Gas turbine assembly for a power plant and method for maintaining a gas turbine assembly

Abstract

A gas turbine for a power plant; the gas turbine includes: at least one turbine provided with at least one row of stator vanes configured to direct a hot gas flow, wherein each vane is outwardly connected to a housing and includes an inner platform; a carrier arranged inwardly relative to the inner platform of the vane; a film seal extending from the carrier to the inner platform of the vane to isolate a lower pressure hot gas from a volume located below the inner platform of the vane, downstream of the film seal, and supplied with a higher pressure cooling air; wherein the film seal includes: an inner end received in a groove in the carrier; an outer end received in a groove in the inner platform of the vane; a film connecting the inner end and the outer end; wherein the film is offset towards the higher pressure air volume at the outer end so that, in operating conditions, the upstream surface of the vane groove does not contact the film.

Description

[0001] Cross-reference to related applications

[0002] This patent application claims the priority of European Patent Application No. 19209597.4 filed on November 15, 2019, the entire disclosure of which is incorporated herein by reference. Technical field

[0003] The present invention relates to the technical field of gas turbine assemblies for power plants. As is known, in a gas turbine assembly (or simply referred to as a gas turbine), an incoming air flow is compressed and then, in a combustor, ignited together with added fuel to produce a hot gas flow that will expand in the turbine to produce rotational work on a rotor, which in turn is connected to a generator. Due to the very high temperature of the working hot gas, some components of the gas turbine assembly must be cooled, and thus, in addition to the main hot gas flow, a cooling flow (usually cooling air) also circulates inside the gas turbine. In addition, some components of the gas turbine are configured to allow a certain relative movement depending on the operating conditions. In view of the above, within the field of gas turbine assemblies for power plants, the present invention specifically relates to the technical problem of how to seal the gas turbine interfaces to separate the cooling air flow from the hot gas flow, where this seal must be ensured during all operating or working conditions. Background art

[0004] As is known, a gas turbine assembly for a power plant includes a compressor assembly, a combustor assembly, and a turbine assembly. The compressor assembly is configured to compress the incoming air supplied at the compressor inlet. The compressed air leaving the compressor assembly flows into a volume (referred to as a "plenum") and from there into the combustor assembly. The combustor assembly typically includes a plurality of burners configured to inject fuel (at least one type of fuel) into the compressed air flow. The mixture of fuel and compressed air enters a combustion chamber where the mixture is ignited. The resulting hot gas flow leaves the combustion chamber and, by passing through the turbine assembly, does rotational work on a rotor connected to a generator. As is known, the turbine assembly includes multiple stages or rows of rotating blades that are interleaved with multiple stages or rows of stator guide vanes. The rotating blades are supported by the rotor, and the stator guide vanes are supported by a concentric housing (referred to as a "vane carrier") that surrounds the turbine assembly.

[0005] For high efficiency, the hot gas flow must have a very high turbine inlet temperature. However, generally, this high temperature involves undesirably high NOx emission levels. To reduce this emission and increase operational flexibility without reducing efficiency, the so-called "sequential" gas turbines are particularly suitable. Generally, a sequential gas turbine includes a first burner or combustion stage and a second burner or combustion stage, where each burner is provided with a plurality of burners in a combustion chamber. Nowadays, at least two different types of sequential gas turbines are known. According to a first embodiment, the first burner and the second burner are of an annular shape and are physically separated by a turbine blade stage called a high-pressure turbine. Downstream of the second burner, there is a second turbine unit (called a low-pressure turbine). This type of gas turbine is produced by the present applicant and is available on the market as "GT26". According to a second embodiment of the sequential gas turbine, the gas turbine is not provided with a high-pressure turbine, and the burner assembly is implemented in the form of a plurality of cylindrical burners. Each cylindrical burner includes a first burner and a second burner arranged one directly downstream of the other inside a housing of a common cylindrical shape. Moreover, this second type of sequential gas turbine is produced by the present applicant and is available on the market as "GT36". These two examples of gas turbine assemblies (GT26 and GT36) are cited only as non-limiting examples, where the present invention (i.e., the improved film seal as will be described hereinafter) can be applied during maintenance or can be incorporated from the start.

[0006] In almost all types of gas turbine assemblies, seals are used to seal the interface between stationary and moving components and / or to separate the hot gas flow from the cooling air flow. In particular, sealing the circumferential interface between the burner outlet and the first row of guide vanes of the turbine is a challenge nowadays because the seal must be ensured during all operating conditions. In fact, in this interface, the guide vanes usually have to be cooled (the temperature of the hot gas leaving the burner is very high), and, in addition, the guide vanes can move in both the axial and radial directions.

[0007] According to a first prior art example, the seal arrangement at the burner-turbine interface includes a honeycomb seal arrangement. In this example, each first guide vane includes an inner platform. The term "inner" means close to the turbine axis, as opposed to the outer end of the guide vane that is connected to the outer support housing surrounding the turbine. This platform consists of labyrinth teeth that act on a honeycomb seal fixed to the burner end (i.e., fixed to the end of the burner liner segment carrier). The above honeycomb seal arrangement involves some drawbacks. In particular, complete seal loss may occur during operation, or just partial delamination of the seal may result in a high leakage rate at this interface.

[0008] According to a second prior art example, the seal arrangement at the burner-turbine interface includes a film seal arrangement that extends radially from the end of the burner liner segment carrier to the vane platform. In particular, this film seal is referred to as a "dogbone" seal because it includes a film that connects an inner spherical end in a seat or groove in the burner liner segment carrier and an outer spherical end in a seat or groove in the inner vane platform. This solution allows for relative vane movement, particularly in the axial direction. However, by using a standard dogbone film seal (as will be clear during its description in Figure 4 and Figure 5 ), in some cases, the vane platform contacts the film. This contact reduces the pressure loading and unloads the inner circumferential seal line and the outer circumferential seal line located between the spherical end and the groove. Additionally, in this condition, the seal can vibrate, leading to an increased risk of wear and seal damage.

[0009] Although the above seal problem has been described with respect to the burner-turbine interface, the same problem also affects other locations within the gas turbine where it is necessary to separate the cooling air flow from the hot gas flow and where, depending on the operating conditions, some parts move relative to other parts. Summary of the Invention

[0010] Accordingly, a main object of the present invention is to provide a gas turbine suitable for overcoming the above problems of the prior art. In particular, a main object of the present invention is to provide a gas turbine having an improved film seal arrangement suitable for accommodating larger vane displacements without shortening the lifespan. A gas turbine suitable for being provided with the novel seal arrangement of the present invention is a gas turbine assembly for a power plant, which includes:

[0011] - at least one turbine provided with at least one row of stator vanes (first vanes) configured to direct the hot gas flow; wherein each vane is outwardly connected to the casing, includes an inner platform, and is configured to move axially and radially depending on the operating conditions;

[0012] - a liner segment carrier (stationary) that is radially inwardly positioned relative to the inner vane platform;

[0013] - a film seal arrangement located between the liner segment carrier and the inner vane platform for isolating the hot gas from a volume filled with higher pressure air located below (radially inward) the inner vane platform and on the downstream side of the film seal, and this high-pressure volume is supplied with cooling air (compressed air).

[0014] Although not all components are listed above, a person skilled in the art of gas turbines for power plants will know that a gas turbine for this purpose is an assembly that includes:

[0015] - A compressor that sucks in air and is configured to compress the air;

[0016] - A burner assembly configured to inject fuel into the compressed air and to combust the mixture to generate a hot gas flow.

[0017] The term "at least one turbine" means that the gas turbine according to the invention may include a single-stage turbine or a high-pressure turbine followed by a low-pressure turbine. Moreover, the burner assembly may include a single-stage burner or two sequential combustion stages. In the last case, the two combustion stages may be separated by a turbine stage (the above-mentioned high-pressure turbine) or may be directly serially connected in a housing of a common cylindrical shape (so-called cylindrical burner gas turbine).

[0018] The film seal arrangement of the present invention can be placed at many different positions inside the gas turbine. In particular, a critical position is the burner-turbine interface, i.e., the interface including the upstream end of the burner housing or liner segment carrier and the first row of guide vanes located downstream. As is known, at this interface, there is a main hot gas flow flowing from upstream to downstream into the guide vane airfoils and a cooling air flow (e.g., compressed air) flowing in a channel or volume (radially inwards) below the guide vane platform having a seal. The volume located on the downstream side of the seal is supplied with cooling air and is called the "higher-pressure air volume", while the volume located on the upstream side of the seal is called the "lower-pressure volume". As is common in the field of gas turbines, the terms "upstream" and "downstream" refer to the direction of the hot gas flow. The terms "inward / internal" and "outward / external" refer to the radial position relative to the axis of the gas turbine. Moreover, the terms "axially", "radially", and "circumferentially" refer to the axis of the gas turbine.

[0019] The cooling air flowing (radially inwards) below the guide vane platform cited above is necessary for cooling the guide vanes, and, importantly, for preventing this cooling air from leaking into the hot gas flow flowing (radially outwards) above the guide vane platform. In view of the above, the closest prior art of the present invention relates to a film seal arrangement called a "dogbone" seal. This known film seal includes:

[0020] - An inner spherical end received in a groove in a carrier (burner liner segment carrier or rotor cover housing);

[0021] - An outer spherical end received in a groove in the inner platform of the guide vane;

[0022] - A film connecting the inner spherical end and the outer spherical end. In a standard "dogbone" seal, the film can be defined as "in a straight line" relative to the spherical ends, i.e., the film is straight along the line connecting the centers of the spherical ends.

[0023] Starting from this prior art, the present invention provides an improved film seal arrangement (or simply referred to as a film seal), which comprises:

[0024] - An inner end, which is not necessarily spherical, and is received in a groove in a carrier (a burner liner segment carrier or a rotor cover housing);

[0025] - An outer end, which is not necessarily spherical, and is received in a groove in the inner platform of the guide vane;

[0026] - A film, which connects the inner end and the outer end, wherein the film of the present invention is not aligned with the line at the center of the connecting end, but is offset relative to the outer end towards the higher pressure volume (i.e., the volume located on the downstream side of the seal and supplied with cooling air). The term "offset" means that the outer end of the film is not connected to the middle part of the outer end, but is on the downstream side, thus producing a "P"-shaped outer end. The film is also offset relative to the inner end towards the lower pressure volume located on the opposite side of the seal.

[0027] In view of the above, according to the present invention, there is a larger gap between the upstream side of the guide vane groove and the film (offset downstream). This gap can be further increased by: making the straight film have a kink so that the angle of the film at the outer end is steeper than the angle of the film at the inner end, or making the film completely non-straight. The resulting larger gap allows the guide vane to perform a larger axial displacement during operation without contact between the guide vane platform and the seal film (see the example in Figure 10 .

[0028] According to an embodiment, the film of the present invention has through-holes, which are positioned close to the outer end and are configured to supply cooling air to the seal contact surface in the guide vane platform groove.

[0029] Alternatively or in combination with the holes above, the film may be provided with through-holes, which are positioned close to the inner end to cool the seal contact surface in the carrier groove and supply the cooling air as purge air to the lower pressure volume located on the upstream side of the film seal.

[0030] Preferably, the outer end is not spherical, but has a larger radial depth towards the downstream side to reduce the contact stress and the material loss rate when the guide vane platform wears into the outer end, so as to produce a continuous circumferential seal contact line, which minimizes the cooling air leakage. The outer edge is beveled to at least provide the same component introduction ability as the spherical end.

[0031] Preferably, the outer end has a smaller radial depth towards the upstream side to form a sacrificial wear surface. The guide vane row consists of individual guide vanes. Between the individual guide vane platforms, there is always an axial misalignment caused by production tolerances and assembly tolerances. During operation, the thermal movements of the individual guide vanes are different, thus enhancing the misalignment between the individual guide vane platforms. The sacrificial surface is shaped to allow the misaligned guide vane platforms to wear quickly and easily into the outer end of the film seal, and to form a continuous circumferential seal line on the lower pressure side to minimize the leakage of cooling air.

[0032] Preferably, the upstream sides of the inner and outer ends of the film seal are coated with a wear coating that minimizes wear and extends the life of the seal. Preferably, the upstream side of the guide vane groove is also coated with a wear coating that minimizes wear and extends the life of the guide vane. Preferably, the upstream side of the carrier groove is also coated with a wear coating that minimizes wear and extends the life of the carrier.

[0033] The large gas turbine has an upper part housing and a lower part housing bolted together at the parting line. Therefore, the film seal cannot be a continuous ring. The film seal arrangement of the present invention also includes an upper part and a lower part with overlapping ends to minimize the leakage amount. These overlapping ends require a cross-sectional shape different from the basic seal cross-sectional shape described previously. In particular, these end cross-sectional shapes are configured such that when overlapping, the same cross-sectional shape as the remaining part of the film seal is obtained (see the example in Figures 11 - 13 ).

[0034] Preferably, at least one film seal arrangement of the present invention is arranged at the burner-turbine interface.

[0035] Finally, the present invention relates to a method for retrofitting in a gas turbine, wherein a previous standard "dogbone" film seal is replaced with a new film seal according to the present invention (i.e., an "offset" film seal).

[0036] It will be understood that both the foregoing summary 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 be apparent from the following description, drawings, and claims.

[0037] In the appended claims, the features considered novel of the present invention are particularly set forth.

[0038] After carefully reading the detailed description with appropriate reference to the drawings, additional benefits and advantages of the present invention will become apparent. Description of the Drawings

[0039] However, the present invention itself may be most clearly understood in the following detailed description of exemplary embodiments of the present invention with reference to the accompanying drawings, in which:

[0040] Figure 1 and Figure 2 are schematic views of two gas turbine assemblies in which a seal according to the present invention may be applied;

[0041] Figures 3 - 5 are schematic views of two different examples of prior art solutions to the problem of sealing the burner-turbine interface;

[0042] Figure 6 discloses an embodiment of the present invention assembled in a gas turbine;

[0043] Figures 7 - 8 discloses two embodiments of a film seal according to the present invention;

[0044] Figures 9 - 10 discloses different operating conditions of a gas turbine including a film seal according to the present invention;

[0045] Figures 11 - 13 discloses two overlapping ends of a film seal according to the present invention. Detailed Description

[0046] In the following, the technical content and details of the present invention will be described in accordance with preferred embodiments in cooperation with the drawings. These embodiments are not intended to limit the scope of implementation of the present invention. The claims of the present invention cover all equivalent variations and modifications made in accordance with the appended claims.

[0047] Now, the present invention will be described in detail with reference to the drawings.

[0048] Now refer to Figure 1 , Figure 1 is a schematic view of a first non-limiting example of a gas turbine assembly (or simply referred to as a gas turbine) for a power plant that may be provided with a novel film seal according to the present invention. According to Figure 1 's embodiment, the gas turbine includes a so-called "sequential combustion gas turbine" provided with a high-pressure turbine and a low-pressure turbine. Along the main gas flow 2, Figure 1The gas turbine 1 includes a compressor 3, a first combustor 31, a high-pressure turbine 5, a second combustor 32, and a low-pressure turbine 7. The compressor 3 and the two turbines 5, 7 are parts of or connected to a common rotor 8 that rotates about an axis 9 and is surrounded by a concentric housing 10. The compressor 3 is supplied with air and is provided with rotating blades 18 and stator vanes 19 configured to compress the air entering the compressor 3. The compressed air leaving the compressor flows into a plenum 11 and from there into a number of first burners 12 of the first combustor 31, which is arranged in a circular pattern about the axis 9. Each first burner 12 is configured to inject at least one type of fuel (connected to at least one first fuel source 13) into the compressed air stream. Preferably, the first burner 12 can be defined as a "premix" burner since it is configured to mix the compressed air and the injected fuel before ignition. The fuel / compressed air mixture flows into an annular-shaped first combustion chamber 4, where the mixture is ignited. During startup, the mixture is initially ignited by an igniter (e.g., by a spark igniter); once ignited, the ignition is self-sustaining and the igniter is turned off. The resulting hot gas leaves the first burner chamber 4 and expands partially in the high-pressure turbine 5, thereby doing work on the rotor 8. Downstream of the high-pressure turbine 5, the partially expanded hot gas flows into a row of second burners 33, where at least one type of fuel is injected by fuel guns 14 (each burner has one gun). The partially expanded gas has a high temperature and contains sufficient oxygen to undergo further combustion by autoignition in a second combustion chamber 6 arranged downstream of the row of second burners 33. These second burners 33 are also referred to as "reheat" burners. The reheated hot gas leaves the second combustion chamber 6 and flows into the low-pressure turbine 7, where the reheated hot gas expands, thereby doing work on the rotor 8. The low-pressure turbine 7 includes multiple stages or rows of rotor blades 15 arranged in series along the main flow direction. Such multiple rows of rotor blades 15 are interspersed with multiple rows of stator vanes 16. The rotor blades 15 are connected to the rotor 8, and the stator vanes 16 are connected to a vane carrier 17, which is a concentric housing surrounding the low-pressure turbine 7. Although not visible in Figure 1 , the gas turbine 1 has a number of seals, in particular, for example, a number of film seals arranged at the second combustor - first turbine vane interface. In view of the above, a gas turbine as disclosed in Figure 1 can be provided with the novel film seals of the present invention.

[0049] Now referring to Figure 2 , Figure 2FIG. is a schematic view of a second non - limiting example of a gas turbine that may be provided with a novel film seal according to the present invention. This gas turbine 20 is also a "sequential combustion gas turbine" that may be provided with innovative features according to the present invention. In particular, Figure 2 A partial view of a gas turbine 20 having a compressor 29, a turbine 21, and a sequential combustor 22 is shown. Figure 2 The sequential combustor 22 has a number of so - called can - type combustors, that is, a number of bolted - together housings, in which there are a number of first burners 24, for example, a first combustion chamber 25, a second combustion chamber 27, four first burners 24, and a number of second burners 26. Upstream of the second burners, a mixer may be provided for adding air to the hot gas leaving the first combustion chamber 25 and generating turbulence in the air / hot - gas mixture. The sequential combustor arrangement is at least partially housed in an outer casing 28, and the outer casing 28 supports individual can - type combustors 22 arranged in a circular pattern around the turbine axis 23. At least one type of fuel is introduced into the first burners 24 via a first fuel injector (not shown), where the fuel is mixed with the compressed air supplied by the compressor 29. Moreover, each of the first burners 24 in this embodiment is a "premix" burner configured to produce a premixed flame. When the hot gas leaves the second combustion chamber 27, the hot gas then expands in the turbine 21, thereby doing work on the rotor 30. Although not visible in Figure 2 this, the gas turbine 20 includes a number of seals, including, for example, a film seal that may be arranged at the can - type burner - first turbine vane interface. In view of the above, the gas turbine as disclosed in Figure 2 may be provided with the novel film seal of the present invention.

[0050] Reference Figure 3 , Figure 3 FIG. is a schematic view of a burner - first vane turbine interface of a gas turbine unit according to the prior art. This gas turbine 34 includes at least one burner schematically denoted by reference numeral 35 that generates a hot gas stream 36. In particular, Figure 3 shows the interface between the burner and the turbine 37. This interface is defined by:

[0051] - a burner liner 38,

[0052] - a first row of vanes 39, which is outwardly supported by a casing 40 and inwardly includes a vane platform 41 having labyrinth teeth 42 on the side adjacent to the burner liner 38,

[0053] - a rotor cover 43, which is located between the rotor (not shown) and the vanes 39, has a cooling air passage 44 (i.e., a compressed air supply), and

[0054] - A honeycomb seal 45, which is attached to the burner liner 38 and, together with the guide vane labyrinth teeth 42, seals the cooling air from the hot gas flow 36.

[0055] As described in the prior art section, the practice of this prior art has deficiencies. In particular, the ability to accommodate the operating movement of the guide vane 39 is limited. The practical solutions of the prior art are shown as assembled in Figure 4 and operating as in Figure 5 In Figures 4 - 5 the seal arrangement includes a standard "dogbone" film seal 46. The "dogbone" film seal 46 has an upper portion and a lower portion that overlap near the gas turbine housing parting line. And, as shown, it has an inner spherical end 47, an outer spherical end 48, and a straight central film 49 (provided with purge air holes 50) that connects the two spherical ends. The inner spherical end 47 is received in a circumferential groove 51 in a carrier 52. The carrier 52 can be part of the burner liner or part of the rotor cover. The outer spherical end 48 of the seal is received in a circumferential groove 53 in the guide vane platform 41. As previously discussed, the two spherical ends require continuous circumferential contact lines with the guide vane platform and the carrier at positions 57 and 58 to minimize the leakage of cooling air into the main flow hot gas. Additionally, the film 49 has holes 50 (radially positioned near the center of the film) to supply cooling air 56 to purge the hot gas from the cavity 59 located between the guide vane platform and the liner segment. The "dogbone" film seal 46 separates the hot gas flow 36 from the cooling air 56. In the case of defining upstream and downstream with respect to the main direction of the hot gas flow 36, the volume 54 upstream of the seal 46 is the lower pressure side, while the downstream volume 55 (supplied with cooling air) is the higher pressure side. Figure 4 The assembled position is shown. In operation, the guide vane platform 41 has a large radially inward and axially downstream movement relative to the outer spherical end 48; Figure 5 shows the upstream face of the groove in the guide vane platform 41 contacting the film in the steady-state operating position. The pressure drop above the "dogbone" seal 49 generates a high force at the contact point 59 shown in Figure 5 This contact portion 59 limits the ability of the "dogbone" seal to accommodate the operating movement of the guide vane 39 relative to the carrier 52. The seal pressure load is transferred to the contact point 59, which shortens the life of the "dogbone" seal because the inner seal contact line 58 and the outer seal contact line 57 unload, allowing the "dogbone" film to vibrate. This vibration increases wear and creates a risk of seal rupture during operation. In the case of using a standard symmetric "dogbone" film seal, as in Figure 5As shown, the ability to accommodate a large displacement of the guide vane 39 relative to the carrier 52 can only be achieved by increasing the radial height of the membrane. However, due to the design requirements of the gas turbine, it is generally not possible to increase the radial height of the membrane seal.

[0056] As disclosed in the Summary of the Invention section of the present invention, the present invention provides a solution for overcoming the above deficiencies and allowing a large relative movement of the guide vane in the downstream direction (i.e., downstream relative to the main hot gas flow direction). According to the general definition of the present invention, the proposed new solution is to provide a membrane seal having an end in the groove guide vane platform and an end in the carrier, and the ends are connected by a membrane offset toward the higher pressure side (i.e., the downstream side of the seal). Figure 6 An embodiment according to the present invention is shown, in which the elements also shown in Figure 4 and Figure 5 are labeled with the same reference numerals. Since the present invention also relates to a maintenance method for replacing the "dog bone" seal, all adjacent parts are the same as those shown in Figure 4 and Figure 5 . As shown in Figure 6 , this example of the improved membrane seal 60 includes an inner end 61 configured to produce a continuous circumferential sealing contact line 57 on the upstream surface of the circumferential groove 51 of the carrier. The new membrane seal 60 includes an outer end 62 configured to produce a continuous circumferential sealing contact line 63 on the upstream surface of the circumferential groove 53 of the guide vane platform. The membrane 64 is offset toward the higher pressure side 55, that is, it is connected to the outer end 62 not in the centered position but near the higher pressure side downstream of the outer end 62. Therefore, the outer part of the seal has a "P" shape profile. Due to this specific profile shape, there is a large gap between the upstream surface of the guide vane groove ( Figure 6 reference numeral 67 in Figure 6 embodiment, the membrane 64 is also not connected to the inner end 61 in a centered manner; the membrane is offset toward the lower pressure side of the seal at the inner end 61 so that the seal can accommodate a large relative movement of the guide vane without contact between the membrane and the carrier. According to this configuration, the complete cross-sectional profile of the membrane seal 60 can be defined as being substantially "S" shaped. Figure 6The seal in [ ] also shows two rows of through - holes 66, 65 at the position of the membrane 64 near the inner end 61 and the outer end 62 of the seal. The cooling air passing through the inner row of through - holes 65 cools the inner end of the seal and the groove contact surface, and performs the main function of purging the hot air from the cavity located between the guide vane platform and the bushing segment; the cooling air passing through the outer row of through - holes 66 cools the outer end of the seal and the groove contact surface before being discharged into the cavity to purge the hot air. The outer row of through - holes 66 is inclined to direct the cooling air onto the groove surface near the contact line 63.

[0057] Figure 7 Another embodiment of the seal according to the present invention (i.e., a seal having a membrane offset towards the higher - pressure side of the seal at the outer end) is disclosed. In this example, the inner end 61 is circular, and the membrane 64 is connected to the inner end 61 centrally. In this embodiment, the membrane 64 has a single row of through - holes for supplying cooling air into the cavity to purge the hot air; the outer end 62 of the seal at the seal contact line 63 has a reduced radial depth and is a sacrificial wear surface.

[0058] Figure 8 Another embodiment of the seal of the present invention is shown. In this example, the seal is very similar to the seal in [ ]. However, Figure 7 the seal in [ ] Figure 8 includes two rows of holes, as shown in [ ], and the outer end 62 has circumferentially patterned slots 68 on the higher - pressure side to reduce the circumferential stress caused by the asymmetry of the end profile. Figure 6

[0059] Figure 9 Figure 10 and Figure 6 shows Figure 9 two different operating conditions of the seal in [ ]. In particular, Figure 6 shows the position of the components after assembly (as in [ ]), while Figure 10 Figure 10 shows the position of the components under steady - state operating conditions (the guide vane platform is displaced radially inwards and axially downstream relative to the carrier). As shown in [ ], due to the membrane offset, the seal of the present invention can accommodate a relatively large guide vane movement relative to the carrier. At the operating position in [ ], there is neither contact between the guide vane platform and the membrane connecting the two ends of the seal nor contact between the carrier and the membrane connecting the two ends of the seal. Figure 10

[0060] ​​​​The seal of the present invention has an upper portion and a lower portion that overlap near the gas turbine housing parting line so that the entire circumference of the burner-vane interface is sealed. The seal cross-sectional profile at the overlap of the upper and lower portions of the seal is such that when the seal is assembled, the cross-sectional profile of the seal at the overlap position is the same as that at the seal cross-sectional segments at all other circumferential positions. Figure 11 and Figure 12 shows the ends of the upper and lower portions which, when overlapped as shown in Figure 13 have the same cross-sectional profile as the remainder of the seal.

[0061] Although the invention has been explained with reference to the (one or more) preferred embodiments of the invention as mentioned above, it will be understood that many other possible modifications and variations can be made without departing from the scope of the invention. Accordingly, it is intended that the appended claims or claim will cover such modifications and variations that fall within the true scope of the invention.

Claims

1. A gas turbine for a power plant; The gas turbine (1, 20) comprises: - at least one turbine (5, 7, 21, 34) provided with at least one row of stator vanes (16, 39) configured to direct a hot gas flow (36); each vane (16, 39) is outwardly connected to a casing (17, 40) and includes an inner vane platform (41); - a bushing segment carrier (52) positioned radially inwards relative to the inner vane platform (41); - a film seal (46) extending from the carrier (52) to the inner vane platform (41) to isolate the hot gas from a higher pressure cooling air located below the inner vane platform (41) and downstream of the film seal (46) in a higher pressure volume (55) supplied with cooling air (56); wherein the film seal (46) comprises: - an inner end (61) received in a first groove (51) in the carrier (52); - an outer end (62) received in a second groove (53) in the inner vane platform (41); - a film (64) connecting the inner end (61) to the outer end (62); characterized in that the film (64) is offset towards the higher pressure volume (55) at the outer end (62) such that the film seal (46) includes an outer portion having a "P"-shaped profile.

2. The gas turbine according to claim 1, wherein, The film seal (46) is provided with a row of through-holes (66) positioned adjacent the outer end (62) of the film seal to supply cooling air to the upstream side of the second groove (53) in the inner vane platform (41).

3. The gas turbine according to claim 1, wherein, The film seal (46) is provided with a row of through-holes (65) positioned adjacent the inner end (61) or centrally located (67) on the film (64) to supply cooling air as purge air to a lower pressure volume (54) upstream of the film seal (46).

4. A gas turbine according to any one of claims 1 to 3, wherein, The outer end (62) has circumferentially patterned slots (68).

5. A gas turbine according to any one of claims 1 to 3, wherein, The outer end (62) includes an upstream portion (63) having a reduced radial depth to act as a sacrificial wear surface such that a continuous circumferential seal line is formed between the outer end (62) of the film seal and a separate axially misaligned second groove (53).

6. A gas turbine according to any one of claims 1 to 3, wherein, The adjacent surfaces of the upstream side of the outer end (62) and the upstream side of the film (64) are coated with a wear coating.

7. A gas turbine according to any one of claims 1 to 3, wherein, The adjacent surfaces of the upstream side of the inner end (61) and the upstream side of the film (64) are coated with a wear coating.

8. A gas turbine according to any one of claims 1 to 3, wherein, The cross-sectional profile of the film seal (46) is substantially "S"-shaped.

9. A gas turbine according to any one of claims 1 to 3, wherein, The inner end (61) and / or the outer end (62) is substantially circular in shape.

10. A gas turbine according to any one of claims 1 to 3, wherein, The gas turbine includes an upper portion of the film seal and a lower portion of the film seal in the circumferential direction, the film seal (46) having a cross-sectional profile at the inner and outer ends such that when overlapping, the same cross-sectional profile is produced at all other circumferential positions on the film seal (46).

11. A gas turbine according to any one of claims 1 to 3, wherein, The gas turbine includes at least one burner, and the film seal (46) is arranged at the burner-turbine interface.

12. A method for retrofitting a gas turbine for a power plant that is already in a maintenance state; the method comprising the steps of: a) providing a gas turbine (5, 7, 21, 34) according to the preamble of claim 1, which includes a plurality of previous film seals; b) providing a plurality of film seals (46) according to the characterizing part of claim 1; c) replacing the previous film seals with film seals according to the characterizing part of claim 1.

13. The method according to claim 12, wherein, The step a) of providing the gas turbine is carried out by providing a sequential gas turbine (1) including a high-pressure turbine (5) and a low-pressure turbine (7).

14. The method according to claim 12, wherein, The step a) of providing the gas turbine is carried out by providing a sequential gas turbine (20) including a plurality of cylindrical burners (22).

15. The method according to any one of claims 12 to 14, wherein, The step c) of replacing the previous film seals with the film seals (46) according to claim 1 is carried out by replacing the previous film seals arranged at the burner-turbine interface.

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