressed air leaving the compressor flows into the plenum and from there into the combustor unit. The combustor unit includes a plurality of nozzles configured to inject fuel into the compressed air stream. The mixture of fuel and compressed air flows into a combustion chamber where it is combusted. The resulting hot gases leave the combustion chamber and expand in the turbine to perform work on the rotor. As is well known, turbines include multiple stages or rows of rotor blades inserted by multiple stages or rows of stator vanes. The rotor blades are connected to the rotor and the stator vanes are connected to a vane carrier, which is a concentric casing around the turbine unit.

To achieve high efficiency, high turbine inlet temperatures are required. Generally, however, such high temperatures involve undesirably high levels of NOx emissions. In order to reduce these emissions and increase operational flexibility without decreasing efficiency, so-called "sequential" gas turbines are particularly suitable. In general, a sequential gas turbine comprises two combustors or combustion stages in series, wherein each combustor is provided with a plurality of burners and at least one opposing combustion chamber. Along the main gas flow direction, the upstream or first burner typically comprises a plurality of so-called "premix" burners. The term "premix" emphasizes the fact that each burner of the first burner is not only configured for direct injection of fuel into the compressed air (e.g. with a so-called diffusion flame), but is also configured for mixing (with swirling) the compressed air and the fuel before injecting the mixture into the combustion chamber. The downstream or second combustor is referred to as the "reheat" or "sequential" combustor, and is fed by the hot gas exiting the first combustor. In addition, the reheat combustor is provided with a plurality of reheat burners configured for injecting fuel into the hot gas from the first combustor. Operating conditions downstream of the reheat burner allow the fuel/air mixture to spontaneously/spontaneously ignite due to the high gas temperature.

Today, two different types of sequential gas turbines are known. According to a first embodiment, the premix and reheat combustors are annular and physically separated by a turbine blade stage called a high pressure turbine. According to a second embodiment, the gas turbine is not provided with a high pressure turbine, and the burner unit is realized in the form of a plurality of can-combustors. In this embodiment, each can combustor includes a premix (first stage) and reheat (second stage) combustor disposed directly downstream of one another within a common can-shaped shell that terminates in a tubular element (referred to in the art as a "transition" tube) configured for directing hot gases exiting the reheat (second stage) combustor toward the turbine inlet.

As other components of the gas turbine engine come into contact with the hot gases, the transition duct needs to be cooled to avoid damage caused by overheating and in order to extend the service life. For the purpose of cooling the transition duct, a portion of the total airflow is typically extracted from the compressor (i.e., before the combustor) and used as convective cooling air acting on the outer surface of the transition duct. Today, there is a need to improve the cooling of the transition duct in order to allow further increases in the ignition temperature within the transition duct itself. A common solution to this problem is to use a higher amount of cooling air. However, the higher air consumption for cooling purposes reduces the efficiency of the gas turbine. The applicant proposes a solution consisting in providing the tubular wall of the transition pipe (i.e. the liner) with a plurality of cooling channels extending through the tubular wall itself between an upstream end and a downstream end, wherein these cooling channels are fed by cooling air in order to cool the liner in a convective manner. In view of the above-described structure with cooling channels, the tubular wall of the transition duct includes an inner wall portion (referred to in the art as a "hot shell"), an outer wall portion (referred to in the art as a "cold shell"), and a plurality of ribs that act as partitions of adjacent cooling channels and connect the inner wall portion and the outer wall portion. In view of the cross-section of the transition duct, these ribs may be defined as radial ribs, even though not the entire transition duct may disclose a circular cross-section. These ribs serve as a rigid structure, for which reason it is in fact possible to realise a thin hot shell. Furthermore, these radial ribs involve a higher creep resistance than the design of the single layer. The sequential liner (with cooling passages extending through the tubular wall separated by radial ribs) works well as long as the thermal barrier coating provided on the hot shell is not lost during operation of the gas turbine. In case of loss of the thermal barrier coating (or in case of a need for a very high temperature difference between the hot and cold shells), the static stress level in the wall structure increases due to the presence of these rigid connections (i.e. radial ribs).

Thus, there is a need today to improve the above-mentioned sequential bushings with cooling channels extending through the tubular wall in order to allow a further increase of the hot gas temperature and also to ensure safe operation in case of unexpected loss of the thermal barrier coating.

Disclosure of Invention

It is therefore a primary object of the present invention to provide a transition duct for a gas turbine configured for guiding hot gases from a combustor to a turbine, wherein the transition duct is further adapted to overcome the aforementioned prior art limitations. In particular, according to the invention, a transition tube is provided, wherein the transition tube comprises a tubular wall provided with a plurality of cooling channels, and wherein the transition tube allows a higher thermal gradient between a hot shell and a cold shell forming the wall. To achieve these results, according to the present invention, there is provided a transition duct for a gas turbine, wherein the transition duct comprises:

a tubular wall structure defining a path for the hot gas to leave the combustion chamber and guiding the hot gas towards the turbine (first turbine vane), wherein the tubular wall structure is provided with an inner surface (or inner wall portion) in contact with the hot gas and an outer surface (or outer wall portion) in contact with the cooling air (i.e. the portion of compressed air leaving the compressor);

-a connection structure extending within the tubular wall structure between the inner surface and the outer surface and configured for defining a plurality of cooling channels extending through the tubular wall structure.

Starting from this general configuration, according to a main aspect of the invention, the connection structure is a flexible connection structure configured for allowing circumferential movement of the outer surface (or outer wall portion) relative to the inner surface (or inner wall portion). The circumference is defined with respect to the longitudinal axis of the tube (main hot gas flow direction), which generally comprises at least a portion with a circular cross section. Due to the fact that the outer surface may grow in a circumferential direction (instead of the rigid structures in prior art practice), it is possible to increase the thermal mismatch between the outer surface and the inner surface without the risk of damaging the wall structure. The main features of the present invention have been defined above in a functional manner, as the skilled person can easily provide a number of different embodiments allowing the claimed result. In any event, in the description of the drawings, a number of different embodiments of the invention will be described. For example, this feature may be achieved by providing the tube with at least one groove extending from the cold shell towards the hot shell and/or providing an extension of the cold shell directed into the channel towards the hot shell or out of the channel. The elongated portion may preferably be V-shaped, U-shaped or half-moon shaped or any similar shape allowing for circumferential movement of the cold shell. The grooves and/or extensions may also be provided with through holes to allow cooling air to enter the channels.

Preferably, the transition duct of the present invention is a single piece made by an additive manufacturing method, such as a selective laser melting process.

According to a preferred embodiment, the connection structure may also be configured for controlling another movement of the wall structure. In particular, the connection structure may also be configured for substantially avoiding any radial movement of the inner surface or wall structure. Thus, according to this embodiment, radial movement of the structure is inhibited/prevented even in the event of creep. As previously mentioned, the radial direction is defined relative to the longitudinal axis (or main hot gas flow direction) of the tube, which generally comprises at least a portion having a circular cross-section.

Moreover, this additional feature of the connection structure has been defined above in a functional manner, as the skilled person can easily provide a number of different embodiments allowing this increased result. In any event, in the description of the figures, some different embodiments provided with this feature will be described. This feature is achieved, for example, by providing the tube with circumferential teeth marks that are received in a circumferential seat or extend as a lip over a portion of the cold shell.

Preferably, the new connection structure of the present invention does not affect the shape, geometry or configuration of the inner surface of the tube. Thus, embodiments of the present invention do not interfere with the thermal barrier coating applied to the inner surface.

The invention also relates to a gas turbine for a power plant, wherein the gas turbine has an axis and comprises:

-a compressor for compressing air;

-a burner section for mixing the compressed matter with at least one fuel and combusting the mixture;

at least one turbine for expanding the combustion hot gas flow leaving the burner section and applying work on the rotor.

Specifically, the burner section of the present invention comprises a plurality of can combustors, wherein each can combustor may comprise a single combustion stage or a first combustor, a second combustor in series. Each can combustor further comprises a transition duct realized in accordance with the appended claims for guiding hot gases from the combustor to the turbine.

It is to be understood that both the foregoing general description and the following detailed description are exemplary and are intended to provide further explanation of the invention as claimed. Other advantages and features of the present invention will become apparent from the following description, drawings, and claims.

The features of the invention believed to be novel are set forth with particularity in the appended claims.

Drawings

Other benefits and advantages of the invention will become apparent upon careful reading of the detailed description with appropriate reference to the accompanying drawings.

The invention itself, however, may be best understood by reference to the following detailed description of this invention, which describes an illustrative embodiment of the invention along with the accompanying figures in which:

FIG. 1 is a side elevation view, taken along an axial longitudinal plane, of a gas turbine assembly that may be provided with a transition duct according to the present invention;

fig. 2 is an enlarged view of the portion marked with reference number II in fig. 1;

fig. 3 and 4 are front and rear views, respectively, of a transition duct having a tubular wall provided with a plurality of cooling channels;

fig. 5A is an enlarged view of the portion marked with reference V in fig. 4, in particular fig. 5A discloses a transition duct wall structure according to prior art practice;

FIG. 5B is a graph showing the static stress load acting on the transition tube wall structure of FIG. 5A during operation due to thermal mismatch between the hot and cold shells;

fig. 6-11 are schematic illustrations of some different embodiments of transition duct wall structures according to the present invention.

Detailed Description

The technical contents and detailed description of the present invention are hereinafter described according to preferred embodiments with reference to the accompanying drawings, not for limiting the execution scope thereof. Any equivalent variations and modifications made in accordance with the appended claims are intended to be covered by the claims of this invention.

The present invention will now be described in detail with reference to the accompanying drawings.

Referring to fig. 1, fig. 1 is a side elevation view of a gas turbine assembly, taken along an axial longitudinal plane, which may be provided with a transition duct according to the present invention. Specifically, FIG. 1 discloses a simplified view of a gas turbine assembly, generally indicated by reference numeral 1. The gas turbine assembly 1 includes a compressor 2, a combustor assembly 3, and a turbine 4. The compressor 2 and the turbine 5 extend along a main axis a. The burner assembly 3 disclosed in the example of fig. 1 is a can burner 5, which may be a sequential burner or a single stage burner. Thus, in this embodiment, the burner assembly 3 comprises a plurality of sequential can-type burners 5 circumferentially arranged around the main axis a. As is well known, the compressor 2 of the gas turbine engine 1 provides a compressed air stream that is enriched with fuel and combusted in a can combustor 5. For cooling purposes, a portion of the air flow delivered by the compressor 2 is also supplied to the combustor assembly 3 and the turbine section 4.

Referring now to fig. 2, fig. 2 is an enlarged view of the portion of fig. 1 labeled with reference II. Specifically, FIG. 2 discloses a can combustor 5 of the gas turbine assembly 1 of FIG. 1. The can combustor 5 disclosed in fig. 2 comprises a first stage combustor 6 and a second stage combustor 7 and a transition duct 8, which are arranged in sequence and define a hot gas path. More precisely, the first stage burner 6 comprises a first stage burner unit 9 and a first stage combustion chamber 10. The second stage burner 7 is arranged downstream of the first stage burner 6 and comprises a second stage burner unit 17 and a second stage combustion chamber 18. The second stage combustor 7 is also coupled to the turbine 4 by a transition duct 8, not shown here. The second stage combustor 12 extends in the axial direction downstream of the first stage combustor 6. In this embodiment, the second stage combustor 12 includes an outer liner 13 and an inner liner 14, wherein the outer liner 13 surrounds the inner liner 14 a distance therefrom such that a convective cooling channel 15 is defined between the outer liner 13 and the inner liner 14.

Reference is now made to fig. 3 and 4, which are front and rear views, respectively, of the transition duct 8 of fig. 2. As is well known, the disclosed transition duct 8 is a component of the hot gas path that is subject to the most severe thermal stresses. For cooling purposes, the transition duct 8 therefore comprises a tubular wall structure 17 provided with a plurality of cooling channels 16 supplied by cooling air. The tubular wall structure 17 has: an upstream end 18 having a circular cross-section, a downstream end 19 having a substantially rectangular cross-section, an inner surface 20 and an outer surface 21. The upstream end 18 is joined to the second stage combustor 12, while the downstream end 19 faces the turbine 4. The inner surface 20 defines a hot gas flow volume through which hot gas flows to the turbine 4. Thus, the inner surface 20 is directly exposed to the hot gas flowing through the hot gas path. In view of the above, the inner surface may be at least partially coated with a thermal barrier coating. The cooling channels 16 extend through the tubular wall structure 17 between the upstream end 18 and the downstream end 19 and are evenly distributed in the circumferential direction of the tubular wall structure 17. At the upstream end 18, the cooling channels 16 are in fluid communication with the convective cooling channels 15 of the second stage combustor 12. In this embodiment, the cooling channels 16 extend in the axial longitudinal direction of the tubular wall structure 17.

Referring now to fig. 5A, fig. 5A is an enlarged view of the portion of fig. 4 labeled with reference numeral V. This fig. 5A discloses a wall 17 (rigid connection) according to prior art practice. As disclosed in fig. 5A, the cooling channels 16 separate different portions of the tubular wall structure 17. In particular, it is possible to define an inner wall portion 22 or hot shell of the tubular wall structure 17, which is a portion of the wall between the inner surface 20 and the cooling channel 16. Thus, the outer wall portion 23 or cold shell of the tubular wall structure 17 may be defined as the portion of the wall between the outer surface 21 and the cooling channel 16. Adjacent cooling channels 16 are separated by a partition or partition in the form of radial ribs 24, or radial ribs 24 extend between the inner wall portion 22 and the outer wall portion 23 of the tubular wall structure 17. During operation, the temperature of the hot shell (inner wall portion 22) increases by several 100K relative to the temperature of the cold shell (outer wall portion 23). Of course, any further increase in the heat load or loss of the thermal barrier coating applied to the hot shell results in a corresponding increase in this temperature differential. The above temperature differences are related to the high static stress loads acting on the wall structure due to the rigid connection of the prior art provided by the ribs extending between the hot and cold shells. In particular, since the hot shell expands more in the circumferential direction than the cold shell, the cold shell must resist high static stress loads acting in the circumferential direction. Next to fig. 5A, direction references A, R and C are disclosed, wherein direction a is the axial direction (main hot gas flow direction), direction C is the circumferential direction, and direction R is the radial direction. The graph of fig. 5B schematically discloses the static stress load acting on the wall structure.

As described in the section on the general definition of the invention, in order to allow an increase in the thermal load and also to ensure safe operation in the event of loss of the above-mentioned thermal barrier coating, the proposed solution consists in providing the wall 17 with flexible features configured for allowing the cold shell 23 to grow in the circumferential direction. In this way, the stress level acting on the cold shell due to the thermal mismatch is reduced and compensated by this circumferential movement. The following description of fig. 6-11 will mention some specific embodiments of the invention, i.e. embodiments in which the rigid prior art connection (radial ribs) between the hot shell and the cold shell has been modified to allow free growth of the cold shell in the circumferential direction. Regardless, in all disclosed embodiments, radial movement of the hot shell is inhibited/prevented even in the event of creep.

Referring now to fig. 6, fig. 6 is a schematic view of a first embodiment of a transition duct wall structure according to the present invention. According to this embodiment, the transition tube wall structure 17 comprises at least one radial groove or slot 25 (extending at least partially axially along the tube), which radial groove or slot 25 is realized in the rib 24 and extends from the cold shell 23 towards the hot shell 22. Thus, at this radial groove 25, the thickness of the transition tube wall structure 17 is reduced to a separate heat shell 22. In other words, at this radial groove or slot 25 in the circumferential direction, the integrity of the transition duct wall structure 17 is only performed at the point where the hot shell 22 acts as a hinge 28 for this structure. In contrast, at radial groove or slot 25, cold shell 23 discloses two spaced facing edges to allow the cold shell to move in circumferential direction C. In order to avoid movement of the structure in the radial direction R (in particular to prevent any radial movement of the hot shell), the intermediate portion of the groove or channel 25 in the radial direction is provided with a circumferential tooth trace 26 at one side and a circumferential seat 27 at the opposite side, at least partially accommodating the circumferential tooth trace 26.

Referring now to fig. 7, fig. 7 is a schematic view of an alternative embodiment of a transition duct wall structure in accordance with the present invention. This embodiment differs from the previous embodiments in the location of the circumferential teeth 26 along the radial direction of the groove or flute 25. As disclosed, in fig. 7, the tooth trace 26 extends outside the groove or recess 25, in particular radially covering (spacing) the groove or recess 25 outside the cold shell 23. Also in this case, the tooth marks 26 allow preventing any radial movement of the hot shell.

Referring now to fig. 8, fig. 8 is a schematic view of an alternative embodiment of a transition duct wall structure in accordance with the present invention. According to this embodiment, the transition tube wall structure 17 comprises at least one V-shaped groove or lowered portion 29 of the cold shell 23 at the channel 16, wherein the V-shaped portion is directed towards the hot shell 22. A bridge 30 may also be provided to connect the apex of the V-shaped portion 29 to the heat shell 22. At the V-shaped portion 29, the rigidity of the cold shell is reduced to allow the cold shell itself to move in the circumferential direction.

Referring now to fig. 9, fig. 9 is a schematic view of an alternative embodiment of a transition duct wall structure in accordance with the present invention. This embodiment differs from the previous embodiment in that the V-shaped portion 29' does not extend towards the hot shell 23 but extends outside the channel 16. Also in this case, at the V-shaped portion 29', the rigidity of the cold shell 23 is reduced to allow the cold shell itself to move in the circumferential direction.

Referring now to fig. 10, fig. 10 is a schematic view of an alternative embodiment of a transition duct wall structure in accordance with the present invention. This embodiment includes all the features of the embodiment of fig. 6, and in addition, fig. 10 discloses at least one cooling hole 30 connecting the groove 25 to the adjacent channel 16. In order to maximize the cooling effect of the hot shell 22, cooling holes 30 are realized at the end of the groove 25 opposite to the cold shell 23. Arrows F in fig. 10 represent the cooling flow through the grooves 25 and holes 30 into the channels 16.

Reference is now made to fig. 11, which is a schematic illustration of an alternative embodiment of a transition duct wall structure in accordance with the present invention. This embodiment is very similar to the embodiment disclosed in fig. 7. However, the embodiment of fig. 11 includes cooling holes 30 (as in the previous example) for connecting the grooves 25 to the channels 16, and it is not provided with teeth marks 26 covering the grooves 25. In practice, the teeth 26 may limit the air flow into the grooves 25.

While the invention has been explained with respect to the preferred embodiments thereof as mentioned above, it will be understood that many other possible modifications and variations may be made without departing from the scope of the invention. It is therefore contemplated that the appended claims will cover such modifications and variations as fall within the true scope of the invention.

Claims (14)

1. A transition duct (8) for a gas turbine assembly (1) and configured for guiding hot gases from a combustion chamber to a turbine, the transition duct (8) comprising:

-a tubular wall structure (17) provided with an inner surface (20) in contact with hot gas and an outer surface (21) in contact with cooling air;

-a connection structure extending between the inner surface (20) and the outer surface (21) and configured for defining a plurality of cooling channels (16) extending through the tubular wall structure (17);

the connection structure is a flexible connection structure comprising a plurality of ribs (24) extending from the outer surface (21) to the inner surface (20) for separating adjacent cooling channels (16),

it is characterized in that the method comprises the steps of,

the transition tube (8) comprises at least one groove (25) extending from the outer surface (21) towards the inner surface (20) within a rib (24).

2. The transition pipe (8) according to claim 1, wherein the transition pipe (8) is a single piece.

3. The transition pipe (8) according to claim 2, wherein the transition pipe (8) is made by an additive manufacturing method.

4. The transition pipe (8) according to any one of the preceding claims, wherein the connection structure is further configured for avoiding radial movement of the inner surface (20).

5. A transition duct (8) according to any one of claims 1-3, wherein the connection structure does not affect the inner surface (20).

6. Transition pipe (8) according to claim 1, wherein the intermediate portion of the groove (25) is provided with a tooth trace (26) configured for avoiding radial movement of the inner surface (20).

7. Transition pipe (8) according to claim 1, wherein the outer surface (21) is provided with a lip covering the groove (25) on the outside.

8. The transition pipe (8) according to claim 1, wherein the outer surface (21) comprises an elongated portion (29') extending towards the inner surface (20) within the cooling channel (16).

9. Transition pipe (8) according to claim 8, wherein a bridge connecting the elongated portion (29') with the inner surface (20) is provided.

10. The transition pipe (8) according to claim 1, wherein the outer surface (21) comprises an elongated portion (29') extending outside the cooling channel (16).

11. Transition pipe (8) according to claim 1, wherein the groove (25) comprises at least one hole (30) connecting the groove (25) to a cooling channel (16).

12. The transition pipe (8) according to any one of the preceding claims 8 to 10, wherein the elongated portion comprises at least one hole (30) connecting the elongated portion to a cooling channel (16).

13. The transition pipe (8) according to claim 11, wherein the hole (30) is provided at an end of the groove (25) opposite to the outer surface (21).

14. A gas turbine for a power plant; the gas turbine (1) has an axis (a) and comprises, in the direction of the air flow:

a compressor (2) for compressing air,

a burner section (3) for mixing and burning the compressed material with at least one fuel,

-a turbine (4) for expanding a flow of combustion hot gases exiting the burner section (3);

wherein the burner section (3) comprises a plurality of can-type burners (5), each of which accommodates a single burner or a first burner (6), a second burner (7) and a transition pipe (8) in series, the transition pipe (8) being realized in accordance with any one of the preceding claims.